Category Paving the Way for Apollo 11

Ranger 4 arrived at the Cape on 26 February 1962. The countdown ran smoothly to liftoff at 20:50 GMT on 23 April. The Atlas performed satisfactorily, the Agena achieved the desired parking orbit and then made the translunar injection manoeuvre as planned. But when the spacecraft appeared above the horizon at Johannesburg it was transmitting a carrier signal without encoded telemetry, which meant that it was not possible to determine the state of the systems. Unable to lock onto the Sun, the initial instabilities imparted by separation from the spent stage caused the spacecraft to tumble. It was concluded that the master clock in the computer/sequencer must have failed. Ranger 4 had transmitted telemetry during the ascent, but the clock had stopped at some point in the gap in coverage between the vehicle passing beyond the final station of the Eastern Test Range and its coming into range of Johannesburg. In effect, the Agena had released a lobotomised robot. Ironically, the radar tracking by Woomera showed that the slight discrepancy in the trajectory would have been well within the capacity of the spacecraft to correct.

The target for Ranger 4 was the same as for its predecessor. The Moon was ‘full’ on 20 April and would be ‘last quarter’ on 27 April. James Webb, W. H. Pickering, Oran Nicks, Clifford Cummings and James Burke congregated at Goldstone on 26 April as the spacecraft approached the Moon. Without solar power, the battery had expired, terminating the carrier wave from the main transmitter, but the fact that the independently powered surface package was transmitting enabled radio tracking to continue. At 12:47 GMT, some 64 hours after launch, the spacecraft passed behind the leading limb of the Moon. Calculations showed that it impacted 2 minutes later.

For the first time, American hardware had hit the Moon, marking a success for

the launch vehicle – but this was little consolation for the spacecraft engineers, and the scientists gained nothing. On the one hand, in the original concept of the Block II it was deemed that three flights would be necessary to have a reasonable probability of achieving a fully successful flight that would deliver the scientific objectives of the project. However, the scientists appeared to think that every flight should succeed! It was difficult to precisely determine why Ranger 4’s clock had stopped, because the failure occurred when out of communication and it prevented the transmission of telemetry. But because the telemetry was present when the spacecraft was last seen on the Agena and absent following its release, attention focused on the separation procedure – in particular, when the umbilical plug was withdrawn from the bus and the onboard computer/sequencer issued the power-up command to the systems. It was inferred that there must have been a short circuit at this time, and the obvious root cause was the heat sterilisation process. Additional waivers were issued, but the computer/sequencer for Ranger 5 had already been treated. With mass no longer an issue for the Block II, it was decided to install a backup clock in order to ensure that telemetry would be produced in the event of the computer/sequencer being disabled.

When the Apollo Site Selection Board met on 30 March 1967 the Apollo officials announced that whilst they would seek further Lunar Orbiter data, that from the first three missions satisfied “the minimal requirements of the Apollo program for site survey for the first Apollo landing”. By now nine mare sites in the Apollo zone were deemed to be suitable as ‘prime sites’ for the early Apollo landings: one in Mare Foecunditatis, two in Mare Tranquillitatis, one in Sinus Medii and five in Oceanus Procellarum. These were designated ‘Set B’.4 For each such site, the US Geological Survey produced geological maps at scales of 1:25,000 and 1:100,000 to supplement the 1:1,000,000 regional maps. It was noticed that the sites on the eastern maria had high densities of large but shallow craters, and the sites on the western maria were generally flatter but rougher in detail. The astronomers had long ago noted that there was a difference in spectral hue, with the eastern maria being bluish and the western maria reddish.

On 15 December 1967 the Apollo Site Selection Board convened at the Manned Spacecraft Center to refine the target list for the first Apollo landing. All of the sites of Set B were acceptable in terms of their approach routes. However, as a landing in Mare Foecunditatis would not allow sufficient time after rounding the eastern limb for radio tracking to verify the lander’s trajectory prior to powered descent, this site was discarded. Five sites were short-listed as ‘Set C’. It was decided that three of these must be selected as options for the first landing mission, forming a prime site and two backups spaced in lunar longitude to accommodate successive 2-day delays in launch. It was recognised that the need for the crew to familiarise themselves with three sites would increase their training burden, but there would be no impact on the surface activities because the first landing was not to include a mapped traverse. In east to west sequence, the five sites were II-P-2, II-P-6, II-P-8, III-P-11 and II-P-13. Whilst it was clear that the prime site would be in the eastern hemisphere, the meeting did not specify whether it should be II-P-2 or II-P-6.

On 26 September 1968 the Set C ellipses were ‘stretched’ from 5.3 x 7.9 km to 5.0 x 15.0 km to allow for uncertainties in the Moon’s gravitational field that might cause a lander to come in either ‘short’ or ‘long’ of the designated aim point. On 3 June 1969 the Set C sites were renamed Apollo Landing Site (ALS) 1 through 5 respectively.

They were I-P-1 in Mare Foecunditatis, II-P-2 and II-P-6 in Mare Tranquillitatis, II-P-8 in Sinus Medii, and II-P-11, III-P-9, III-P-11, III-P-12 and II-P-13 in Oceanus Procellarum.

In mid-1966 Sam Phillips had hoped that AS-206 would be able to launch LM-1 in April 1967, and Kurt Debus, estimating that it would take 6 months to check out the spacecraft, had asked Grumman to send it to the Cape in September 1966. But it was delayed by manufacturing issues and combustion instabilities in the ascent engine. Nevertheless, AS-206 was erected on Pad 37 in January 1967 in the expectation of launching in April. However, because the AS-204 launch vehicle on Pad 34 had not been damaged by the fire that destroyed the Apollo 1 spacecraft, on 20 March it was reassigned to LM-1. Accordingly, by 11 April AS-206 had been returned to storage and AS-204R – as this was redesignated – erected in its place. In the absence of the spacecraft, Grumman built a plywood mockup on the pad for facilities verification.

On 12 May George Low confided to headquarters that although Grumman had promised to deliver LM-1 in June, he was sceptical. John J. Williams headed a 400- man operations team at the Cape. After the arrival of the ascent and descent stages on 23 June, LM-1 was mated on 27 June. However, the initial examination identified a significant number of departures from specification. On 26 July Carroll Bolender was reassigned to Houston as ASPO’s LM Manager. LM-1 was de-mated in August to repair leaks in the ascent stage. After another leak developed in September, it was de-mated and a number of items extracted for return to Grumman. After the testing was finally completed, the spacecraft, minus its legs, was mechanically mated to the launch vehicle on 19 November and a nose cone fitted in place of the absent CSM. The flight readiness tests were finished in late December. The cabin closeout was on 18 January 1968, during the countdown demonstration test. Loading the hypergolic propellants into LM-1 was delayed by procedural issues, but the ensuing tests ended on 19 January.

The terminal countdown began on 21 January, at T-10 hours 30 minutes. The spacecraft went onto internal power at T-42 minutes and, several hours later than planned, AS-206 lifted off as Apollo 5 at 22:48:08 GMT on 22 January 1968.

The S-IVB achieved an orbit ranging between 88 and 120 nautical miles, shed the nose cone and then splayed the four SLA panels. LM-1 was released at 000:53:50. After manoeuvring clear using its attitude control thrusters, the LM adopted a ‘cold – soak’ orientation, which its guidance system successfully maintained with a minimal engine duty cycle.

The mission plan called for two descent propulsion system manoeuvres, an abort staging, and an ascent propulsion system manoeuvre. The first manoeuvre was to occur on the third revolution and last 38 seconds. It would run at 10 per cent throttle for the first 26 seconds, then be concluded at full throttle. The thrust profile of the second manoeuvre was to be representative of flying a lunar landing, involving five phases over a total of 734 seconds. The abort staging sequence would be initiated at

full throttle, and was to include descent propulsion system shutdown and a ‘fire in the hole’ 5-second burn by the ascent propulsion system. The final burn was to run to propellant depletion and end the primary mission. The guidance system initiated the first descent propulsion system firing at 003:59:42 but the buildup of thrust did not satisfy the programmed velocity-time criteria, and the guidance system, sensing that the spacecraft was not accelerating as rapidly as expected, aborted the burn after just 4 seconds. In fact, this was a design feature, since on a manned mission it would allow the crew time to analyse the situation and decide whether to restart the engine to continue. In normal circumstances the engine would have fired with full tank pressurisation and achieved the desired thrust in 4 seconds, but in this case the tanks were only partially pressurised and it would have taken 6 seconds to build up thrust. The premature cutoff was merely the result of inadequate coordination between the guidance and propulsion teams, not a problem with the spacecraft. Mission Control sent a command to deactivate the guidance system in order to permit the remainder of the mission to be controlled from Earth using a preplanned sequence which would address the minimum requirements of the mission.

At 006:10:00 the onboard automatic sequencer initiated this program. It began by using the backup control system to control the vehicle’s attitude. In performing two burns in this mode, the descent engine gimballed properly and responded smoothly to throttle commands. But the short duration of the three descent propulsion system firings precluded a full evaluation of the thermal aspects of the supercritical helium pressurisation system. In abort staging, all system operations and vehicle dynamics were satisfactory for manned flight. The primary control system was then reselected to control the vehicle’s attitudes and rates. However, as this had been off during the abort staging sequence its computer program did not know of the change of mass resulting from this action and its computed thruster firing times were based on the mass of the two-stage vehicle and caused an extremely high rate of propellant usage. The final ascent propulsion system firing started at 007:44:13 and ran to thrust decay at 007:50:03. Since the attitude control system had by that time exhausted its own propellants, this burn was initiated with the thrusters drawing from the tanks of the ascent propulsion system. This continued until the sequencer automatically closed the interconnect valves, whereupon, with the thrusters starved and the ascent engine still firing, the vehicle started to tumble. The rates were soon of such a magnitude as to impede the flow of propellants to the engine, and helium ingestion induced thrust decay prior to propellant depletion. The vehicle had been in a retrograde orientation during the controlled portion of its final manoeuvre, and calculations indicated that it entered the atmosphere over the Pacific Ocean.

On 26 January the LM-2 flight requirements meeting determined that: (1) apart from minor anomalies, LM-1 had achieved all its flight objectives; (2) it should be possible to achieve the objectives for LM-2 either by additional ground testing or on a manned mission; and (3) it was not necessary to undertake additional unmanned flights to ‘man rate’ the LM. Grumman’s own view was that there should be two test flights, but the company relented after the review of the LM-1 data by the Manned Space Flight Management Council on 6 February. On 6 March NASA cancelled the shipment of LM-2 to the Cape. If AS-502 repeated the success of its predecessor, then AS-503 would indeed be manned, and hopefully be launched before the end of the year with CSM-103 and LM-3.

MISSILES AND SPACE

When a team of German rocket experts surrendered to the US Army in May 1945 and General Holger ‘Ludy’ Toftoy, an artillery officer serving as Chief of Ordnance Technical Intelligence in Europe, set out to arrange their relocation to the USA, the V-2 missile was seen as an important military technology. However, this perception changed with the introduction of the atomic bomb in August against Japan. In the immediate post-war years the US military felt that strategic aircraft carrying atomic bombs would enable it to defeat any enemy. In this context, a ballistic missile which could fly only several hundred kilometres to deliver about 1,000 kg of conventional explosive was insignificant. Consequently, upon being settled in El Paso, Texas, the German team led by Wernher von Braun found themselves with little to do.

Although the ballistic missile had seemingly become obsolete as a weapon, it held out the prospect of serving a more benign role, and in November 1945 the US Navy recommended the development of a satellite. The Army Air Force agreed. However, each service felt that it alone should be assigned this task.

In 1946 the RAND Corporation, created as a ‘think tank’ for the Army Air Force, said: ‘‘The achievement of a satellite craft by the United States would inflame the imagination of mankind, and would probably produce repercussions in the world comparable to the explosion of the atomic bomb. […] Since mastery of the elements is a reliable index of material progress, the nation which first makes significant achievements in space travel will be acknowledged as the world leader in both military and scientific techniques. To visualise the impact on the world, one can imagine the consternation and admiration that would be felt here if the US were to discover suddenly that some other nation had already put up a successful satellite.’’

Meanwhile, von Braun was showing the Army how to assemble, prepare and fire V-2 missiles at the White Sands Proving Grounds in New Mexico. They were made from parts either recovered from Germany or manufactured to his specifications in America. In 1948, while in Texas, von Braun wrote a book, Das Marsprojekt, in which he outlined how an expedition to explore Mars might be undertaken. It was a

‘grand design’ which left the details to be developed in due course. He set out “more or less to project the technology that existed then’’ to motivate young engineers. He argued that a mission would be feasible ‘‘in 15 to 20 years’’ if a nuclear-powered ion engine could be created. The expedition would involve ten space ships with a crew totalling around 70 people. The ships were to be assembled in Earth orbit, with three carrying ‘landing boats’ for Mars. Later in 1948, von Braun’s team was relocated to the Redstone Arsenal of the Army Ordnance Corps in Huntsville, Alabama. It was a new establishment on the site of facilities used by the Chemical Corps in the Second World War, and was to undertake research and development of rockets and missiles.

In September 1949 the Soviets exploded an atomic bomb – at least 3 years earlier than the US had expected. Although the Soviet bomb was not yet a weapon, it was evident that America would soon lose its monopoly. In early 1950 President Harry S. Truman authorised the hydrogen bomb. In 1951 funding was made available for preliminary work for what would become the Atlas intercontinental-range ballistic missile. The hydrogen bomb test at Eniwetok Atoll on 1 November 1952 was not a viable weapon, because it weighed 60 tonnes. But as the bomb’s weight was reduced for carriage by aircraft it was realised that if it were to prove possible to make the device even smaller, it might become feasible to develop a ballistic missile capable of delivering it. The Air Force (which had gained its independence from the Army in 1947) created a committee chaired by physicist John von Neumann. This was asked to predict the trend in weight-to-yield ratio of hydrogen bomb development, estimate the warhead that a ballistic missile might deliver over intercontinental range by the end of the decade, and assess whether the probable accuracy would make a warhead of that yield a viable weapon. In February 1954 the committee reported that progress with warheads would make missiles viable. The RAND Corporation endorsed this conclusion. Although the Air Force responded by assigning the development of an intercontinental-range ballistic missile ‘top priority’, Secretary of Defense Charles E. Wilson, who was in tune with the ‘economic conservatism’ of the administration of President Dwight D. Eisenhower deliberated on the matter for over 12 months until informed in 1955 that a recently established radar intelligence station in Turkey that was operated by the US had discovered that the Soviets were well advanced in the development of their own intercontinental-range ballistic missile – test flights were launched from a site east of the Black Sea and passed across Soviet territory to fall near the Kamchatka Peninsula. America had felt safe because the USSR had no strategic bombers, but a ballistic missile would be able to circumvent America’s air defences. The risk was that when the Soviet missile entered service with a nuclear warhead it would be able to wipe out the US bomber bases in a ‘first strike’ which would prevent retaliation against the Soviet Union. The US therefore simply had to have its own fleet of missiles.

On 15 June 1962 Brainerd Holmes issued Requirements for Data in Support of Project Apollo, in which he called for three types of information about the Moon as a matter of priority, certainly within the next few years. First, environmental data on particles and fields in space near the Moon to assist in the design of manned spacecraft and assure the safety of crews both in flight and on the Moon. Second, information on the physical properties of the lunar surface in order to confirm the design of the Apollo landing gear. Third, photo-reconnaissance and topographical data in order to facilitate early selection of Apollo landing sites. Holmes had not consulted Homer Newell in drawing up this list of requirements, he simply expected that since Apollo was the agency’s pre-eminent program Newell would arrange for the information to be provided as soon as possible – and pay the bill out of his own office’s budget. But Newell’s Space Sciences Steering Committee had its own priorities.

Holmes supported the unmanned lunar projects which would provide information for Apollo, but opposed those intended to undertake tasks which astronauts would soon be able to do. He therefore opposed Prospector, which was to collect and return lunar samples to Earth. Accepting this logic, Newell set out to ensure that astronauts performed useful science while on the Moon.[21] Soon after being appointed Director of the new Office of Space Sciences, Newell arranged for the Space Science Board of the National Academy of Sciences, now chaired by the Princeton geologist Harry H. Hess, to arrange a series of joint workshops to discuss the best way to undertake space science. The first Summer Study was held at the University of Iowa between 17 June and 10 August 1962, with over 100 representatives of NASA, academia and industry. The aim was to evaluate past and current programs, and recommend future

programs. Afterwards, the Space Science Board issued a summary report, A Review of Space Research, in which it acknowledged that Apollo would start off as “an engineering effort”, but expressed the hope that “scientific investigations will later become the primary goals”.

In September 1962 Gene Shoemaker began a 12-month secondment to NASA to assist the Office of Space Sciences. His motivation for taking this post was partly to increase his chances of becoming one of the astronauts who would have the good fortune to undertake field geology on the Moon. Don Elston served as Acting Chief of the Branch of Astrogeology in Shoemaker’s absence.

On 11 October 1962 Robert Seamans called in Homer Newell, Brainerd Holmes, W. H. Pickering and Oran Nicks, and told the Office of Manned Space Flight and the Office of Space Sciences to coordinate their lunar activities. Newell was told that his priority was to support Apollo’s requirement for data about the Moon. Nicks was to coordinate, and report how unmanned missions could best contribute to Apollo. In particular, could further Rangers provide some of this data by delivering a surface package incorporating a penetrometer to measure the strength of the lunar surface. When on 15 October Newell publicly announced five additional Rangers for 1964 equipped with the high-resolution TV system, he emphasised they “would increase the probability of obtaining lunar surface detail information that could be used in the manned landing system’’. Newell was also considering another series for 1965 which would deliver surface capsules. If these rough landers were funded, they would be primarily for scientific research. For these, Newell asked the Aeronutronic Division of Ford, which had developed the seismometer, to investigate a small TV camera capable of being delivered to the lunar surface in a capsule. On 22 October Holmes and Newell announced that a Joint Working Group would be formed, composed of representatives of their two offices. Chaired by Gene Shoemaker, at least during his period of secondment to NASA, it would be responsible for recommending to the Office of Manned Space Flight “a detailed program of scientific exploration’’ and for recommending to the Office of Space Sciences “a program of data acquisition to assure a timely flow of environmental information into planning for manned projects’’. It would also be responsible “for establishing and maintaining close liaison with field centers, government agencies and universities in the development of an integrated scientific program for manned space flights’’.

Seamans’s directive that the Office of Space Sciences fly lunar missions primarily in support of Apollo, rather than for purely scientific purposes, renewed Newell’s determination to ensure that Apollo crews conducted proper science whilst orbiting the Moon and on its surface – the Manned Spacecraft Center, being fully occupied with the engineering challenge of sending men to the Moon, was slow to pursue this aspect of the program.

A BOUNCY LANDING

The 6-month hiatus after Surveyor 2 was not due to concern over the loss of that spacecraft, but to the wait for the restartable version of the Centaur. The two-burn configuration traded payload capacity against the hardware to restart the engines and the cryo-propellants that would be vented while coasting in parking orbit, but it offered Surveyor launches in winter months,[33] considerably lengthened the launch windows, and increased the flexibility in selecting the arrival time for optimal illumination at the landing site. The Centaur stage demonstrated its restart capability by a launch on 26 October 1966, thereby completing its test program.

The target for Surveyor 3 was a 60-km-diameter circle in the southeastern part of Oceanus Procellarum, centred 120 km southeast of the crater Lansberg. Its attraction was that although it was crossed by a ray from Copernicus 370 km to the north, at telescopic resolution it was sparsely cratered. The smooth patch of mare material was broken about 20 km to the west by rough hummocky terrain and isolated hills, and was bounded to the east by low ridges. It had been photographed at medium resolution by Lunar Orbiter 1 as target I-P-7, and at high resolution by Lunar Orbiter 3 as III-P-9c. These pictures revealed the presence of sub-telescopic craters in the target circle – with one, just over 1 km in diameter, situated very near the aim point. As a smooth-looking patch of Oceanus Procellarum, it bore a similarity to the Surveyor 1 landing site in the Flamsteed Ring, some 650 km to the west.

Surveyor 3 lifted off from Pad 36B at 07:05:01 GMT on 17 April 1967. The Atlas jettisoned its booster section at T+ 142 seconds and the sustainer engine shut down at T + 238. Once free of the Atlas, the Centaur established the desired circular parking orbit at an altitude of 160 km – with insertion at T+569 seconds. The coasting phase would vary between 4 minutes and 25 minutes, depending on the geometry of the translunar injection – in this case it was to be 22 minutes 9 seconds. While coasting, the Centaur first fired two 50-pound-thrust hydrogen peroxide thrusters to settle the remaining propellants in their tanks, then continuously fired two 3-pound thrusters to maintain this condition. It had two clusters of 3.5 and 6- pound thrusters to control its attitude, and while maintaining its longitudinal axis to the local horizontal it rolled at a rate of 0.17 degree per second in order to even out solar heating of its payload and vented any propellant boil-off.

Owing to the predawn launch, the Centaur emerged from the Earth’s shadow at 07:21:25. About 40 seconds prior to the translunar injection, the 50-pound thrusters fired again to guarantee that the propellants would enter their feed pipes. The main engines were shut down when the inertial guidance system sensed that the requisite velocity had been achieved – in this case at 07:38:49, after a 108-second burn. As on earlier missions, after it had configured and released the spacecraft, the spent stage performed the separation manoeuvre. Once free, Surveyor 3 stabilised itself and then adopted its cruise attitude. The midcourse manoeuvre at 05:00:03 on 18 April lasted 4.3 seconds and the 13.8-ft/sec change in velocity was entirely devoted to achieving the ‘critical component’ required to reduce the divergence from the centre of the target circle from the initial 480 km down to a mere 5 km.

The pre-retro manoeuvre in which the spacecraft departed from its cruise attitude involved starting a yaw of-158 degrees at 23:23:30 on 19 April and a pitch of-76.8 degrees at 23:30:17 to align the thrust axis with the velocity vector. The final roll of -64 degrees initiated at 23:34:35 was to optimise the RADVS. The initial approach was at 23.6 degrees to the local vertical. This would require a significant gravity turn during the vernier phase of the descent to force the trajectory to vertical.

The altitude marking radar was enabled at 23:59:33, and issued its 100-km slant – range mark at 00:01:12.829 on 20 April. The delay to the initiation of the braking manoeuvre was specified as 5.090 seconds. The verniers ignited precisely on time, and the new ‘high-impulse’ retro-rocket 1.1 seconds later – at which time the vehicle was travelling at 8,618 ft/sec. The acceleration switch noted the peak thrust of 9,550 pounds fall to 3,500 pounds at 00:02:00.587, giving a burn duration of 40.0 seconds. After allowing the thrust to tail off, the casing was jettisoned at 00:02:12.429. At burnout, the angle between the vehicle’s thrust vector and velocity vector was 21.1 degrees.

When the RADVS-controlled phase of the flight began at 00:02:14.642, the slant range was 36,158 feet (and because the velocity vector at burnout was offset to vertical, the altitude was 32,900 feet) and the total velocity was 483 ft/sec (and since the vehicle had maintained its thrust along the velocity vector extant at the time of retro ignition, the longitudinal rate was 462 ft/sec). The altimeter had locked on at a slant range of 43,700 feet, only to drop out again. So when the RADVS was given control it aligned the thrust axis along the velocity vector extant at retro burnout

and flew with the verniers at 0.9 lunar gravity, very slowly accelerating as it descended. When the altimeter locked on again at 00:02:15.786, attitude control was switched from inertial to radar, and the thrust axis was swung in line with the instantaneous velocity vector to initiate the gravity turn. On intercepting the ‘descent contour’ at 00:02:33.816, the slant range was 22,300 feet and the speed was 495 ft/sec. By the 1,000-foot mark at 00:03:53.023, the vehicle was descending almost vertically with a sink rate of 103.3 ft/sec. When the 10-ft/sec mark was issued at a height of 46 feet at 00:04:10:623, it seemed to be home and dry.

But at 00:04:13.275, at a height of 30 feet, one of the three angled radar beams lost its lock on the surface. As the flow of data to the closed-loop computer abruptly ceased at 00:04:13.387, the control system reverted to its inertial guidance system to maintain its attitude and throttled the verniers to cancel out 0.9 of lunar gravity. But because the RADVS was no longer operative, it was unable to issue the 14-foot mark intended to cut off the verniers!

At 00:04:18.050 the vehicle touched down with a vertical rate of about 6 ft/sec. Although it was level at this time, the ground was sloping down to the west, causing leg no. 2 to make contact first. In response to the tilt induced by the other two legs touching down, the flight control system – which was in attitude-hold mode and did not realise that it was on the ground – increased the thrust of

verniers no. 1 and 3 to re-establish a level attitude, and this additional thrust caused the vehicle to lift off.

After peaking at about 38 feet, the vehicle made second contact at 00:04:42.030 some 50 feet west of the initial point, this time at a vertical rate of 4 ft/sec. Just as previously, the slope caused leg no. 2 to touch down first and in its effort to hold its attitude the vehicle lifted off again. The engines were cut off by a command from Earth at 00:04:53.907 – at which time the vehicle had peaked at a height of 11 feet and was at 3 feet and falling. Since a portion of the thrust had been aimed laterally at each liftoff, this had built up a horizontal component, with the result that when the vehicle struck the surface its vertical rate was only 1.5 ft/sec but it had a horizontal rate of 3 ft/sec. The elasticity in its legs caused it to rebound several inches and hop another 18 inches further downslope before it settled at 00:04:54.420, some 36 feet west of its second point of contact. The gyroscopes indicated the lander to be tilted towards the west at an angle of about 12.5 degrees from vertical.

An investigation concluded that the most likely cause of the RADVS dropping out as it neared the surface was that its logic ordered a ‘break lock’ as one of the beams crossed a field of rocks – to a microwave radar, angular rocks would have appeared much as broken mirror fragments would to a searchlight. The circuitry was designed to make the radar tracking circuits select the strongest signal if several were present. It was essential to ignore antenna ‘side lobes’ when the radar was preparing for the gravity turn. As Surveyor 3 made the final vertical descent, the scintillating side lobe had obliged the system’s logic to break its lock. As the probability of losing lock on the main beam during the vertical phase of the descent was negligible, it was decided that this problem would be eliminated on future missions by having the flight control system inhibit this side-lobe rejection logic upon receiving the 1,000- foot mark.

The first 200-line picture was received 58 minutes after landing, and a total of 55 wide-angle pictures were obtained in this mode. For this mission, a small visor had been added to the hood of the camera to prevent direct sunlight from penetrating the optical train in the hope of reducing the glare that had afflicted Surveyor 1 when the Sun was above 45 degrees of elevation. Surveyor 3 arrived approximately 23 hours after local sunrise, and in the orientation in which the vehicle landed the camera was on the eastern side with the Sun 11 degrees above the horizon. There was therefore surprise that many of the preliminary pictures were partially or completely obscured by a veiling glare. It was concluded that either engine efflux or fine particles stirred up by the engines during the ‘hot’ landings had coated part of the camera’s mirror such that when that part of the mirror was directly illuminated by sunlight the view of the lunar surface was obscured. In addition, any scene that included terrain which strongly reflected sunlight was similarly degraded. Later, intermittent sticking of the mirror in both its azimuth and elevation motions implied that dust had penetrated its mechanism. The hood rotated in azimuth with the mirror, and the mirror could be rotated in elevation to seal the hood, but the engineers had been reluctant to start off in that configuration in case the mirror failed to open. A better hood was already in development. As events would show, the camera’s operational issues would impair the imaging schedule and the glare would degrade the results. A telemetry problem meant that scanning for the Sun and Earth could not start until 06:32. This issue had appeared at the time of the second contact in the protracted arrival. It proved to be a signal processing failure. The fact that the inoperative RADVS lost its high-voltage supply at the same time implied that the signal processing problem was the result of an electrical arc. After a detailed study of the performance of the system identified a number of short circuits, a work-around was devised to minimise the impact on the surface activities. Meanwhile, the Sun and Earth acquisition was completed at 08:15, and the first 600-line picture was taken at 08:42. By handing over in succession, the Deep Space Network stations at Goldstone in California, Canberra in Australia and Madrid in Spain maintained continuous communication with the lander.

An analysis of the early pictures determined that the horizon was 5 degrees higher than it would have been if the lander were on a level plain – indicating that it was in a shallow depression. From the tilt, it was inferred to be on the eastern interior wall of a medium-sized crater.

In-flight tracking could locate the landing site only to within a few kilometres, but the crater in which the vehicle had settled was able to be identified by comparing the landscape observed at ground level with the overhead view of frame H-154 taken by Lunar Orbiter 3 – although obviously the lander was not present at the time that this picture was taken. This showed that Surveyor 3 was within 2.8 km of the aim point. The area was in frame H-125 taken by Lunar Orbiter 4 in May 1967, but because that mission was mapping from high altitude the resolution was insufficient to show the lander – nevertheless, the resulting refinement of the selenodetic grid enabled the coordinates of the site to be measured to an accuracy of better than +0.01 degree in each ordinate.

Once the crater in which Surveyor 3 landed had been found in overhead imagery, its diameter was measured at about 200 metres. It was actually the largest of a tight cluster of craters arranged in a pattern which would later be dubbed the Snowman. Photoclinometry of H-154 suggested that the crater was about 20 metres deep, that there was a smooth transition from the concave floor, that the slopes of the interior walls averaged 10 to 15 degrees, and that the rim was low and gently convex. There was an inflection in the profile from concave to convex about half way between the centre and the rim crest, in both radial and vertical directions. The ‘ground truth’ of Surveyor 3 offered a means of checking the automated photoclinometry of overhead imagery – in particular, the depth of 20 metres was seen to have been overestimated by about 5 metres.

The overhead view resolved about 100 small craters scattered over the floor, inner slopes and rim of the main crater. These ranged in size from 25 metres down to the effective limiting resolution of 1 metre. Most had gentle interior slopes and rounded rims, but a few had steep interior slopes and sharp rims. Since blocks were of higher albedo than the surface material it was possible to discern blocks down to half a metre in size, and it was evident that most were related to three of the largest craters superimposed on the main crater. By taking bearings on features and relating these to the overhead perspective, the location of the lander could be pin-pointed to within 0.5 metre – it was almost half way between the centre and the rim crest. Because it was at the inflection of the slope, its foot pads were about 7 metres below the rim

crest and about 7 metres above the centre of the cavity. In fact, the eastern rim of the crater proved to be beyond the camera’s horizon in the upslope perspective. In an exercise analogous to field surveying, the wealth of detail within the crater enabled a topographic chart of its interior to be compiled – it was an excellent example of how lunar orbiters and landers could work together. The tilt of the lander was measured by a variety of methods and estimated at 12 degrees, inclined almost due west. The fact that this was several degrees steeper than the local slope of 10 degrees was the result of foot pad no. 1, which was on the downslope side, having come to rest in a small depression.

Whereas Surveyor 1’s verniers had cut off as intended at a height of 12 feet and – as hoped – had not disturbed the surface, the fact that Surveyor 3’s verniers had kept firing through two touchdowns offered an opportunity to investigate how an intense gaseous plume affected the lunar surface material. Although the imprints of the first contact were unidentifiable owing to the highly foreshortened view of the rim of the crater and the problem with the camera made it difficult to look for erosional effects beneath the lander, the site of the second contact was conveniently positioned about

“less than 0.5 mm” which was made by tilting the camera’s mirror to view almost directly downward and taking high-resolution pictures of the surface beneath the camera.

After the success of Surveyor 1, it was decided to introduce the soil mechanics surface sampler that had been intended for later Surveyors, possibly for carriage by a rover. The original design included sensors for direct measurement of position, force and acceleration, but because the telemetry and commanding capability of the initial form of the lander could not support this complexity the position measuring system, strain gauges and accelerometers were deleted. Instead, the actions of the arm would have to be monitored by observing it using the camera, and a limited amount of data would be able to be inferred from measuring the current that the motors drew whilst in operation. The experiment comprised the articulation mechanism, the electronics compartment and their supporting structures and electrical cabling. The mechanism was mounted on the space frame immediately to the left of leg no. 2, in the position formerly occupied by the approach TV camera. The electronics compartment was at the same level and almost midway between legs no. 2 and 3. The electromechanical mechanism comprised an arm and a scoop. The arm consisted of tubular aluminium cross members which could be extended and retracted in a pantograph fashion. The scoop was a container about 13 cm long and 5 cm wide, was rigidly affixed to the arm and its door was opened and closed using an electric motor. Three electrical motors operated through drive trains to extend and retract the arm and to rotate it independently in azimuth and elevation. It was spring-loaded, extended by having an electric motor unreel a metal tape, and retracted by reeling in the tape. The arm had a maximum extension of about 1.5 metres, but was unable to access the ground immediately below its mount. It could be elevated to raise the scoop about 1 metre off the ground. On this mission the azimuth arc of the mechanism subtended 112 degrees, ranging from the left edge of pad no. 2 towards pad no. 3 – although because the legs were spaced at 120-degree intervals it stopped short of pad no. 3. In all, the sampler had an arcuate operating area of 2.2 square metres. It was controlled from Earth, but because there was no onboard memory for complex sequences it had to be operated one command at a time. The fact that it used the same radio channel as the camera meant their use had to be interleaved. The principal investigator was Ronald F. Scott, an engineer at Caltech who had worked on the Surveyor rover proposal, but the hardware had been designed and built by Hughes. Of the overall mass of 15 pounds, the mechanical assembly accounted for 8.4 pounds.

The soil mechanics surface sampler, which was also referred to as the ‘scratcher’, could manipulate the lunar surface material in a number of ways. A trench could be made by opening the door of the scoop to expose its blade, driving the scoop into the ground and retracting the arm. The scoop could hold up to 100 cubic centimetres of loose granular material, or a small rock fragment. The mechanism was designed to dig to a depth of 0.45 metre, providing that the material permitted this. An impact test involved raising the scoop and disengaging the elevation motor by releasing the clutch to allow a torsional spring to assist lunar gravity in drawing the scoop down to disturb the surface. If dropped on top of a rock, the scoop’s blade could serve as a rudimentary geological hammer. There was a 2.5 x 5.1-cm strip on the lower edge of

TV TARGET

FOOTPAD г

The configuration of the Surveyor 3 lander.

the door in order to place a flat face on the lunar surface. A static bearing test would involve placing the scoop, door closed, directly above the target and then driving the scoop down until the motor stalled, with the current providing a measure of the force applied. The arm could also be manipulated to push rocks aside in order to inspect either the underside of the rock or its imprint on the surface.

The checkout of the soil mechanics surface sampler started at 10:00 on 21 April, shortly before the end of Goldstone’s second session. After a pyrotechnic was fired to release the mechanism, JPL engineer Floyd I. Roberson commanded the arm to extend. The picture taken to confirm this showed that the arm had not advanced as far as expected. The command sequence was repeated, and the next picture showed that the arm was in the desired position. He then put it through a series of actions to verify that it could move in azimuth and elevation, checking its progress at each step by TV. This done, the arm was drawn back.

On 22 April the arm was swung to the middle of its operating area, and at 05:15 made its first bearing test of the lunar surface. The scoop was raised, the arm was swung to the right, and the scoop, door open, was driven into the surface at 09:14, after which the arm was retracted in order to scrape its first trench. Next the arm was swung left, beyond the bearing test position. After making a shallow scrape, it was raised and repositioned to make a second scrape on the same line. This time the motor stalled after just 10 cm – evidently it was more difficult to scrape an already existing trench. Meanwhile, the camera had suffered a difficulty moving in azimuth

A model of the soil mechanics surface sampler carried by Surveyor 3.

that limited its ability to support the sampling activity. Work on the second trench resumed on 23 April with a third scrape being made along the same line. Arm work was suspended on 24 and 25 April owing to the heat. Because the latitude of the site was 3 degrees, the maximum solar elevation was 87 degrees. The arm had been left at the inner end of the second trench. Having noticed what appeared to be a rock at that position, the team decided to scoop it up on 26 April, but in the process of doing so the object crumbled. The arm was swung as far right as it could traverse and the sample was deposited on the upper surface of foot pad no. 2 so that the camera could inspect the clump of fine-grained material in colour at high resolution.

On 27 April the arm swung slightly left, away from foot pad no. 2, and conducted a second and third bearing test. It then moved a little further left and scraped a third trench involving 26 retraction steps, with a wide-angle picture being taken after each step and later sequenced to produce a ‘stop-motion’ movie. On 28 April the scoop picked up a small bright object from near the most recently made trench. This was added to the material dumped onto foot pad no. 2 for inspection, but there was loose material in the scoop from the trench and on falling from the scoop this covered the white object. When the scoop was dragged across the pile to expose the object of interest, it was observed to have darkened. Next, the arm made two parallel scrapes

successively offset to the left of the third trench in order to widen it, and then a bearing test was performed on its floor. On 29 April half a dozen impact tests were conducted in an arc beyond the recent trench, with the scoop being released from a variety of heights in order to vary the force of the impact. The arm was swung to the left of its operating area on 30 April and the scoop manipulated to draw a partially buried bright object onto the surface – it proved to be a fragment of hard rock, and it was photographed in colour. On 1 May two additional scrapes were made to deepen the second trench and then the scoop was dropped four times with its door open to loosen the floor prior to a final scrape. With the Sun sinking in the west the lander’s shadow masked ever more of the arm’s operating area, so on 2 May the arm ended its operations by swinging over to the right to scrape a short fourth trench alongside the broad third trench.

The results of the arm operations indicated that the material was fine-grained and had sufficient cohesion to create loose aggregations up to several centimetres in size, although such ‘clods’ readily fell apart. When the scoop was pressed on the surface for a bearing-strength test, it left a smooth imprint which had a raised ridge of lumpy

V ‘MPACT □ BEARING О CONTACT

material around the edge. This implied that although the material was compressible, it was only moderately so, and after a certain compression the vertical force tended to displace material sideways. Impact tests were performed, but the ‘spring constant’ of the torsional spring proved insufficient to determine the density in this manner. In general, the first scrape of a trench excavated to a depth of about 7.5 cm, and each successive scrape on the same line gained an additional 5 cm – with the arm having to work harder to achieve this. Bearing tests on the floor of a trench showed that the strength of the material increased significantly at a depth of several centimetres. The deepest excavation achieved was about 18 cm, which was less than half of that for which the arm’s range of operation had been designed. Nevertheless, it provided a valuable insight into the third dimension of the enigmatic fragmental debris layer. There was no indication of textural layering in the walls of the trenches. If there was any change in the grain size, this was on a scale finer than the camera’s resolution. It simply seemed that the upper few centimetres were porous, and hence compressible, whereas the essentially similar material below was more consolidated. Its cohesivity was confirmed by the fact that the trench walls did not collapse. As in the case of the material disturbed by the foot pads, the subsurface was significantly darker than the undisturbed surface – in retrospect, it was more as if the uppermost few millimetres had somehow been lightened. From the fact that no bright angular fragments were uncovered in trenching, it was speculated that while buried they became coated with dark fine-grained material and in this darkened condition were difficult to see in a trench. By implication, it seemed that after a rock had been exposed on the surface for a time it was ‘cleaned off’ by some form of weathering. There were only a few rocks within the arm’s operating area. Most were small and partially buried. The arm picked up one rock for a close examination, but it was too small for its mass to be measured. The jaws of the scoop picked up a small white rock which was about 1.2 cm in size – a task involving 90 minutes of careful remote-control manipulation. The 100-psi pressure which the scoop exerted would have crushed a weak terrestrial rock such as a siltstone or friable sandstone, but the lunar rock remained intact.

The lesson for Apollo was that whilst the lunar material was very fine-grained, it was moderately cohesive and its bearing strength increased significantly at shallow depth.

Surveyor 3’s view was confined to the 200-metre-diameter crater in which it had landed – it could not see the plain beyond. The craters in view ranged in size from 10 cm up to 25 metres. Most of the craters that were less than 3 metres in diameter were fairly shallow, and either had very subdued raised rims or were rimless. Most of the craters between 3 metres and 12 metres in diameter were subdued, but 25 per cent had raised rims and relatively steep walls. It was apparent that most of the small craters had not penetrated beneath the fragmental debris layer within the main crater, and had merely redistributed the material that was already exposed at the surface. The size-frequency distribution was similar to that for this size range seen on the maria by Rangers 7 and 8.

The angular-to-rounded fragments ranged in size from tiny grains up to blocks of about 1.5 metres. The albedo of the undisturbed surface was 8.5 (±2) per cent, and in some cases the albedo of the blocks was one-third brighter. Although the camera

operated most effectively when the Sun was high in the sky, in such illumination the absence of shadows made subtle terrain relief almost impossible to discern – but on the other hand in such illumination it was straightforward to chart the distribution of blocks. Most blocks were relatively angular, with many wedge-shaped and some even tabular. Some of the angular rocks were partially buried, but most of the well – rounded fragments were fairly deeply buried.

In addition to the sparse and random litter of blocks, there were two prominent ‘strewn fields’ of coarse blocks. One was clearly associated with a sharp raised-rim crater about 13 metres in diameter that was embedded in the northeastern rim of the main crater, some 80 metres from the lander. The other was associated with a pair of subdued craters that were located high on the southern wall. The line of sight provided a view inside the northeastern crater, revealing its interior to be full of similar blocks. Exterior to the rim, there were radial lines of blocks. The blocks associated with this crater were the largest, coarsest and most angular in the lander’s field of view. It was evident that they were ejected by the impact that created the crater, and derived from material at a depth of 2 or 3 metres beneath the rim of the main crater. Some of the blocks had almost planar faces, as though they had broken along pre-existing joints. The tabular ones displayed grooves and ridges on their narrowest sides suggestive of lamination parallel to their longest dimension, such as would be produced in flow-banded lavas. The blocks associated with the southern craters were of similar size, but were more rounded and tended to be more deeply buried. Their source was probably the larger of the two craters there, which was 15 metres in diameter. These observations suggested that large blocks associated with subdued craters tended to be more rounded than those associated with sharp raised-rim craters, and those around subdued craters were more buried than those of sharp raised-rim craters. This suggested that freshly exposed blocks were not only eroded by the rain of meteoritic material, but also tended to be reburied as material accumulated – either by the arrival of further ejecta or as a result of downslope motion of loose debris. The fact that the rounded blocks had a pitted texture whereas the angular ones did not, implied that the pitting was caused by the same process that rounded off the angular blocks.

It was also apparent that the surface on the interior of a sizeable shallow crater on a mare plain was similar to that on a relatively level area between such craters. The 200-metre crater in which Surveyor 3 landed had probably been partially filled in by the downslope motion of material on its interior walls, thickening the debris towards the centre. Loose material piled up against the upslope sides of the larger blocks was interpreted as evidence of this process. The strewn fields of coarse blocks associated with 13-15-metre-diameter craters on or near the rim of the main crater implied that the fragmental debris layer was about 2 metres thick there. In contrast, the fact that a 20-metre crater near the centre of the main crater had not excavated blocks served to confirm that the layer there had been thickened. When the main crater was created, its floor would have been several tens of metres deeper and its rim several metres higher and sharper than it is today. By the ‘hinge-flap’ effect of an impact, the debris that formed the rim would have been excavated from the deepest point. In effect, a blocky crater on the rim of a larger crater serves as a ‘drill hole’. The crater on the

A northward-looking section of a panorama taken by Surveyor 3. The outline shows the area covered by the next illustration. (Courtesy of Philip J. Stooke, adapted from International Atlas of Lunar Exploration, 2007)

A portion of the previous illustration featuring a strewn field of boulders around a small crater on the northeastern rim of the crater in which Surveyor 3 landed. (Courtesy of Philip J. Stooke, adapted from International Atlas of Lunar Exploration, 2007)

northeastern rim would offer visiting astronauts an opportunity to recover material excavated from beneath the fragmental debris layer.2

After a detailed analysis of the Surveyor 3 imagery, Gene Shoemaker introduced the term ‘regolith’ to lunar science. This was familiar to terrestrial geologists as the collective name for the rock wastes of whatever origin and however transported that rest on bedrock and nearly everywhere form the surface of the land. On Earth, there are many erosional processes and the regolith includes volcanic ash, glacial drift, alluvial deposits, eolian deposits and soils rich in humus. On the Moon, the primary erosional process was meteoritic impact. Harold Urey had introduced ‘gardening’ for the manner in which the poorly sorted fragmental debris layer was turned over by impacts. The pictures from the two Oceanus Procellarum landing sites hinted that the thickness of the layer increased with age. To start with, the surface would have been bare rock. Any significant impact would have been capable of breaking up and scattering the material. This ejecta would have been progressively eroded by the rain of smaller projectiles. Over time, the layer of debris would have thickened, requiring ever larger impacts to reach the substrate. An important aspect of this process was that it would yield a continuous distribution of fragment sizes, which was expressed by saying that the lunar regolith was seriate.

One scientific task for Surveyor 3 was to monitor the Moon’s passage through the Earth’s shadow. With the Moon at ‘full’ phase and the Earth masking the Sun, this was a lunar eclipse to a terrestrial observer and a solar eclipse to the lander. It was the first opportunity to observe the thermal effects of such an event from the lunar surface and assess inferences drawn from telescopic studies. In particular, Surveyor science team member John M. Saari had been involved in infrared scanning of the Moon’s disk during the lunar eclipse of 19 December 1964, the data from which was processed into isothermal contours that indicated the presence of many ‘anomalies’, mostly associated with craters, where the heat that had been absorbed while the Sun was shining at lunar noon was radiated again at the onset of the eclipse. In addition to monitoring the temperature, Surveyor 3 gave the scientists a bonus: the mirror of its TV camera had the same 35-degree elevation limit as previously, but because the lander was west of the meridian and inclined due west at an angle of 12 degrees by virtue of having settled on a slope, the field of view of the wide-angle frame was just able to include Earth, east of the zenith. Optical observations of the eclipse would yield the first direct measurement of the distribution of the refracted sunlight which weakly illuminates the lunar disk at such times.

Fortunately, the eclipse on 24 April occurred during a period when the Moon was still just above Goldstone’s horizon. A total of 20 images were taken in two sets: the first at 11:24 and the second 37 minutes later. They were taken at two iris apertures, and with several exposures for each of the three colour filters. In the first set, an arc along the northwestern limb of Earth refracted light that varied greatly in brightness, and with a fainter glow at each end containing bright ‘beads’. In the second set, the

And in fact Apollo 12 would do so at precisely this spot.

The locations of lunar transient events reported over the years by various observers and infrared ‘hot spot’ anomalies measured during a lunar eclipse on 19 December 1964. (Courtesy of John M. Saari and R. W. Shortfall, Isothermal and Isophotic Atlas of the Moon, NASA, 1967)

brightest refraction had migrated to the northeastern limb. To assess the distribution of cloud on the limb, the geographic coordinates of the ‘beads’ were later calculated and compared to pictures taken by the ESSA 3 meteorological satellite in low polar orbit on the day prior to the eclipse. In some areas cloud in the troposphere occulted some of the refracted sunlight, but the bright ‘beads’ occurred where sunlight passed through regions free of cloud. Of course, the refracted sunlight made it impossible to view the much fainter solar corona.[34]

The thermal properties of the lunar surface inferred from the lander’s temperature data differed from inferences made from telescopic studies, in that the in-situ data showed a higher thermal inertia. However, the data was provided by sensors in place to monitor the thermally controlled compartments, not by instruments specifically designed to study the thermal properties of the lunar surface, and therefore was too crude to draw definitive conclusions.

The plan had called for attempting a ‘liftoff and translation’ manoeuvre by firing the verniers, but this was ruled out by thermal factors. The decision rested upon the temperatures of the thrust chambers of the engines, the flight control electronics, the helium tank, the shock absorbers and the roll control actuator on vernier no. 1 – all of which depended upon solar heating and shadowing in the orientation in which the vehicle came to rest. The key issue was the temperatures of the engines. Irrespective of the orientation, by the time the elevation of the Sun reached about 35 degrees one or other of the engines was sure to exceed its permitted pre-ignition temperature of 105°C. Hence, any attempt to lift off had to be made either early in the morning or late in the afternoon. For Surveyor 3, the thermal situation was complicated by the fact that the lander was on a slope, which significantly altered the manner in which shadows were cast. To preserve the option of this experiment, the helium tank had not been vented. Also, given that at no time during the protracted landing had the forces on the shock absorbers imparted even half the load endured by Surveyor 1, it had been decided not to lock the legs. Whilst the helium was significantly depleted, as the tank absorbed solar heat its pressure increased to 2,735 psia.[35] By the time of the eclipse, however, it had been decided not to attempt to fire the verniers. The legs were monitored to determine whether the rapid decline in temperature at the onset of the eclipse prompted the shock absorbers to leak, but they retained their integrity. At 20:36 on 24 April, the helium tank was finally vented.

Surveyor 3 had two flat beryllium mirrors situated to enhance the camera’s view of the underside of the vehicle. That is, the camera was between legs no. 2 and 3 and the mirrors were affixed to leg no. 1. One mirror was 35 x 22 cm and gave a view of the lower portion of crushable block no. 3 and the area beneath vernier no. 3. The other mirror was 9 x 33 cm and viewed the area beneath vernier no. 2. The fact that the verniers were cut off at a height of only 3 feet suggested there might be signs of surface erosion, but since the vertical velocity at the third contact was only 1.5 ft/sec the crushable blocks probably did not strike the surface. Unfortunately, the hopes of viewing beneath the lander were foiled by the fact that when the Sun was low in the sky the pictures were ‘washed out’ by the glare from the coating on the main mirror of the camera, and when the Sun was high in the sky the area of interest was in the lander’s shadow! It had been hoped to fire the downward-pointing cold-gas thruster on leg no. 2 to follow up on Surveyor 1’s surface erosion experiment, but this time obtaining good ‘before’ and ‘after’ pictures of the test area. However, pictures taken when the Sun was low in the east were washed out by glare, pictures taken when the Sun was high lacked the shadows required to highlight the minute changes likely to result from such a weak thrust, and in the late afternoon the area was in the lander’s shadow – so the test had to be cancelled.

Between 10:29 and 11:06 on 30 April, Surveyor 3 snapped wide-angle pictures of Earth illuminated as a crescent with the dawn terminator running the length of South America. These were the first colour pictures of Earth taken from deep space.[36] The filters had been revised to provide an improved spectral match to standard colorimetry functions.

Following sunset at 18:38 on 3 May Surveyor 3 monitored the rate at which the temperature fell, and at 00:02 on 4 May was commanded to hibernate.

In total, the camera took 6,326 pictures. Owing to the glare from contamination of the main mirror, usable images could be obtained only over a limited azimuth range during the early morning and late afternoon. This glare, combined with the difficulties in moving the mirror, made it impossible to obtain all of the systematic surveys of the landscape which had been planned. About 8 per cent of the pictures were taken at wide-angle to provide panoramas at specific illumination phases. The glare impaired detailed photometry, but the colorimetry confirmed that the surface was essentially grey. The glare precluded photographing the stars for use as celestial references to precisely determine the orientation of the camera, and hence the true orientation of the lander, but on one occasion it did manage to photograph Venus, which helped to some extent.

Although Surveyor 3 was unable to view the plain surrounding the crater in which it landed, scientists were delighted to have the opportunity to survey the interior of a medium-sized crater on a mare! In effect, therefore, its observations complemented those by Surveyor 1 of the open plain. In general, the character of the lunar surface material appeared to be similar at the two sites. The soil mechanics surface sampler was active for a total of 18.3 hours. It executed 5,879 commands, during which it made seven bearing-strength tests, thirteen impact tests and four trenches to provide data on the strength, texture and structure of the lunar material to a depth of 18 cm. In particular, it found that the bearing strength of the material increased with depth, even although there was no discernible change in the grain size – it was just that the uppermost few centimetres were more porous. Although sequences of commands to enable the sampler to perform complex operations had been stored on magnetic tape for step-by-step uplinking, for much of the time it was actually operated in real-time and monitored by TV.

Surveyor 3 evidently succumbed to the chill of the lunar night, because attempts to reactivate it after sunrise were unsuccessful.

Robert Seamans resigned as NASA’s Deputy Administrator on 2 October 1967, and departed 3 months later. On 5 February 1968 the Senate accepted the nomination of Thomas O. Paine, a senior manager of the General Electric Corporation, and he was sworn in as Deputy Administrator on 25 March.

AS-502 was the second in the series of ‘A’ missions designed to ‘man rate’ the Saturn V launch vehicle. The payload was LTA-2R and CSM-020, a Block I with some Block II modifications for certification, including a heat shield with an actual unified crew hatch to be tested in lunar return conditions. In view of the fact that the mission would be unmanned, certain systems had been deleted from the command module in order to accommodate an electromechanical control sequencer.

The countdown proceeded without unplanned holds, and Apollo 6 lifted off from Pad 39A at 12:00:01 GMT on 4 April 1968. It contrast to the perfect performance of the first Saturn V, this one suffered a number of problems.

Firstly, between T+110 and T+140 seconds it underwent a ‘pogo’ oscillation, as a longitudinal structural mode frequency coupled with the resonant frequency of the oxidiser lines feeding the S-IC’s engines. The greatest disturbance was in the range 5.2 to 5.5 hertz. The oscillations in the engine chamber pressures built up to a peak – to-peak maximum of 8 to 10 psia at T+ 125, and the condition was reinforced by the consequent variation in thrust. The +0.6-g oscillation measured in the spacecraft exceeded the design criteria, and would have given astronauts a very rough ride. The emergency detection system cast one ‘vote’ for terminating the mission. Had it gone on to cast a second vote, an abort would have become mandatory and the command module would have been drawn clear by the launch escape system.

Ground-based and airborne cameras noted three small pieces and five or six large pieces separating from the vicinity of the SLA between T + 133.31 and T + 133.68, at which time strain, vibration and acceleration sensors in the S-IVB, IU, SLA, LTA and CSM reported abrupt changes. Subsequent analysis determined that one of the SLA panels had suffered structural failure and shed some of its skin. Fortunately, the supporting elements were able to sustain the loads for the remainder of the powered flight.

The S-IC shut down at T+ 148.21 and separated cleanly. The five J-2 engines of the S-II ignited at T+150 and performed satisfactorily for 169 seconds, but at T+319 the hydrogen flow rate to engine no. 2 suddenly increased and its thrust declined by 23 psi. The engine ran at this level until T + 412.92, whereupon the temperature in the engine bay suddenly rose and the engine shut down. Engine no. 3, which had shown no sign of distress, cut off 1.26 seconds later.

Despite the loss of 40 per cent of its power, the S-II was able to gimbal the three remaining engines to hold itself stable. However, the controller in the Instrument Unit had been configured to react only with a single-engine-failure contingency and was unable to take into account the loss of the second engine. It naively attempted to recover its trajectory as if it had four good engines. When the IU started to adjust the propellant mixture ratio to optimise consumption leading up to S-II shutdown, it also adopted a mode designed to stabilise the vehicle for separation. If all had been going to plan, at that time the S-II would have been within 5 seconds of cutoff but the loss of two engines reduced the rate of propellant consumption and the other engines had to burn for much longer than planned to trigger the fuel-depletion cutoff. In fact, the burn was 57.81 seconds longer than usual, and because the vehicle held its attitude fixed during this time it climbed higher than intended and the space – fixed velocity at cutoff was 335.52 ft/sec less than nominal.

The inherited trajectory anomaly presented the S-IVB with a serious challenge. It ignited at T + 577.28 and burned for 166.52 seconds, some 28.95 seconds longer than nominal. On finding itself high, slow and short, it pitched down 50 degrees in order to lose altitude, accelerate and gain range. On achieving the desired altitude it raised its nose above the local horizon to overcome the negative radial velocity acquired in descending, whilst simultaneously minimising further increasing its horizontal rate. On activating terminal guidance at T + 712.3, the system set the altitude constraints to zero and focused on achieving the desired velocity. Although it ordered the vehicle to pitch up beyond vertical and travel backwards, the 1 deg/sec rotation rate meant that when the space-fixed velocity exceeded the objective and the stage shut down at T+747.04 the angle had reached only 65 degrees. The velocity was 160 ft/sec greater than nominal, the surface range was 269.15 nautical miles longer than nominal and the flight path angle was slightly negative. A circular orbit at 100 nautical miles had been planned, but the actual orbit of 93.49 x 194.63 nautical miles did not preclude continuing with the mission.[50]

Once the S-IVB had realigned its axis with the horizon it initiated a sequence of manoeuvres to be undertaken on the first revolution. First, it rolled 180 degrees and pitched down 20 degrees, and then it pitched up 20 degrees and rolled to resume its original attitude. The only appreciable effect was the sloshing of liquid oxygen at the onset of each change in pitch, but this was rapidly damped out. This qualified such manoeuvres to orient the vehicle to enable astronauts to perform landmark tracking while in parking orbit.

The Apollo 6 plan was to reignite the S-IVB at the end of the second revolution to simulate the translunar injection. Although this was to achieve lunar distance, the apogee was to be away from the Moon in order not to complicate the evaluation of the guidance system in deep space. Immediately after cutoff, the S-IVB was to pitch through 155 degrees and release the CSM in an attitude appropriate for a retrograde burn. A 254-second burn by the service propulsion system would put the spacecraft into an ellipse with an apogee reduced to about 12,000 nautical miles and with an atmosphere-intercepting perigee. The vehicle was to coast in an attitude that would ‘cold soak’ the heat shield to approximate lunar return thermal conditions. Late in the descending portion of the ellipse, the spacecraft would fire its engine again to accelerate to 36,500 ft/sec with a flight path angle of-6.5 degrees to simulate lunar return, jettison the service module and orient itself for atmospheric entry leading to splashdown in the Pacific.

Despite the additional propellant consumed in attaining orbit, the S-IVB was still capable of performing the planned program. However, the restart system failed to produce hydraulic pressure. On noting this failure, the IU cancelled the ignition command at 003:13:50.33 and advanced to the next programmed item – just as if it had achieved the burn. At 003:14:10.33 it initiated the 155-degree pitch manoeuvre, and 15 seconds later was still rotating when the ground commanded the spacecraft to separate. The hinges had been designed to rotate the SLA panels to 45 degrees from the vehicle’s longitudinal axis within 1.3 seconds, but in this case the S-IVB was still in the process of pitching and the fact that the spacecraft suffered a disturbance in pitch of 1.5 deg/sec during separation implied that one of the panels had nudged the spacecraft. A preplanned alternative mission was selected in which the service propulsion system would fire to obtain the desired apogee of 12,000 nautical miles, although the propellant required to achieve this would rule out the follow-on burn to accelerate to lunar return velocity. As planned, the spacecraft adopted an attitude to ‘cold soak’ its heat shield for 6 hours. At high altitude, a 70-mm camera snapped 370 colour pictures of Earth in daylight, and instruments monitored how efficiently the command module’s wall blocked the charged-particle radiation circulating in the van Allen belts.

On jettisoning the service module, the command module oriented itself for entry, making contact at 009:38:29 at an inertial velocity of 32,830 ft/sec (10 per cent less than intended) on a flight path angle of-5.85 degrees. It splashed down 49 nautical miles short of USS Okinawa, on station at the recovery point for the simulated lunar return. For the first time, a capsule adopted the apex-down flotation attitude, but was promptly righted by a set of airbags that inflated on its nose. It was recovered when the ship arrived 6 hours later.

Even while CSM-020 was coasting in space, NASA set up a meeting at the Cape with the contractors to investigate the problems suffered by the upper stages. Within 24 hours a review of the recorded telemetry established that:

• 70 seconds after S-II ignition, sensors in the engine compartment began to report chilling and the flow of liquid hydrogen to engine no. 2 increased;

• at +110 seconds the thrust of this engine began to decline gradually, and then dropped sharply at +169 seconds, at which time the load on the mechanism that gimballed the engine suddenly increased to counter a lateral component of thrust;

• at +263 seconds the thrust fell sharply again, there was a sudden increase in temperature in the engine compartment, the automatic system to shut down an engine if its thrust fell below a specific value intervened; and

• one second later engine no. 3, which had shown no anomalous behaviour, suddenly cut off!

It was soon realised that engine no. 3 had cut off because it received the command intended for its ailing sibling. It was inferred that the control wires for the solenoids of the liquid oxygen prevalves of engines 2 and 3 must have been erroneously cross­connected. A check of the records showed that the wiring had been modified several months previously. Thus the intervention had shut down engine no. 2 by cutting off its fuel supply and shut down engine no. 3 by cutting off its oxidiser.

The telemetry from the S-IVB showed that its J-2 had also behaved anomalously by starting to chill at +68 seconds, the thrust starting to decline at +107 seconds, the rate of decline increasing at +115 seconds, and then the engine compartment temperature increasing at +119 seconds. However, in this case the engine continued until it was shut down nominally at +170 seconds.

Thus two engines on different stages had misbehaved in a strikingly similar way: namely, the onset of chilling in the engine compartment about 70 seconds into the burn, the thrust starting to decay about 40 seconds later, and then a sudden increase in temperature – although at different times.

The chilling implied a propellant leak – and the fact that a temperature colder than liquid oxygen was measured indicated a leak of hydrogen somewhere in the feed to the engine. The way in which the engine compartment of the S-II had been chilled supported the case for a hydrogen leak – that is, it had affected engine no. 5, which was in the centre of the cluster, but the engine on the opposite side of the cluster was unaffected. The structure of the engine and its orientation within the cluster implied that the leak was associated with a stainless steel pipe about 1 inch in diameter that wound its way ‘upwards’ from the middle of the assembly. This pipe carried liquid hydrogen to the ignition system, a small chamber in the middle of the injector plate in the roof of the combustion chamber. Another pipe delivered liquid oxygen. Spark plugs ignited this mix and issued a jet of flame into the chamber to start and sustain main combustion. The hydrogen pipe incorporated three flexible bellows to permit movement, with one bellows located precisely where the leak was inferred to have occurred. Evidently, the leaking pipe sprayed liquid hydrogen into the engine bay, chilling it. This flow created the lateral component of the thrust which the gimbal counteracted. The reduced flow of hydrogen to the igniter made the flame from the igniter oxygen-rich and turned it into a ‘torch’ which so eroded the structure as to let blazing gases enter the engine compartment, heating it. Rocketdyne rigged an engine at the Santa Susanna Test Laboratory in California to replicate the leak, and within a month was able to reproduce the behaviour of engine no. 2.

In the case of the S-IVB it was concluded that when the feed pipe broke, it did so completely, and with no hydrogen reaching the igniter there was no ‘torch’ to erode the structure. However, gas was able to pass from the combustion chamber through the igniter and out of the severed pipe to cause external heating. The cryogenic leak froze the hydraulic fluid, with the result that during the restart attempt both the main and auxiliary hydraulic pumps cavitated and yielded essentially no system pressure. In any case, with no hydrogen reaching the igniter there was no way that the engine could have been ignited even if everything else had worked.

Within a month, therefore, the investigators knew what had happened to the J-2s, but not why. To discover why the pipe was vulnerable, the telemetry from AS-501

was compared with that from AS-502. It was noted that on AS-502 the power of the engines was greater, and that the liquid hydrogen pressures in the pipes would have been marginally greater. Tests by Rocketdyne established that the bellows section vibrated at the increased flow rate, but even on a prolonged firing remained intact. The entire engine was vibrated prior to ignition to simulate the pogo of the S-IC in case this had been a factor, but the engine fired as before with the bellows vibrating. After several weeks, it was decided to vibrate the engine while passing fluid at the increased rate – but for safety using gaseous nitrogen at room temperature instead of liquid hydrogen. As the flow rate was being adjusted to the value desired, and prior to the vibration being introduced, the bellows failed. The same thing happened to a replacement. An inspection indicated metal fatigue arising from vibration induced by the flow.

Attention then turned to the environmental factors. The flight engine had been in near-vacuum in the upper atmosphere. The firing tests had been at sea level where, with liquid hydrogen at -253°C, a coating of ice would have formed on the bellows and tended to dampen out the vibration induced by the flow. Rocketdyne put an engine into an altitude chamber and pumped liquid hydrogen through the pipe, and the bellows failed after 100 seconds. In retrospect, the mystery was not that failures had occurred in flight; it was that the other engines had not succumbed to the same problem! Regardless, the fix was to use pipes which incorporated bends capable of absorbing the movements that the flexible bellows had been intended to counter.

Meanwhile, the pogo had been overcome by modifying the prevalves in the liquid oxygen ducts of the S-IC engines to incorporate a cavity containing helium which, by compressing, would dampen pressure fluctuations and thereby maintain a smooth combustion.

An analysis of the structural failure of one of the SLA panels found that this was unrelated to the pogo. Aerodynamic heating had increased the pressure of moisture inside the aluminium honeycomb material sufficiently for part of the facing sheet to puncture and tear off in the slipstream. It was decided to apply a layer of cork to the exterior of the adapter to absorb moisture, and to drill holes to prevent a build up of pressure within the underlying honeycomb.

Sam Phillips said this post-flight investigation was “one of the most aggressive, thorough and determined engineering test and analysis programs I have ever seen”.

In fact, AS-502 was a successful test flight precisely because it revealed problems which were then fixed.

Notwithstanding the problems suffered by AS-502, on 23 April George Mueller called for AS-503 to be manned. Although Sam Phillips directed the next day that this launch vehicle be prepared with CSM-103 and LM-3 for a manned mission, he also instituted contingency planning to reconfigure it with BP-30 and LTA-B in the event of the decision to undertake a third unmanned test. The Kennedy Space Center replied that, given sufficient notice of the configuration, the ‘boilerplates’ would be able to be launched in mid-October but the manned mission would not be ready until late November at the earliest. On 26 April James Webb approved this planning for a manned mission, subject to a resolution of the anomalies which afflicted AS-502. In seeking to overcome the pogo, the Marshall Space Flight Center asked whether the

emergency detection system could be configured to trigger an abort automatically. When Deke Slayton argued against doing so, George Low ordered the development of a ‘pogo abort sensor’ with a display in the command module to enable the crew to judge whether to initiate an abort. On 17 August, by which time it was clear that the pogo would be able to be eliminated, Low recommended that work on this sensor be terminated and a week later Phillips concurred.

Meanwhile, when the fuel injector of the ascent propulsion system developed by the Bell Aerospace Company for the LM continued to suffer combustion instability into the summer of 1967, NASA hired Rocketdyne to develop an alternative injector as a contingency measure. In April 1968 Grumman was instructed to coordinate the testing of Rocketdyne’s injector in Bell’s engine. In May George Low decided that this hybrid should be used, and told Rocketdyne to perform the integration work. By mid-August it was clear from qualification testing that the modified engine was free of instabilities, and there would be no need to mount another test flight. On 13 May Low met Chris Kraft, Deke Slayton and Maxime Faget to consider whether the ‘fire in the hole’ staging demonstration should be on the ‘D’ or ‘E’ mission. A key factor was that LM-3 would be the last to have the development flight instrumentation for monitoring the systems. Faget argued that whilst such data was desirable, it was not essential – it would be sufficient to take photographs of the base of the ascent engine following the rendezvous. In view of this line of argument, and the fact that the ‘fire in the hole’ staging demonstration would increase the complexity of the ‘D’ mission, Low postponed making a decision on whether to do it on the ‘E’ mission until the performance of the engine on LM-1 had been thoroughly analysed. On 17 May Kraft advised Low that the ‘E’ mission was already a complex affair, and that as further objectives were added the probability of achieving them diminished. In fact, Kraft saw little need for a ‘fire in the hole’ test. He understood the engineers’ desire to test all the systems in space in both normal and backup modes, but the first ‘fire in

the hole’ test at the White Sands Test Facility on 22 December 1967 had achieved all its objectives and further ground testing would provide the data needed to calculate the pressure and temperature transients pertaining to lunar lift off.5 In parallel with these discussions, the Manned Spacecraft Center was studying extending the apogee of the ‘E’ mission to almost lunar distance to investigate navigation, communica­tions and thermal control issues in the event of the lunar orbital ‘F’ mission being deleted, and this alternative mission was labelled ‘E-prime’.

On 7 May 1968 CSM-101 passed its final customer acceptance review, and at the end of the month was delivered to the Cape. The inspectors were delighted to find fewer discrepancies than on any previous spacecraft. But Wally Schirra, who would be in command of flying the vehicle, did not accept it as flightworthy until after its altitude chamber tests in June. In contrast, when LM-3 was delivered on 14 June the inspectors found over 100 deficiencies, many of which were classified as major. In July, George White, the Chief of Reliability and Quality Assurance in the Office of Manned Space Flight, briefed George Mueller on the issues the Certification Review Board would require to consider.

On 7 August George Low advised the Manned Space Flight Management Council that the delivery of CSM-103 was imminent, but LM-3 was unlikely to be ready for launch until February 1969.

Low felt that for Apollo to have a chance of achieving a lunar landing in 1969, the first manned Saturn V must be flown in late 1968. By this point, the pogo problem was heading towards resolution and the other issues that marred AS-502 had been fixed. In April 1967 the Manned Spacecraft Center had outlined a contingency for a lunar mission involving only the CSM. On 8 August Low asked Kraft to consider sending CSM-103 to the Moon, and then he flew to the Cape with Carroll Bolender, ASPO’s LM Manager, Scott Simpkinson, Chief of ASPO’s Test Division, and Owen Morris, ASPO’s Chief of Reliability and Quality Assurance, to discuss AS-503 with Sam Phillips, Kurt Debus, Director of the Kennedy Space Center, Rocco Petrone, Director of Launch Operations and Roderick Middleton, the Apollo Manager at the Cape.

At 08:45 local time in Houston on 9 August, Low met Kraft and Robert Gilruth to discuss the CSM-only option. Kraft said that it was feasible and Gilruth was enthusiastic. At 09:30 Deke Slayton was called in, and offered his support. Low then telephoned Sam Phillips, who had remained in Florida, and a meeting was arranged at 14:30 in Huntsville. In attendance were Low, Gilruth, Kraft and Slayton from the Manned Spacecraft Center; Wernher von Braun, Eberhard Rees (his deputy), Lee James (his Saturn V Program Manager) and Ludie Richard from the Marshall Space Flight Center; Kurt Debus and Rocco Petrone from the Kennedy Space Center; and Phillips and George Hage (his deputy) from headquarters. Low opened by saying that if the Apollo 7 evaluation of CSM-101 went well, it would be technically possible to send CSM-103 out to the Moon in December. However, if CSM-101 had

In the event, no ‘fire in the hole’ separation was demonstrated in space.

problems then it would be necessary to confine CSM-103 to Earth orbit to continue the evaluation of the spacecraft’s systems. Kraft pointed out that a mere ‘loop’ around the far-side of the Moon and then back to Earth would be insufficient – the spacecraft would have to enter orbit in order to contribute significantly to the lunar landing mission (which, depending on whether the ‘F’ mission was undertaken, might be the next mission to venture to the Moon). There was general agreement to initiate the planning for this contingency in secret, pending the decision and public announcement. The meeting broke up at 17:00. On returning to Houston, at 20:30 Low briefed George Abbey (his technical assistant), Kenneth Kleinknecht (his CSM Manager), Carroll Bolender (his LM Manager) and Dale Myers (Apollo Program Director for North American Aviation).

On 10 August Slayton offered this ‘new’ mission to James McDivitt, who was earmarked to fly AS-503, but McDivitt opted to await LM-3 so as to keep the ‘D’ mission for which his crew had been training. In contrast, Frank Borman, who had eagerly followed the discussions to extend the apogee of his ‘E’ mission out to lunar distance, readily accepted, and thereby regained the first manned Saturn V mission.

Kraft told Low on 12 August that a daylight launch would be required to allow an Atlantic recovery after an abort, and this meant the lunar launch window would open on 20 December. Low selected LTA-B as a stand-in for the LM because it had been assigned to fly with BP-30 on the unmanned mission and was already in preparation.

On 14 August the original twelve conferees, minus Rees, were joined at a meeting at NASA headquarters by Deputy Administrator Thomas Paine, Julian Bowman and William Schneider (the latter both of the Office of Manned Space Flight) to make a formal recommendation. With the discussion in progress, George Mueller telephoned from Vienna in Austria, where he and James Webb were at a United Nations Conference on the Exploration and Peaceful Uses of Outer Space. Mueller was sceptical of the proposal, and said he would not be able to discuss it until 22 August. Paine, playing the devil’s advocate, pointed out to the conferees that until recently there had been doubts about whether the Saturn V was safe for manned flight, and here they were considering having it send a spacecraft on an impromptu mission to orbit the Moon, then he invited comments. Von Braun pointed out that once it was decided to man AS-503, it did not matter how far the spacecraft went. Hage noted that there were a number of points in the mission where go/no-go decisions could be made, managing the risk. Slayton opined that not to pursue this option would significantly diminish the likelihood of achieving President Kennedy’s deadline. Debus had no technical reservations about the launch. Nor did Petrone. Bowman said it would be a ‘shot in the arm’ for the program. Lee James said it would enhance the safety of later flights. Ludie Richard said it would improve lunar capability. Schneider was fully in favour. Gilruth pointed out that although it was an impromptu mission, it would improve the chance of being able to achieve the overall goal of the program. Kraft reiterated that it should be a lunar orbital rather than a circumlunar mission. Low pointed out that if Apollo 7 succeeded, they could either fly this impromptu mission or await LM-3 and launch in February or March. In view of the deadline for the lunar landing, Low said the decision was obvious: they should send Apollo 8 to the Moon. Paine concurred. Phillips ordered planning to continue.

Phillips and Paine discussed the plan with Mueller and Webb on the telephone the next day. Mueller had warmed to the idea overnight. Webb was “fairly negative” (as Phillips later put it) but asked for information to be sent by telegram. On 16 August Webb called Paine and agreed to the mission planning, with the proviso that there be no public announcement. On 17 August Phillips told Low that although Webb had authorised a manned Saturn V launch in December, there must be no ‘leak’ that the spacecraft might venture to the Moon – that decision was contingent on the outcome of Apollo 7.

Meanwhile, CSM-103 had arrived at the Cape and a start had been made on the modifications required to send it to the Moon. On 19 August Phillips directed that if AS-503 was manned and did not carry a LM, irrespective of where the spacecraft went the mission was to be designated ‘C-prime’. Then McDivitt’s crew would fly the ‘D’ mission riding AS-504 with CSM-104 and LM-3. The ‘E’ mission had been deleted. That same day, Phillips told the press that if the mission of CSM-101 was a success, AS-503 would be manned and that because the LM would not be ready this would be a CSM-only mission; by not mentioning the option of leaving Earth orbit he readily conveyed the impression of an Earth orbit mission. On 3 September Low directed that if the ‘C-prime’ mission went to the Moon, it would make ten orbits over a period of 20 hours and then head home. If it was confined to Earth orbit, it would undertake the parking orbit preparations for translunar injection and then fly one of a number of alternative missions, each of which would involve simulating the transposition, docking and extraction of the LM. On 9 September Borman’s crew began to train in the simulator at the Cape for the lunar mission. Ten days later, with the AS-502 investigations finished and the remedies implemented, Mueller declared the Saturn V to be ‘man rated’. CSM-103 was mated with AS-503 on 7 October, the launch escape system was added on 8 October, and the next day the space vehicle was rolled out and installed on Pad 39A.

On 31 March 1968 Lyndon Johnson announced that he would not seek and would not accept his party’s nomination for the presidential election in November. The two main candidates were Hubert H. Humphrey and Richard M. Nixon. Perhaps because James Webb knew that neither would retained him as NASA Administrator, he informed Johnson on 16 September that he would stand down on 6 October, which was his 62nd birthday. Thomas Paine was promoted to Acting Administrator.

Meanwhile, on 20 September CSM-101 passed its flight readiness review, and later that day Wally Schirra announced to reporters that Apollo 7 would be his final mission because he intended to retire from NASA.

The first International Polar Year was held between 1882 and 1883 to coordinate meteorological, magnetic and auroral studies. The eruption of Krakatoa in

Indonesia on 20 May 1883 had a temporary but significant effect on the atmosphere. A second International Polar Year was held 50 years later. In 1950 the International Council of Scientific Unions proposed to exploit the technologies developed in the years since the Second World War to undertake geophysical research on a global basis to study the solar-terrestrial relationship. In early 1952 it was agreed that this International Geophysical Year would run from July 1957 to December 1958, a period which was expected to coincide with the time of maximum solar activity in the 11-year cycle of sunspots. In early 1954 the National Security Council said the US “should make a major effort during the International Geophysical Year”, and directed the Pentagon to provide “whatever support was necessary to place scientists and their instruments in remote locations” to make observations.

In August 1953 physicist Fred Singer outlined to the International Congress of Astronautics a 45-kg satellite for MOUSE (Minimum Orbital Unmanned Scientific Experiment). He spent the next year promoting it. In October 1954 he canvassed the US delegation to the meeting in Rome, Italy, of the International Geophysical Year’s Steering Committee, and as a result a resolution was passed which encouraged participants to investigate the possibility of launching a satellite as the highlight of the program.

In November 1954 Charles Wilson told journalists he did not care if the Soviets were first to put up a satellite. Despite the National Security Council directive for “a major effort’’ in support of the International Geophysical Year, it was not until 1955 that Wilson endorsed a satellite. In July 1955 Eisenhower announced that the US would put up a satellite for the International Geophysical Year. Eisenhower saw it as a one-off scientific venture. He assigned to the Pentagon the decision for how it should be achieved. There was intense rivalry between the services, because such a spectacle would boost that service’s claim to be assigned a greater responsibility for long-range missiles. Shortly before Eisenhower’s announcement, Donald Quarles, Chief of Research and Development at the Pentagon, had set up a committee chaired by Homer Joe Stewart, a physicist at the University of California at Los Angeles, to review the capabilities of the services. The National Security Council had stipulated that the satellite must not impede the development of the Atlas missile, which was only now beginning to gear up as a ‘crash’ national program. This ruled out the Air Force.

The Army proposed Project Orbiter, claiming that if the Redstone missile, which was an improved V-2, were to be fitted with three upper stages, a satellite would be able to be launched by January 1957, which was before the start of the International Geophysical Year. The Navy had Project Vanguard, in which an improved form of the Viking ‘sounding’ rocket introduced in 1949 for stratospheric research would be augmented with two upper stages. Part of the rationale for the Stewart Committee selecting Vanguard was the perceived greater reliability of requiring only two upper stages, instead of three. In addition, the Committee was impressed by the in-line configuration of the Vanguard stages, as opposed to clustering small solid rockets to form the upper stages of the Redstone launch vehicle. Nevertheless, Stewart himself had supported the Army’s proposal. One factor was that whereas the Redstone was a weapon and was classified, the Viking was not classified. Another rationale, added later, was that it would be better to use a ‘civilian’ rocket for this scientific project. The Committee was not concerned that Vanguard would not deliver as early as the Army claimed for Orbiter – it was simply presumed that the first satellite would be American, and provided that it was launched within the period of the International Geophysical Year it would serve its purpose. On 9 September 1955 the Pentagon endorsed the Committee’s recommendation. The spherical Vanguard satellite would weigh about 1.5 kg, and would transmit a radio signal that would allow the study of electrons in the ionosphere and thus make a unique contribution to the International Geophysical Year.

Since the services were only loosely controlled by the Department of Defense, the Army set out to contest the decision, emphasising that the Redstone could launch a satellite without impeding military work. When on the Stewart Committee, Clifford C. Furnas of Buffalo University had sided with the Army. Now at the Pentagon, he advised the Army to have its missile ready as a backup in case Vanguard faltered.

On 1 February 1956 the Army Ballistic Missile Agency was established at the Redstone Arsenal, Major General John B. Medaris commanding. It was to develop an intermediate-range ballistic missile named Jupiter. As the warhead would enter the atmosphere at a faster speed and be subjected to greater heating than that of the short-range Redstone, it was decided to test the new re-entry vehicle by firing it on a ‘stretched’ Redstone equipped with two upper stages made by clustering small solid rockets. The fact that this ‘Jupiter-C’ would enable the Army to develop and test a vehicle capable of launching a satellite was, of course, entirely coincidental! When the first test flight on 20 September 1956 reached a peak altitude of 1,000 km and flew 4,800 km down the Air Force’s Eastern Test Range from Cape Canaveral, the Pentagon directed Medaris to personally guarantee that Wernher von Braun did not inadvertently place anything into orbit! One criticism of Vanguard was that although its first stage was based on the Viking, the project really involved developing a new integrated vehicle in a period of only 2 years. With Vanguard running late, Medaris sought permission to launch a satellite, but the Secretary of the Army refused – in fact, the Army Ballistic Missile Agency was ordered to destroy the remaining solid rockets obtained for the upper stages. In response, Medaris decided to leave them in storage to ‘assess’ their shelf life!

In public, Eisenhower maintained that launching a satellite was a one-off venture for the International Geophysical Year. In fact, this was a cunning ruse, because the aim was to use Vanguard to set the precedent of a US satellite passing over foreign territory, and thus preclude a legal challenge when the US began to send up satellites for military functions such as reconnaissance.

Soon after the US announced that it would launch a satellite for the International Geophysical Year, the Soviet Union said it intended to do the same. In mid-1957 the Soviet magazine Radio told its readers how to go about ‘listening’ to this satellite. In late August the TASS news agency announced the successful test flight of a ‘‘super long range’’ missile which was capable of striking ‘‘any part of the world’’. When a Soviet delegate at an International Geophysical Year meeting in Washington in late September was asked whether the promised satellite was imminent, he replied: ‘‘We won’t cackle until we’ve laid our egg.’’ In other words, wait and see!

On 4 October the 84-kg Sputnik was placed into an orbit which ranged in altitude between 220 and 950 km and transmitted its incessant ‘beep, beep, beep’ signal.

The news caused a world-wide sensation, but Eisenhower was not concerned. At a press conference on 9 October he dismissed Sputnik as a ‘‘small ball in the air’’ that ‘‘does not raise my apprehensions, not one iota’’. On the other hand, the mass of the satellite showed the capability of the Soviet intercontinental-range ballistic missile, and Eisenhower ordered an end to the administrative difficulties that were impeding funding for the American missile programs.

Lyndon Baines Johnson was not only the senior Democratic senator for Texas, as the Democratic leader in the Senate he essentially controlled majority congressional support for the legislative program: put simply, without his backing, the Republican administration was ineffective. Johnson saw Sputnik in terms of national security – the satellite could well have been an orbital bomb, waiting to be instructed to fall on an American city. He ordered a Congressional investigation into the state of national security preparedness. As a result, the public became aware that there was a ‘‘missile gap’’; and, almost overnight, ‘space’ was transformed from a fantasy into something that the US should be leading, since otherwise national prestige would be damaged.1

After the launch on 3 November of a heavier Sputnik with a canine passenger, Eisenhower demanded an increase in the pace of Vanguard, which was in trouble, and also authorised the Army Ballistic Missile Agency to prepare a Jupiter-C in case Vanguard should fail. Medaris had the solid rockets for the upper stages retrieved from storage and let von Braun loose.

On 6 December 1957 Vanguard ignited, lifted a few centimetres off the pad, then collapsed back and exploded in a fireball. On 31 January 1958 the Army launched a satellite using essentially the same vehicle configuration as the Stewart Committee had rejected. On being asked for permission to inform Washington of the success, Medaris reputedly said: ‘‘Not yet, let them sweat a little.’’ The satellite, Explorer 1, was integrated into the solid rocket of the final stage and inserted into an orbit which ranged between 360 and 2,550 km. The Geiger-Mueller tube it carried was supplied by James van Allen, a physicist at the University of Iowa, and detected the presence of charged-particle radiation trapped within the Earth’s magnetic field, far above the atmosphere.

With the development of nuclear-armed intercontinental-range ballistic missiles threatening to make manned strategic bombers obsolete, the Air Force reacted to the prospect of its strike force becoming ‘silo rats’ by claiming that it needed to develop a manned space flight capability. Its Ballistic Missile Division, headed by General Bernard Schriever, devised Man In Space Soonest. This envisaged a progression of steps that would result in an Air Force officer landing on the Moon in 1965. When this was submitted to the Pentagon in March 1958 the response was lukewarm – in

part owing to the estimated cost of $1.5 billion, but also due to the absence of a clear military necessity. In fact, the proposal was an example of what would be referred to in today’s parlance as a demonstration of ‘the vision thing’.

No sooner had the Army developed its Jupiter intermediate-range ballistic missile than the Pentagon assigned operational control of all land-based missiles with ranges exceeding 320 km to the Air Force, thus limiting the Army to ‘battlefield’ missiles. In fact, the Air Force had no use for the Jupiter, since it had just developed its own Thor intermediate-range ballistic missile.

The only prospect for the Army Ballistic Missile Agency was therefore to develop powerful launch vehicles for satellites. On 19 December 1957 the Army proposed the National Integrated Missile and Space Vehicle Development Program. Like the Air Force, the Army saw itself as the obvious service to explore space. In 1959 it proposed Project Horizon to achieve a manned lunar landing in 1965, but this was received no more enthusiastically than the rival Man In Space Soonest.

There was eagerness for the final Block II to provide close-up pictures and radar reflectivity of the Moon’s surface, as well as (hopefully) seismometry. Ranger 5 had been scheduled for June 1962, but was postponed to allow the first pair of Mariner interplanetary missions to be dispatched in July and August.

On 30 August Rolph Hastrup, in charge of sterilisation, recommended that heat – treatment not be applied to the Block III. The use of ‘clean rooms’ to assemble the spacecraft, and the infusion of gaseous ethylene oxide to sterilise it within the Agena shroud shortly prior to launch should be continued. Clifford Cummings postponed a decision until after the next mission.

Ranger 5 arrived at the Cape on 27 August. As a result of recent modifications, it was about 10 kg heavier than its predecessors. The countdown on 16 October was scrubbed when a short circuit occurred in the spacecraft’s radio system. A launch the next day was ruled out by high winds. The mission got underway at 18:00 GMT on 18 October. Despite suffering a glitch, the Atlas responded to steering commands from the Cape, and the Agena achieved the desired parking orbit. This time, tracking ships were stationed in the Atlantic in order to provide continuous monitoring of the spacecraft’s telemetry. It had been decided that if the trajectory from the Agena’s second burn were to be beyond the spacecraft’s ability to correct, then the scientific priority would be to obtain gamma-ray data, rather than to snap flyby pictures of the Moon. This was because the Block III would provide TV, whereas there would be no gamma-ray spectrometers on any spacecraft that would head into deep space any time soon. The Agena made the translunar injection and released its payload. For the first time, the Deep Space Instrumentation Facility had two missions to keep track of in space. Mariner 2 was cruising to Venus, but the lunar mission would have priority call on resources during its 3-day flight.

When the Woomera tracking station acquired Ranger 5, it had deployed its solar panels and locked onto the Sun. The next task was to roll in order to acquire Earth as the second point of reference. But the temperature in the power switching and logic module of the computer/sequencer rose sharply and power from the solar panels was lost – there had been a short circuit. Patrick Rygh had replaced Marshall Johnson in charge of the Space Flight Operations Center, to free Johnson to manage the design and construction of the new Space Flight Operations Facility. James Burke, at the Cape, directed Rygh to have the spacecraft make a midcourse manoeuvre before its battery expired, to ensure that it would hit the Moon. But because the spacecraft had not acquired Earth its actual orientation in space was indeterminate. It was therefore decided to set up the manoeuvre using only the Sun as a reference. A command was uplinked to gimbal the high-gain antenna away from the nozzle of the engine on the base of the bus. The spacecraft initiated the ad hoc 30- minute manoeuvre sequence, but before it could be completed the transmitter fell silent. It appeared that electrical shorts had drained the battery. The Moon was ‘last quarter’ on 20 October. The inert vehicle flew by the trailing limb on 21 October at an altitude of 720 km and passed on into solar orbit – its progress once again being tracked by the transmitter in the surface package.

On 22 October W. H. Pickering ordered an investigation staffed by JPL personnel who were not involved in the project. When this issued its report on 13 November, it lamented that the mass limit imposed on the Block II prevented it from having any redundancy – in order to achieve its mission, the spacecraft required every system to

work. Burke was criticised for (in the opinion of people not involved) having spent too much of his time on launch vehicles, launch operations and space experiments, as opposed to the spacecraft. Burke was also criticised for the importance he gave to meeting schedules. However, in this he had merely been reflecting NASA’s desire to get ahead of the Soviets within the 36 months that had been assigned to the project. The structure of JPL was also criticised, in that engineers assigned to work on flight projects by the technical divisions often lacked vital experience, and section chiefs unfamiliar with either the project management or the subsystems that their engineers worked on had inadequately reviewed this work. Remarkably, despite the fact that a lack of commonality in the failures implied a reliability issue in the components, the investigation did not address the issue of heat sterilisation, and Hastrup’s memo to Cummings was not discussed. The report concluded that the Block III was unlikely to perform any better. To remedy the situation, it recommended (in part) that Burke be replaced and that his successor review the Block III design, add redundancy, and introduce new project management, inspection and testing procedures.

Neither of the two Block Is had achieved the intended high-apogee orbits (owing to Agena problems) and only one of the three Block Ils had reached the Moon (in an inert state). The project had been acknowledged to be technologically risky when it was commissioned, but no one had expected such poor performance. The spacecraft failures undoubtedly resulted from heat-sterilisation. The only scientific result from the entire exercise was provided by the gamma-ray spectrometer of Ranger 3, which established the existence of ‘hard’ radiation in space. However, absolutely nothing had been learned about the Moon. Nevertheless, the sense of ‘crisis’ would not have come about if the final Block II mission had been a complete success.

Responding to the mood, Homer Newell asked Oran Nicks to establish a Board of Inquiry to review the past performance and future prospects of the Ranger project. It was chaired by Albert J. Kelley, Director of the Electronics and Control Division of the Office of Advanced Research and Technology, and drew its membership from headquarters, field centres not involved in the project, and analysts from Bellcomm Incorporated – a systems engineering group established by the American Telephone & Telegraph Company in March 1962 at the request of the Office of Manned Space Flight to conduct independent analyses in support of Apollo. In particular, it was to submit recommendations ‘‘necessary to achieve successful Ranger operation’’. No thought was given to cancelling the project, because the high-resolution TV from the Block III was required for Apollo. On 30 November the Board issued its report. As regards JPL, it said that because the laboratory was attempting to use a common bus for its lunar and planetary projects, Ranger was more complex than strictly required, and as yet the high order of engineering skill and fabrication technology required for this not to represent an issue had yet to be achieved. It also said that the degree of ground testing was inadequate – the laboratory’s tradition with military missiles was to iron out problems by test flights; this was impractical with spacecraft. The Board judged heat sterilisation to have been a significant factor in the failure rate. Of course, the lack of redundancy in the spacecraft was criticised. JPL was also criticised for trying to run such a major venture simply by superimposing a small project office on top of its divisional structure. The recommendations therefore included strengthening the project office at JPL and revising the procedures for design review, design change control, testing and quality assurance. Heat sterilisation should cease. The Block III objectives should be restated, and all activities which did not directly contribute put aside. If additional versions of the spacecraft were required, then JPL should hire an industrial contractor.

On 7 December 1962 JPL relieved both Clifford Cummings and James Burke of their posts. On 12 December, Brian Sparks, Deputy Director of the laboratory, led a delegation to Washington to discuss the Kelley report with Homer Newell. At this and a second meeting on 17 December it was decided (in part) to delete the eight particles and fields experiments which Newell had added to the Block III in March; to discontinue heat sterilisation and the use of gaseous ethylene oxide; to discard all heat-treated hardware; and that (as originally intended) the sole goal of the Block III would be to obtain high-resolution TV of the lunar surface in support of Apollo. On 21 January 1963 William Cunningham, the Ranger Program Chief at headquarters, told the scientists that their experiments had been deleted from the Block III and, to ease the blow, pointed out that they would be favourably considered for carriage on possible future missions.

Morale at JPL was boosted on 14 December 1962 when Mariner 2 made a close flyby of Venus and became the first deep-space mission to make in-situ observations of another planet, along the way establishing that the solar wind was ‘gusty’.

On 18 December Robert Parks superseded Cummings, and Harris Schurmeier was made Ranger Project Manager – having been Chief of the Systems Division that had handled most of the work, he was the obvious choice. He immediately instituted a Ranger System Design Review Board involving Burke (who remained on the project staff), Gordon Kautz, Allen Wolfe and section chiefs of the supporting engineering divisions. Its primary task was to increase reliability by identifying and eliminating potential weak points in subsystems. The deletion of the experiments from the Block III released 22.5 kg of mass to accommodate redundancy. The enlarged solar panels were retained to provide a healthy power margin. Meanwhile, Bernard P. Miller of the Radio Corporation of America held a thorough review of the high-resolution TV package and recommended that the various wide-angle and narrow-angle cameras, together with their associated electronic assemblies, be split into two independent electrical chains so as to ensure that some pictures would be obtained even if an electrical problem were to disable one chain. Furthermore, to guard against the failure of the computer/sequencer, Miller recommended that a backup timer be added to start the TV system. Schurmeier accepted these recommendations. He also duplicated the gas supply of the attitude control system, and increased the capability of the main engine to make the Block III better able to correct a discrepancy in the translunar injection. And as arcing discharges were the single most worrisome cause of in-flight failures, he ordered that plastic covers be placed over all exposed terminals. W. H. Pickering strengthened the project office by revising the lines of authority and responsibility within the technical divisions so as to make the section chiefs personally involved in project activities, accountable for the quality of their engineers’ work, and no longer able to reassign personnel without the consent of the project manager. Pickering also made Ranger the laboratory’s highest priority flight project – thereby guaranteeing Schurmeier the authority he needed (and Burke had lacked) to drive work through in the manner desired. On 13 February 1963 NASA approved the long list of changes to be made to the Block III. In October the schedule for Block III was set, calling for missions in late January, March, May and July 1964.

Surveyor 4 was similar to Surveyor 3, with a soil mechanics surface sampler, but it also had a magnet on foot pad no. 2 to investigate whether there were magnetic particles in the surface material. It was to employ the last of the single-burn Centaur stages and fly essentially the same direct ascent trajectory as Surveyor 2 to aim for Sinus Medii. It lifted off from Pad 36A at 11:53:29 GMT on 14 July 1967. Both the Atlas and the Centaur performed satisfactorily, with translunar injection at 12:04:57. The spacecraft deployed its legs and omni-directional antenna booms, and, on being released, cancelled the inherited rates, acquired the Sun and deployed its solar panel. When commanded to acquire Canopus some 6 hours later, it did so without incident. It was decided to postpone the midcourse manoeuvre from the nominal 15 hours into the flight, and make it 24 hours later. The 10.5-second burn at 02:30:04 on 16 July imparted a change in velocity of 33.78 ft/sec to trim the initial divergence of 175 km from the centre of the 60-km-diameter target circle to a mere 8.5 km.

The pre-retro manoeuvre in which the spacecraft departed from its cruise attitude involved starting a roll of + 80.4 degrees at 01:24:44 on 17 July and a yaw of + 92.7 degrees at 01:29:34. This aligned the thrust axis with the velocity vector as that would be at retro ignition. The roll of -25.3 degrees at 01:35:05 was to optimise the illumination for post-landing imaging of crushable block no. 3. A landing on the prime meridian involved making an approach at 31.5 degrees to local vertical, as opposed to 23.6 degrees for Surveyor 3 at 23°W and 6.1 degrees for Surveyor 1 at 43°W. This would require a greater gravity turn in the vernier phase to force the trajectory to vertical. If successful, this mission would ‘open the door’ to sending future landers to targets in the eastern portion of the Apollo zone.

The altitude marking radar was enabled at 02:00:17, and issued its 100-km slant – range mark at 02:01:56.080. The programmed delay to the initiation of the braking manoeuvre was 2.725 seconds. The verniers ignited precisely on time, and the retro – rocket 1.1 seconds later – at which time the vehicle was travelling at 8,606 ft/sec. With everything apparently normal, the downlink fell silent at 02:02:41.018, when 40.9 seconds into the predicted 42.5-second duration of the retro-rocket’s burn. The vehicle was at an altitude of 49,420 feet, travelling at 1,092 ft/sec, and nominally 2 minutes

Outcome unknown 315

20 seconds from landing. The Deep Space Network was unable to re-establish contact with it.

The engineering team that studied the telemetry realised that whatever the fault was, it had cut the downlink within an interval of 0.25 millisecond without showing any indication in the preceding telemetry. The cause of the failure was not apparent. The only noteworthy unusual development was a slight modulation in the thrust of verniers no. 1 and 2, but it was not evident how this could have been relevant. The investigation listed four possible causes, without rating them in order of likelihood: (1) the breakage of a critical power lead in a wiring harness, or the failure of an

electrical connector, or the failure of a solder joint; (2) damage to the spacecraft’s circuitry from the rupture of the casing of the retro-rocket; (3) a transmitter failure; or (4) damage to the spacecraft’s circuitry caused by the rupture of a pressure vessel such as a shock absorber or a helium tank, nitrogen tank or vernier propellant tank. Since there was judged to be a “relatively low probability’’ of any of these failure modes recurring, no hardware changes were ordered.

Interestingly, if Surveyor 4’s problem was simply a transmitter failure, then it is highly likely that the vehicle landed safely.