In this article I have put together a collection of information in regard to the technology of the 1960s. Many people believe that it was technologically impossible to get a manned spacecraft to land on the Moon and return to Earth safely. Before Apollo 11 there were attempts to land probes on the Moon to gather images and soil samples – some were successful. These probes were very much smaller than the Apollo Lunar Landing modules.
Computer technology used onboard the LMs was very limiting compared with today’s standards. However, IBM had a fast growing technological industry dealing with mainframe computers. These very large machines could easily have been used to predict flight paths to, and landings on, the Moon’s Surface. So the argument about computer power must really be centred on the rope memory computers used in the LMs and not overall computing technology in the 1960s.
It’s Just Rocket Science
World War II
In 1943, production of the V-2 rocket began in Germany. It had an operational range of 300 km (190 mi) and carried a 1,000 kg (2,200 lb) warhead, with an amatol explosive charge. It normally achieved an operational maximum altitude of around 90 km (56 mi), but could achieve 206 km (128 mi) if launched vertically. The vehicle was similar to most modern rockets, with turbopumps, inertial guidance and many other features. Thousands were fired at various Allied nations, mainly Belgium, as well as England and France. While they could not be intercepted, their guidance system design and single conventional warhead meant that it was insufficiently accurate against military targets. A total of 2,754 people in England were killed, and 6,523 were wounded before the launch campaign was ended. There were also 20,000 deaths of slave labour during the construction of V-2s. While it did not significantly affect the course of the war, the V-2 provided a lethal demonstration of the potential for guided rockets as weapons.
In parallel with the guided missile programme in Nazi Germany, rockets were also used on aircraft, either for assisting horizontal take-off (RATO), vertical take-off (Bachem Ba 349″Natter”) or for powering them (Me 163, etc.). During the war Germany also developed several guided and unguided air-to-air, ground-to-air and ground-to-ground missiles (see list of World War II guided missiles of Germany).
The Allies rocket programs were much less sophisticated, relying mostly on unguided missiles like the Soviet Katyusha rocket.
Post World War II
At the end of World War II, competing Russian, British, and US military and scientific crews raced to capture technology and trained personnel from the German rocket program at Peenemünde. Russia and Britain had some success, but the United States benefited the most. The US captured a large number of German rocket scientists, including von Braun, and brought them to the United States as part of Operation Paperclip. In America, the same rockets that were designed to rain down on Britain were used instead by scientists as research vehicles for developing the new technology further. The V-2 evolved into the American Redstone rocket, used in the early space program.
After the war, rockets were used to study high-altitude conditions, by radio telemetry of temperature and pressure of the atmosphere, detection of cosmic rays, and further research; notably the Bell X-1, the first manned vehicle to break the sound barrier. This continued in the US under von Braun and the others, who were destined to become part of the US scientific community.
Independently, in the Soviet Union’s space program research continued under the leadership of the chief designer Sergei Korolev. With the help of German technicians, the V-2 was duplicated and improved as the R-1, R-2 and R-5 missiles. German designs were abandoned in the late 1940s, and the foreign workers were sent home. A new series of engines built by Glushko and based on inventions of Aleksei Mihailovich Isaev formed the basis of the first ICBM, the R-7. The R-7 launched the first satellite- Sputnik 1, and later Yuri Gagarin-the first man into space, and the first lunar and planetary probes. This rocket is still in use today. These prestigious events attracted the attention of top politicians, along with additional funds for further research.
One problem that had not been solved was atmospheric reentry. It had been shown that an orbital vehicle easily had enough kinetic energy to vaporize itself, and yet it was known that meteorites can make it down to the ground. The mystery was solved in the US in 1951 when H. Julian Allen and A. J. Eggers, Jr. of the National Advisory Committee for Aeronautics(NACA) made the counterintuitive discovery that a blunt shape (high drag) permitted the most effective heat shield. With this type of shape, around 99% of the energy goes into the air rather than the vehicle, and this permitted safe recovery of orbital vehicles.
The Allen and Eggers discovery, initially treated as a military secret, was eventually published in 1958. Blunt body theory made possible the heat shield designs that were embodied in the Mercury, Gemini, Apollo, and Soyuz space capsules, enabling astronauts and cosmonauts to survive the fiery re-entry into Earth’s atmosphere. Some spaceplanes such as the Space Shuttle made use of the same theory. At the time the STS was being conceived, Maxime Faget, the Director of Engineering and Development at the Manned Spacecraft Center, was not satisfied with the purely lifting re-entry method (as proposed for the cancelled X-20 “Dyna-Soar”). He designed a space shuttle which operated as a blunt body by entering the atmosphere at an extremely high angle of attack of 40° with the underside facing the direction of flight, creating a large shock wave that would deflect most of the heat around the vehicle instead of into it. The Space Shuttle essentially uses a combination of a ballistic entry (Blunt body theory) and then at an altitude of about 122,000 m (400,000 ft), the re-entry interface takes place. Here the atmosphere is dense enough for the Space Shuttle to begin its lifting re-entry by reducing the angle-of-attack, pointing the nose down and using the lift its wings generate to “start flying” (gliding) towards the landing site.
Rockets became extremely important militarily as modern intercontinental ballistic missiles (ICBMs) when it was realized that nuclear weapons carried on a rocket vehicle were essentially impossible for existing defense systems to stop once launched, and ICBM/Launch vehicles such as the R-7, Atlas and Titan became the delivery platform of choice for these weapons.
Fueled partly by the Cold War, the 1960s became the decade of rapid development of rocket technology particularly in the Soviet Union (Vostok, Soyuz, Proton) and in the United States (e.g. the X-15 and X-20 Dyna-Soar aircraft). There was also significant research in other countries, such as France, Britain, Japan, Australia, etc., and a growing use of rockets for Space exploration, with pictures returned from the far side of the Moon and unmanned flights for Mars exploration.
In America the manned programmes, Project Mercury, Project Gemini and later the Apollo programme culminated in 1969 with the first manned landing on the moon via the Saturn V, causing the New York Times to retract their earlier editorial implying that spaceflight couldn’t work:
Further investigation and experimentation have confirmed the findings of Isaac Newton in the 17th century and it is now definitely established that a rocket can function in a vacuum as well as in an atmosphere. The Times regrets the error.— New York Times, 17 June 1969 – A Correction
In the 1970s America made further lunar landings, before cancelling the Apollo program in 1975. The replacement vehicle, the partially reusable ‘Space Shuttle’ was intended to be cheaper, but this large reduction in costs was not achieved. Meanwhile, in 1973, the expendable Ariane programme was begun, a launcher that by the year 2000 would capture much of the geosat market[Wikipedia_1].
The Saturn V (spoken as “Saturn five”) was an American human-rated expendable rocket used by NASA between 1967 and 1973. The three-stage liquid-fueledsuper heavy-lift launch vehicle was developed to support the Apollo program for human exploration of the Moon and was later used to launch Skylab, the first American space station. The Saturn V was launched 13 times from the Kennedy Space Center in Florida with no loss of crew or payload. As of 2017, the Saturn V remains the tallest, heaviest, and most powerful (highest total impulse) rocket ever brought to operational status, and holds records for the heaviest payload launched and largest payload capacity to low Earth orbit (LEO) of 140,000 kg (310,000 lb), which included the third stage and unburned propellant needed to send the Apollo Command/Service Module and Lunar Module to the Moon.
The largest production model of the Saturn family of rockets, the Saturn V was designed under the direction of Wernher von Braun and Arthur Rudolph at the Marshall Space Flight Center in Huntsville, Alabama, with Boeing, North American Aviation, Douglas Aircraft Company, and IBM as the lead contractors.
To date, the Saturn V remains the only launch vehicle to launch missions to carry humans beyond low Earth orbit. A total of 15 flight-capable vehicles were built, but only 13 were flown. An additional three vehicles were built for ground testing purposes. A total of 24 astronauts were launched to the Moon, three of them twice, in the four years spanning December 1968 through December 1972.
The origins of the Saturn V rocket begin with the US government bringing Wernher von Braun along with about seven hundred German rocket engineers and technicians to the United States in Operation Paperclip, a program authorized by President Truman in August 1946 with the purpose of harvesting Germany’s rocket expertise, to give the US an edge in the Cold War through development of intermediate-range (IRBM) and intercontinental ballistic missiles (ICBM). It was known that America’s rival, the Soviet Union, would also try to secure some of the Germans.
Von Braun was put into the rocket design division of the Army due to his prior direct involvement in the creation of the V-2 rocket. Between 1945 and 1958, his work was restricted to conveying the ideas and methods behind the V-2 to the American engineers. Despite Von Braun’s many articles on the future of space rocketry, the US Government continued funding Air Force and Navy rocket programs to test their Vanguard missiles in spite of numerous costly failures. It was not until the 1957 Soviet launch of Sputnik 1 atop an R-7 ICBM capable of carrying a thermonuclear warhead to the US, that the Army and the government started taking serious steps towards putting Americans in space. Finally, they turned to von Braun and his team, who during these years created and experimented with the Jupiter series of rockets. The Juno I was the rocket that launched the first American satellite in January 1958, and part of the last-ditch plan for NACA (the predecessor of NASA) to get its foot in the Space Race. The Jupiter series was one more step in von Braun’s journey to the Saturn V, later calling that first series “an infant Saturn”.
The Saturn V’s design stemmed from the designs of the Jupiter series rockets. As the success of the Jupiter series became evident, the Saturn series emerged.
C-1 to C-4
Between 1960 and 1962, the Marshall Space Flight Center (MSFC) designed a series of Saturn rockets that could be used for various Earth orbit or lunar missions.
The C-1 was developed into the Saturn I, and the C-2 rocket was dropped early in the design process in favor of the C-3, which was intended to use two F-1engines on its first stage, four J-2 engines for its second stage, and an S-IV stage, using six RL10 engines.
NASA planned to use the C-3 as part of the Earth Orbit Rendezvous (EOR) concept, with at least four or five launches needed for a single lunar mission. But MSFC was already planning an even bigger rocket, the C-4, which would use four F-1 engines on its first stage, an enlarged C-3 second stage, and the S-IVB, a stage with a single J-2 engine, as its third stage. The C-4 would need only two launches to carry out an EOR lunar mission.
On January 10, 1962, NASA announced plans to build the C-5. The three-stage rocket would consist of: the S-IC first stage, with five F-1 engines; the S-II second stage, with five J-2 engines; and the S-IVB third stage, with a single J-2 engine. The C-5 was designed for a 90,000-pound (41,000 kg) payload capacity to the Moon.
The C-5 would undergo component testing even before the first model was constructed. The S-IVB third stage would be used as the second stage for the C-IB, which would serve both to demonstrate proof of concept and feasibility for the C-5, but would also provide flight data critical to development of the C-5. Rather than undergoing testing for each major component, the C-5 would be tested in an “all-up” fashion, meaning that the first test flight of the rocket would include complete versions of all three stages. By testing all components at once, far fewer test flights would be required before a manned launch.
The C-5 was confirmed as NASA’s choice for the Apollo program in early 1963, and was named the Saturn V. The C-1 became the Saturn I, and C-1B became Saturn IB. Von Braun headed a team at the Marshall Space Flight Center in building a vehicle capable of launching a manned spacecraft on a trajectory to the Moon. Before they moved under NASA’s jurisdiction, von Braun’s team had already begun work on improving the thrust, creating a less complex operating system, and designing better mechanical systems. It was during these revisions that the decision to reject the single engine of the V-2’s design came about, and the team moved to a multiple-engine design. The Saturn I and IB reflected these changes, but were not large enough to send a manned spacecraft to the Moon. These designs, however, provided a basis for which NASA could determine its best method towards landing a man on the Moon.
The Saturn V’s final design had several key design features. Engineers determined that the best engines were the F-1s coupled with the new liquid hydrogen propulsion system called J-2, which made the Saturn C-5 configuration optimal. By 1962, NASA had finalized its plans to proceed with von Braun’s Saturn designs, and the Apollo space program gained speed.
With the configuration finalized, NASA turned its attention to mission profiles. Despite some controversy, a lunar orbit rendezvous for the lunar module was chosen over an Earth orbital rendezvous. Issues such as type of fuel injections, the needed amount of fuel for such a trip, and rocket manufacturing processes were ironed out, and the designs for the Saturn V were selected. The stages were designed by von Braun’s Marshall Space Flight Center in Huntsville, and outside contractors were chosen for the construction: Boeing (S-IC), North American Aviation (S-II), Douglas Aircraft (S-IVB), and IBM (Instrument Unit).
Selection for Apollo lunar landing
Early in the planning process, NASA considered three leading ideas for the Moon mission: Earth Orbit Rendezvous, Direct Ascent, and Lunar Orbit Rendezvous (LOR). A direct ascent configuration would launch a larger rocket which would land directly on the lunar surface, while an Earth orbit rendezvous would launch two smaller spacecraft which would combine in Earth orbit. A LOR mission would involve a single rocket launching a single spacecraft, but only a small part of that spacecraft would land on the moon. That smaller landing module would then rendezvous with the main spacecraft, and the crew would return home.
NASA at first dismissed LOR as a riskier option, given that an orbital rendezvous had yet to be performed in Earth orbit, much less in lunar orbit. Several NASA officials, including Langley Research Center engineer John Houbolt and NASA Administrator George Low, argued that a Lunar Orbit Rendezvous provided the simplest landing on the moon, the most cost–efficient launch vehicle and, perhaps most importantly, the best chance to accomplish a lunar landing within the decade. Other NASA officials were convinced, and LOR was officially selected as the mission configuration for the Apollo program on November 7, 1962.
The Saturn V’s size and payload capacity dwarfed all other previous rockets which had successfully flown at that time. With the Apollo spacecraft on top, it stood 363 feet (111 m) tall, and without fins, it was 33 feet (10 m) in diameter. Fully fueled, the Saturn V weighed 6.5 million pounds (2,950 metric tons) and had a low Earth orbit payload capacity originally estimated at 261,000 pounds (118,000 kg), but was designed to send at least 90,000 pounds (41,000 kg) to the Moon. Later upgrades increased that capacity; during the final three Apollo lunar missions it deployed about 310,000 pounds (140,000 kg)[note 1] to LEO and sent up to 107,100 lb (48,600 kg) spacecraft to the Moon. At a height of 363 feet (111 m), the Saturn V was 58 feet (18 m) taller than the Statue of Liberty from the ground to the torch, and 48 feet (15 m) taller than the Big Benclock tower.
In contrast, the Mercury-Redstone Launch Vehicle used on Freedom 7, the first manned American spaceflight, was just under 11 feet (3.4 m) longer than the S-IVB stage, and delivered less sea level thrust (78,000 pounds-force (350 kN)) than the Launch Escape System rocket (150,000 pounds-force (667 kN) sea level thrust) mounted atop the Apollo Command Module.
The Saturn V was principally designed by the Marshall Space Flight Center in Huntsville, Alabama, although numerous major systems, including propulsion, were designed by subcontractors. It used the powerful new F-1 and J-2 rocket engines for propulsion. When tested, these engines shattered the windows of nearby houses. Designers decided early on to attempt to use as much technology from the Saturn I program as possible. Consequently, the S-IVB-500 third stage of the Saturn V was based on the S-IVB-200 second stage of the Saturn IB. The Instrument Unit that controlled the Saturn V shared characteristics with that carried by the Saturn IB.
Blueprints and other Saturn V plans are available on microfilm at the Marshall Space Flight Center.
The Saturn V consisted of three stages—the S-IC first stage, S-II second stage and the S-IVB third stage—and the instrument unit. All three stages used liquid oxygen (LOX) as an oxidizer. The first stage used RP-1 for fuel, while the second and third stages used liquid hydrogen (LH2). The upper stages also used small solid-fueled ullage motors that helped to separate the stages during the launch, and to ensure that the liquid propellants were in a proper position to be drawn into the pumps.
S-IC first stage
The S-IC was built by the Boeing Company at the Michoud Assembly Facility, New Orleans, where the Space Shuttle External Tanks would later be built by Lockheed Martin. Most of its mass at launch was propellant, RP-1 fuel with liquid oxygen as the oxidizer. It was 138 feet (42 m) tall and 33 feet (10 m) in diameter, and provided over 7,600,000 pounds-force (34,000 kN) of thrust. The S-IC stage had a dry weight of about 289,000 pounds (131 metric tons) and fully fueled at launch had a total weight of 5,100,000 pounds (2,300 metric tons). It was powered by five Rocketdyne F-1 engines arrayed in a quincunx (five units, with four arranged in a square, and the fifth in the center) The center engine was held in a fixed position, while the four outer engines could be hydraulically turned (gimballed) to steer the rocket. In flight, the center engine was turned off about 26 seconds earlier than the outboard engines to limit acceleration. During launch, the S-IC fired its engines for 168 seconds (ignition occurred about 8.9 seconds before liftoff) and at engine cutoff, the vehicle was at an altitude of about 36 nautical miles (67 km), was downrange about 50 nautical miles (93 km), and was moving about 7,500 feet per second (2,300 m/s).
S-II second stage
The S-II was built by North American Aviation at Seal Beach, California. Using liquid hydrogen and liquid oxygen, it had five Rocketdyne J-2engines in a similar arrangement to the S-IC, also using the outer engines for control. The S-II was 81 feet 7 inches (24.87 m) tall with a diameter of 33 feet (10 m), identical to the S-IC, and thus was the largest cryogenic stage until the launch of the Space Shuttle in 1981. The S-II had a dry weight of about 80,000 pounds (36,000 kg) and fully fueled, weighed 1,060,000 pounds (480,000 kg). The second stage accelerated the Saturn V through the upper atmosphere with 1,100,000 pounds-force (4,900 kN) of thrust in vacuum. When loaded, significantly more than 90 percent of the mass of the stage was propellant; however, the ultra-lightweight design had led to two failures in structural testing. Instead of having an intertank structure to separate the two fuel tanks as was done in the S-IC, the S-II used a common bulkhead that was constructed from both the top of the LOX tank and bottom of the LH2 tank. It consisted of two aluminum sheets separated by a honeycomb structure made of phenolic resin. This bulkhead had to insulate against the 126 °F (70 °C) temperature gradient between the two tanks. The use of a common bulkhead saved 7,900 pounds (3.6 t). Like the S-IC, the S-II was transported from its manufacturing plant to the Cape by sea.
S-IVB third stage
The S-IVB was built by the Douglas Aircraft Company at Huntington Beach, California. It had one J-2 engine and used the same fuel as the S-II. The S-IVB used a common bulkhead to separate the two tanks. It was 58 feet 7 inches (17.86 m) tall with a diameter of 21 feet 8 inches (6.604 m) and was also designed with high mass efficiency, though not quite as aggressively as the S-II. The S-IVB had a dry weight of about 23,000 pounds (10,000 kg) and, fully fueled, weighed about 262,000 pounds (119,000 kg).
The S-IVB-500 model used on the Saturn V differed from the S-IVB-200 used as the second stage of the Saturn IB, in that the engine was restartable once per mission. This was necessary as the stage would be used twice during a lunar mission: first in a 2.5 min burn for the orbit insertion after second stage cutoff, and later for the trans-lunar injection (TLI) burn, lasting about 6 min. Two liquid-fueled Auxiliary Propulsion System (APS) units mounted at the aft end of the stage were used for attitude control during the parking orbit and the trans-lunar phases of the mission. The two APSs were also used as ullage engines to settle the propellants in the aft tank engine feed lines prior to the trans-lunar injection burn.
The S-IVB was the only rocket stage of the Saturn V small enough to be transported by plane, in this case the Pregnant Guppy.
The Instrument Unit was built by IBM and rode atop the third stage. It was constructed at the Space Systems Center in Huntsville, Alabama. This computer controlled the operations of the rocket from just before liftoff until the S-IVB was discarded. It included guidance and telemetry systems for the rocket. By measuring the acceleration and vehicle attitude, it could calculate the position and velocity of the rocket and correct for any deviations.
In the event of an abort requiring the destruction of the rocket, the range safety officer would remotely shut down the engines and after several seconds send another command for the shaped explosive charges attached to the outer surfaces of the rocket to detonate. These would make cuts in fuel and oxidizer tanks to disperse the fuel quickly and to minimize mixing. The pause between these actions would give time for the crew to escape using the Launch Escape Tower or (in the later stages of the flight) the propulsion system of the Service module. A third command, “safe”, was used after the S-IVB stage reached orbit to irreversibly deactivate the self-destruct system. The system was also inactive as long as the rocket was still on the launch pad.
The Soviet space program’s counterpart to the Saturn V was Sergei Korolev’s N1-L3. The Saturn V was taller, heavier, and had greater payload capacity, both to low Earth orbitand to translunar injection. The N-1 was a three-stage launch vehicle with more liftoff thrust and a larger first stage diameter than the Saturn V. It was to carry the 209,000 lb (95,000 kg) L3 vehicle into orbit. The L3 contained an Earth departure stage, which would send to the Moon a 51,800 lb (23,500 kg) package which contained another stage for lunar orbit insertion and powered descent initiation, a single-cosmonaut lander, and a two-cosmonaut lunar orbiter for the return to Earth. The N1/L3 would have produced more total impulse (product of thrust and time) in its first four stages than the three-stage Saturn V, but it was not able to convert as much of this into payload momentum (product of mass and velocity).
The N1 never became operational; four test launches each resulted in catastrophic vehicle failure early in flight, and the program was canceled. Korolev elected to cluster 30 relatively small engines for the first stage, rather than develop a large engine like the Rocketdyne F-1.
The three-stage Saturn V grew over its lifetime to a peak thrust of at least 7,650,000 lbf (34,020 kN) (AS-510 and subsequent) and a lift capacity of 310,000 lb (140,000 kg) to LEO. The AS-510 mission (Apollo 15) had a liftoff thrust of 7,823,000 lbf (34,800 kN). The AS-513 mission (Skylab 1) had slightly greater liftoff thrust of 7,891,000 lbf (35,100 kN). By comparison, the N-1 had a sea-level liftoff thrust of about 10,200,000 lbf (45,400 kN). No other operational launch vehicle has ever surpassed the Saturn V in height, weight, total impulse, or payload capability. The closest contenders were the US Space Shuttle and the Soviet Energia.
U.S. Space Shuttle
The Space Shuttle generated a peak thrust of 6,800,000 lbf (30,100 kN), and payload capacity to LEO (excluding the Orbiter itself) was 63,500 pounds (28,800 kg), which was about 25 percent of the Saturn V’s payload. Total mass in orbit, including the Orbiter, was about 247,000 lb (112,000 kg), compared to the Apollo 15 total orbital mass of the S-IVB third stage and Apollo spacecraft, of 309,771 lb (140,510 kg), some 62,800 lb (28,500 kg) heavier than the Shuttle was rated to carry to LEO.
Energia had a liftoff thrust of 7,826,000 lbf (34,810 kN), and flew twice in 1987 and 1988, the second time as the launcher for the Buran shuttle. However, both the Energia and Buran programs were cancelled in 1993. Hypothetical future versions of Energia might have been significantly more powerful than the Saturn V, delivering 10,000,000 lbf (46,000 kN) of thrust and able to deliver up to 386,000 lb (175 t) to LEO in the “Vulkan” configuration. Planned uprated versions of the Saturn V using F-1A engines would have had about 18 percent more thrust and 302,580 pounds (137,250 kg) payload. NASA contemplated building larger members of the Saturn family, such as the Saturn C-8, and also unrelated rockets, such as Nova, but these were never produced.
Some other recent US launch vehicles have significantly lower launch capacity to LEO than Saturn V: the US Delta 4 Heavy capacity is 63,470 lb (28,790 kg), the Atlas V 551 has a capacity of 41,478 lb (18,814 kg), and the planned SpaceX Falcon Heavy has a 141,000 lb (64,000 kg) projected capacity. The European Ariane 5 ES delivers up to 46,000 lb (21,000 kg) and the Russian Proton-M can launch 49,000 lb (22,000 kg).
Space Launch System
NASA’s Space Launch System, planned for its first flight in 2018, in its final configuration is planned to be 400 feet (120 m) tall with payload, and lift up to 290,000 pounds (130,000 kg) into low Earth orbit.
S-IC thrust comparisons
Because of its large size, attention is often focused on the S-IC thrust and how this compares to other large rockets. However, several factors make such comparisons more complex than first appears:
Without knowing the exact measurement technique and mathematical method used to determine thrust for each different rocket, comparisons are often inexact. As the above shows, the specified thrust often differs significantly from actual flight thrust calculated from direct measurements. The thrust stated in various references is often not adequately qualified as to vacuum vs sea level, or peak vs average thrust.
Similarly, payload increases are often achieved in later missions independent of engine thrust. This is by weight reduction or trajectory reshaping.
The result is there is no single absolute figure for engine thrust, stage thrust or vehicle payload. There are specified values and actual flight values, and various ways of measuring and deriving those actual flight values.
The performance of each Saturn V launch was extensively analyzed and a Launch Evaluation Report produced for each mission, including a thrust/time graph for each vehicle stage on each mission.
After the construction and ground testing of a stage was completed, it was then shipped to the Kennedy Space Center. The first two stages were so large that the only way to transport them was by barge. The S-IC, constructed in New Orleans, was transported down the Mississippi River to the Gulf of Mexico. After rounding Florida, it was then transported up the Intra-Coastal Waterway to the Vertical Assembly Building (now called the Vehicle Assembly Building). This was essentially the same route which would be used later by NASA to ship Space Shuttle External Tanks. The S-II was constructed in California and thus traveled to Florida via the Panama Canal. The third stage and Instrument Unit could be carried by the Aero Spacelines Pregnant Guppy and Super Guppy, but could also have been carried by barge if warranted.
On arrival at the Vertical Assembly Building, each stage was inspected in a horizontal position before being moved to a vertical position. NASA also constructed large spool-shaped structures that could be used in place of stages if a particular stage was late. These spools had the same height and mass and contained the same electrical connections as the actual stages.
NASA stacked or assembled the Saturn V on a Mobile Launcher Platform (MLP), which consisted of a Launch Umbilical Tower (LUT) with nine swing arms (including the crew access arm), a “hammerhead” crane, and a water suppression system which was activated prior to launch. After assembly was completed, the entire stack was moved from the VAB to the launch pad using the Crawler Transporter (CT). Built by the Marion Power Shovel company (and later used for transporting the smaller and lighter Space Shuttle), the CT ran on four double-tracked treads, each with 57 ‘shoes’. Each shoe weighed 2,000 pounds (910 kg). This transporter was also required to keep the rocket level as it traveled the 3 miles (4.8 km) to the launch site, especially at the 3 percent grade encountered at the launch pad. The CT also carried the Mobile Service Structure (MSS), which allowed technicians access to the rocket until eight hours before launch, when it was moved to the “halfway” point on the Crawlerway (the junction between the VAB and the two launch pads).
Lunar mission launch sequence
The Saturn V carried all Apollo lunar missions. All Saturn V missions were launched from Launch Complex 39 at the John F. Kennedy Space Center in Florida. After the rocket cleared the launch tower, flight control transferred to Johnson Space Center’s Mission Control in Houston, Texas.
An average mission used the rocket for a total of just 20 minutes. Although Apollo 6 experienced three engine failures, and Apollo 13 one engine shutdown, the onboard computers were able to compensate by burning the remaining engines longer to achieve parking orbit. None of the Saturn V launches resulted in a payload loss.
The first stage burned for about 2 minutes and 41 seconds, lifting the rocket to an altitude of 42 miles (68 km) and a speed of 6,164 miles per hour (2,756 m/s) and burning 4,700,000 pounds (2,100,000 kg) of propellant.
At 8.9 seconds before launch, the first stage ignition sequence started. The center engine ignited first, followed by opposing outboard pairs at 300-millisecond intervals to reduce the structural loads on the rocket. When thrust had been confirmed by the onboard computers, the rocket was “soft-released” in two stages: first, the hold-down arms released the rocket, and second, as the rocket began to accelerate upwards, it was slowed by tapered metal pins pulled through dies for half a second. Once the rocket had lifted off, it could not safely settle back down onto the pad if the engines failed. The astronauts considered this one of the tensest moments in riding the Saturn V, for if the rocket did fail to lift off after release they had a low chance of survival given the large amounts of propellant. A fully fueled Saturn V exploding on the pad would have released the energy equivalent of two kilotons of TNT. To improve safety, the Saturn Emergency Detection System (EDS) inhibited engine shutdown for the first 30 seconds of flight. (See Saturn V Instrument Unit)
It took about 12 seconds for the rocket to clear the tower. During this time, it yawed 1.25 degrees away from the tower to ensure adequate clearance despite adverse winds. (This yaw, although small, can be seen in launch photos taken from the east or west.) At an altitude of 430 feet (130 m) the rocket rolled to the correct flight azimuth and then gradually pitched down until 38 seconds after second stage ignition. This pitch program was set according to the prevailing winds during the launch month. The four outboard engines also tilted toward the outside so that in the event of a premature outboard engine shutdown the remaining engines would thrust through the rocket’s center of gravity. The Saturn V reached 400 feet per second (120 m/s) at over 1 mile (1,600 m) in altitude. Much of the early portion of the flight was spent gaining altitude, with the required velocity coming later. The Saturn V broke the sound barrier at just over 1 minute at an altitude of between 3 and 4 nautical miles. At this point, shock collars, or condensation clouds, could be seen forming around the bottom of the command module and around the top of the second stage.
At about 80 seconds, the rocket experienced maximum dynamic pressure (max Q). The dynamic pressure on a rocket varies with air density and the square of relative velocity. Although velocity continues to increase, air density decreases so quickly with altitude that dynamic pressure falls below max Q.
Acceleration increased during S-IC flight for three reasons. One, increased acceleration increased the propellant pressure at the engines, increasing the flow rate somewhat. This was the least important factor, though this feedback effect often led to an undesirable thrust oscillation called pogo. Two, as it climbed into thinner air F-1 engine efficiency increased significantly, a property of all rockets. The combined thrust of five engines on the pad was about 7.5 million pounds, reaching nearly 9 million pounds at altitude. But the biggest contribution by far was the rocket’s rapidly decreasing mass. The propellant in just the S-IC made up about three-quarters of Saturn V’s entire launch mass, and it was furiously consumed at over 13 metric tonnes per second. Newton’s second law states that force is equal to mass times acceleration, or equivalently that acceleration is equal to force divided by mass, so as the mass decreased (and the force increased somewhat), acceleration rose. Including gravity, launch acceleration was only 1¼ g, i.e., the astronauts felt 1¼ g while the rocket accelerated vertically at ¼ g. As the rocket rapidly lost mass, total acceleration including gravity increased to nearly 4 g at T+135 seconds. At this point, the inboard (center) engine was shut down to prevent acceleration from increasing beyond 4 g.
When oxidizer or fuel depletion was sensed in the suction assemblies, the remaining four outboard engines were shut down. First stage separation occurred a little less than one second after this to allow for F-1 thrust tail-off. Eight small solid fuel separation motors backed the S-IC from the rest of the vehicle at an altitude of about 36 nautical miles (67 km). The first stage continued ballistically to an altitude of about 59 nautical miles (109 km) and then fell in the Atlantic Ocean about 300 nautical miles (560 km) downrange.
The engine shutdown procedure was changed for the launch of Skylab to avoid damage to the Apollo Telescope Mount. Rather than shutting down all four outboard engines at once, they were shut down two at a time with a delay to reduce peak acceleration further.
After S-IC separation, the S-II second stage burned for 6 minutes and propelled the craft to 109 miles (175 km) and 15,647 mph (6,995 m/s), close to orbital velocity.
For the first two unmanned launches, eight solid-fuel ullage motors ignited for four seconds to give positive acceleration to the S-II stage, followed by start of the five J-2 engines. For the first seven manned Apollo missions only four ullage motors were used on the S-II, and they were eliminated completely for the final four launches. About 30 seconds after first stage separation, the interstage ring dropped from the second stage. This was done with an inertially fixed attitude so that the interstage, only 1 meter from the outboard J-2 engines, would fall cleanly without contacting them. Shortly after interstage separation the Launch Escape System was also jettisoned. See Apollo abort modesfor more information about the various abort modes that could have been used during a launch.
About 38 seconds after the second stage ignition the Saturn V switched from a preprogrammed trajectory to a “closed loop” or Iterative Guidance Mode. The Instrument Unit now computed in real time the most fuel-efficient trajectory toward its target orbit. If the Instrument Unit failed, the crew could switch control of the Saturn to the Command Module’s computer, take manual control, or abort the flight.
About 90 seconds before the second stage cutoff, the center engine shut down to reduce longitudinal pogo oscillations. At around this time, the LOX flow rate decreased, changing the mix ratio of the two propellants, ensuring that there would be as little propellant as possible left in the tanks at the end of second stage flight. This was done at a predetermined delta-v.
Five level sensors in the bottom of each S-II propellant tank were armed during S-II flight, allowing any two to trigger S-II cutoff and staging when they were uncovered. One second after the second stage cut off it separated and several seconds later the third stage ignited. Solid fuel retro-rockets mounted on the interstage at the top of the S-II fired to back it away from the S-IVB. The S-II impacted about 2,300 nautical miles (4,200 km) from the launch site.
On the Apollo 13 mission, the inboard engine suffered from major pogo oscillation, resulting in an early automatic cutoff. To ensure sufficient velocity was reached, the remaining four engines were kept active for longer than planned. A pogo suppressor was fitted to later Apollo missions to avoid this, though the early engine 5 cutoff remained to reduce g-forces.
Unlike the two-plane separation of the S-IC and S-II, the S-II and S-IVB stages separated with a single step. Although it was constructed as part of the third stage, the interstage remained attached to the second stage.
During Apollo 11, a typical lunar mission, the third stage burned for about 2.5 minutes until first cutoff at 11 minutes 40 seconds. At this point it was 1,430 nautical miles (2,650 km) downrange and in a parking orbit at an altitude of 103.2 nautical miles (191.1 km) and velocity of 17,432 mph (7,793 m/s). The third stage remained attached to the spacecraft while it orbited the Earth one and a half times while astronauts and mission controllers prepared for translunar injection (TLI).
This parking orbit was quite low by Earth orbit standards, and it would have been short-lived due to aerodynamic drag. This was not a problem on a lunar mission because of the short stay in the parking orbit. The S-IVB also continued to thrust at a low level by venting gaseous hydrogen, to keep propellants settled in their tanks and prevent gaseous cavities from forming in propellant feed lines. This venting also maintained safe pressures as liquid hydrogen boiled off in the fuel tank. This venting thrust easily exceeded aerodynamic drag.
For the final three Apollo flights, the temporary parking orbit was even lower (approximately 93 nautical miles (172 km)), to increase payload for these missions. The Apollo 9Earth orbit mission was launched into the nominal orbit consistent with Apollo 11, but the spacecraft were able to use their own engines to raise the perigee high enough to sustain the 10-day mission. The Skylab was launched into a quite different orbit, with a 234-nautical-mile (434 km) perigee which sustained it for six years, and also a higher inclination to the equator (50 degrees versus 32.5 degrees for Apollo).
On Apollo 11, TLI came at 2 hours and 44 minutes after launch. The S-IVB burned for almost six minutes giving the spacecraft a velocity close to the Earth’s escape velocity of 25,053 mph (11,200 m/s). This gave an energy-efficient transfer to lunar orbit, with the Moon helping to capture the spacecraft with a minimum of CSM fuel consumption.
About 40 minutes after TLI the Apollo Command Service Module (CSM) separated from the third stage, turned 180 degrees and docked with the Lunar Module (LM) that rode below the CSM during launch. The CSM and LM separated from the spent third stage 50 minutes later. This process is known as Transposition, docking, and extraction.
If it were to remain on the same trajectory as the spacecraft, the S-IVB could have presented a collision hazard so its remaining propellants were vented and the auxiliary propulsion system fired to move it away. For lunar missions before Apollo 13, the S-IVB was directed toward the Moon’s trailing edge in its orbit so that the moon would slingshot it beyond earth escape velocity and into solar orbit. From Apollo 13 onwards, controllers directed the S-IVB to hit the Moon. Seismometers left behind by previous missions detected the impacts, and the information helped map the interior structure of the Moon.
On September 3, 2002, astronomer Bill Yeung discovered a suspected asteroid, which was given the discovery designation J002E3. It appeared to be in orbit around the Earth, and was soon discovered from spectral analysis to be covered in white titanium dioxide, which was a major constituent of the paint used on the Saturn V. Calculation of orbital parameters led to tentative identification as being the Apollo 12 S-IVB stage. Mission controllers had planned to send Apollo 12’s S-IVB into solar orbit after separating from the Apollo spacecraft, but it is believed the burn lasted too long, and hence did not send it close enough to the Moon, remaining in a barely stable orbit around the Earth and Moon. In 1971, through a series of gravitational perturbations, it is believed to have entered in a solar orbit and then returned into weakly captured Earth orbit 31 years later. It left Earth orbit again in June 2003.
In 1965, the Apollo Applications Program (AAP) was created to look into science missions that could be performed using Apollo hardware. Much of the planning centered on the idea of a space station. Wernher von Braun’s earlier (1964) plans employed a “wet workshop” concept, with a spent S-II Saturn V second stage being launched into orbit and outfitted in space. The next year AAP studied a smaller station using an S-IVB Saturn 1B second stage. By 1969, Apollo funding cuts eliminated the possibility of procuring more Apollo hardware, and in fact forced the cancellation of some later Moon landing flights. This freed up at least one Saturn V, allowing the wet workshop to be replaced with the “dry workshop” concept: the station (now known as Skylab) would be built on the ground from a surplus Saturn IB second stage and launched on the first two live stages of a Saturn V. A backup station, constructed from a Saturn V third stage, was built and is now on display at the National Air and Space Museum.
Skylab was the only launch not directly related to the Apollo lunar landing program. The only significant changes to the Saturn V from the Apollo configurations involved some modification to the S-II to act as the terminal stage for inserting the Skylab payload into Earth orbit,[specify] and to vent excess propellant after engine cutoff so the spent stage would not rupture in orbit. The S-II remained in orbit for almost two years, and made an uncontrolled re-entry on January 11, 1975.
Three crews lived aboard Skylab from May 25, 1973 to February 8, 1974, with Skylab remaining in orbit until July 11, 1979.
Proposed post-Apollo developments
After Apollo, the Saturn V was planned to be the prime launch vehicle for Prospector intended to land a 330-kilogram (730 lb) robotic rover on the Moon, similar to the Soviet Lunokhod, and the Voyager Mars probes, as well a scaled-up version of the Voyager interplanetary probes. It was also to have been the launch vehicle for the nuclear rocket stage RIFT test program and the later NERVA. All of these planned uses of the Saturn V were cancelled, with cost being a major factor. Edgar Cortright, who had been director of NASA Langley, stated decades later that “JPL never liked the big approach. They always argued against it. I probably was the leading proponent in using the Saturn V, and I lost. Probably very wise that I lost.”
The canceled second production run of Saturn Vs would very likely have used the F-1A engine in its first stage, providing a substantial performance boost. Other likely changes would have been the removal of the fins (which turned out to provide little benefit when compared to their weight); a stretched S-IC first stage to support the more powerful F-1As; and uprated J-2s or an M-1 for the upper stages.
A number of alternate Saturn vehicles were proposed based on the Saturn V, ranging from the Saturn INT-20 with an S-IVB stage and interstage mounted directly onto an S-ICstage, through to the Saturn V-23(L) which would not only have five F-1 engines in the first stage, but also four strap-on boosters with two F-1 engines each: giving a total of thirteen F-1 engines firing at launch.
The Space Shuttle was initially conceived as a cargo transport to be used in concert with the Saturn V, even to the point that a Saturn-Shuttle was proposed, using the winged shuttle orbiter and external tank, but with the tank mounted on a modified, fly-back version of the S-IC. The first S-IC stage would be used to power the Shuttle during the first two minutes of flight, after which the S-IC would be jettisoned (which would then fly back to KSC for refurbishment) and the Space Shuttle Main Engines would then fire and place the orbiter into orbit. The Shuttle would handle space station logistics, while Saturn V would launch components. Lack of a second Saturn V production run killed this plan and has left the United States without a heavy-lift launch vehicle. Some in the U.S. space community have come to lament this situation, as continued production would have allowed the International Space Station, using a Skylab or Mir configuration with both U.S. and Russian docking ports, to have been lifted with just a handful of launches. The Saturn-Shuttle concept also would have eliminated the Space Shuttle Solid Rocket Boosters that ultimately precipitated the Challenger accident in 1986.
From 1964 until 1973, $6.417 billion (equivalent to $73.2 billion in 2016) in total was appropriated for the R&D and flights of the Saturn V, with the maximum being in 1966 with $1.2 billion (equivalent to $16.6 billion in 2016). That same year, NASA received its biggest budget of US$4.5 billion, about 0.5 percent of the gross domestic product (GDP) of the United States at that time.
One of the main reasons for the cancellation of the last three Apollo flights was the cost. In the time frame from 1969 to 1971 the cost of launching a Saturn V Apollo mission was US $185 to $189 million, of which $110 million was for the production of the vehicle (equivalent to $1.26 billion in 2016).
Saturn V vehicles and launches
U.S. proposals for a rocket larger than the Saturn V from the late 1950s through the early 1980s were generally called Nova. Over thirty different large rocket proposals carried the Nova name, but none was developed.
Wernher von Braun and others also had plans for a rocket that would have featured eight F-1 engines in its first stage, like the Saturn C-8, allowing a direct ascent flight to the Moon. Other plans for the Saturn V called for using a Centaur as an upper stage or adding strap-on boosters. These enhancements would have enabled the launch of large robotic spacecraft to the outer planets or send astronauts to Mars. Other Saturn-V derivatives analyzed included the Saturn MLV family of “Modified Launch Vehicles”, which would have almost doubled the payload lift capability of the standard Saturn V and were intended for use in a proposed mission to Mars by 1980.
In 1968, Boeing studied another Saturn-V derivative, the Saturn C-5N, which included a nuclear thermal rocket engine for the third stage of the vehicle. The Saturn C-5N would carry a considerably greater payload to interplanetary destinations. Work on the nuclear engines, along with all Saturn V ELVs, was ended in 1973.
In 2006, as part of the proposed Constellation Program, NASA unveiled plans to construct two Shuttle Derived Launch Vehicles, the Ares I and Ares V, which would use some existing Space Shuttle and Saturn V hardware and infrastructure. The two rockets were intended to increase safety by specializing each vehicle for different tasks, Ares I for crew launches and Ares V for cargo launches. The original design of the heavy-lift Ares V, named in homage to the Saturn V, was 360 ft (110 m) in height and featured a core stage based on the Space Shuttle External Tank, with a diameter of 28 ft (8.4 m). It was to be powered by five Space Shuttle Main Engines (SSMEs) and two five-segment Space Shuttle Solid Rocket Boosters (SRBs). As the design evolved, the SSMEs were replaced with five RS-68 engines, the same engines used on the Delta IV. The switch from the SSME to the RS-68 was intended to reduce cost, the RS-68 being cheaper, simpler to manufacture, and more powerful than the SSME, though the lower efficiency of the RS-68 required an increase in core stage diameter to 33 ft (10 m), the same diameter as the Saturn V’s S-IC and S-II stages.
In 2008, NASA again redesigned the Ares V, lengthening the core stage, adding a sixth RS-68 engine, and increasing the SRBs to 5.5 segments each. This vehicle would have been 381 ft (116 m) tall and would have produced a total thrust of approximately 8,900,000 lbf (40 MN) at liftoff, more than the Saturn V or the Soviet Energia, but less than the Soviet N-1. Projected to place approximately 180 metric tons into orbit, the Ares V would have surpassed the Saturn V in payload capability. An upper stage, the Earth Departure Stage, would have utilized a more advanced version of the J-2 engine, the J-2X. Ares V would have placed the Altair lunar landing vehicle into low Earth orbit. An Orion crew vehicle launched on Ares I would have docked with Altair, and the Earth Departure Stage would then send the combined stack to the Moon.
After the cancellation of the Constellation Program – and hence Ares I and Ares V – NASA announced the Space Launch System (SLS) heavy-lift launch vehicle for deep-space exploration. The SLS, similar to the original Ares V concept, will be powered by four SSMEs and two five-segment SRBs. Its Block I configuration will lift approximately 70 metric tons to low Earth orbit. Block IB will add a second stage, the Exploration Upper Stage, powered by four RL10 engines, to increase payload to LEO and deep space. An eventual Block II variant will upgrade to advanced boosters, increasing LEO payload to at least 130 metric tons.
One proposal for advanced boosters would use a derivative of the Saturn V’s F-1, the F-1B, and increase SLS payload to around 150 metric tons to LEO. The F-1B is to have better specific impulse and be cheaper than the F-1, with a simplified combustion chamber and fewer engine parts, while producing 1,800,000 lbf (8.0 MN) of thrust at sea level, an increase over the approximate 1,550,000 lbf (6.9 MN) achieved by the mature Apollo 15 F-1 engine,
NASA SLS deputy project manager Jody Singer of the Marshall Space Flight Center in Huntsville, in 2012 stated that the vehicle will have a launch cost of approximately $500 million per launch, with a relatively minor dependence of costs on launch capability.
Saturn V displays
List of lunar probes
This is a list of space probes that have flown by, impacted, or landed on the Moon for the purpose of lunar exploration, as well as probes launched toward the Moon that failed to reach their target. Confirmed future probes are included, but missions that are still at the concept stage, or which never progressed beyond the concept stage, are not.
The list does not include the manned Apollo missions.
Lunar probes by date
|Pioneer 0||DOD||17 August 1958||orbiter||failure||first attempted launch beyond Earth orbit; launch vehicle failure; maximum altitude 16 km|||
|Luna E-1 No.1||USSR||23 September 1958||impactor||failure||launch vehicle failure|||
|Pioneer 1|| NASA/
|11 October 1958||orbiter||failure||second stage premature shutdown; maximum altitude 113,800 km; some data returned|||
|Luna E-1 No.2||USSR||12 October 1958||impactor||failure||launch vehicle failure|||
|Pioneer 2|| NASA/
|8 November 1958||orbiter||failure||third stage failure; maximum altitude 1,550 km; some data returned|||
|Luna E-1 No.3||USSR||4 December 1958||impactor||failure||launch vehicle failure|||
|Pioneer 3|| NASA/
|6 December 1958||flyby||failure||fuel depletion; maximum altitude 102,360 km; some data returned|||
|Luna 1||USSR||4 January 1959||flyby||partial success||first spacecraft in the vicinity of the Moon (flew within 5,995 km, but probably an intended impactor)|||
|Luna E-1A No.1||USSR||18 June 1959||impactor||failure||failed to reach Earth orbit|||
|Pioneer 4|| NASA/
|4 March 1959||flyby||partial success||achieved distant flyby; first US probe to enter solar orbit|||
|Luna 2||USSR||14 September 1959||impactor||success||first impact on Moon|||
|Pioneer P-1||NASA||24 September 1959?||orbiter?||failure||designation sometimes given to a failed launch or launchpad explosion during testing; conflicting information between sources|
|Luna 3||USSR||6 October 1959||flyby||success||first images from the lunar farside|||
|Pioneer P-3||NASA||26 November 1959||orbiter||failure||disintegrated shortly after launch|||
|Luna 1960A†||USSR||15 April 1960||flyby||failure||failed to attain correct trajectory|||
|Luna 1960B†||USSR||16 April 1960||flyby||failure||launch vehicle failure|||
|Pioneer P-30||NASA||25 September 1960||orbiter||failure||second stage failure; failed to reach Earth orbit|||
|Pioneer P-31||NASA||15 December 1960||orbiter||failure||first stage failure|||
|Ranger 3||NASA||28 January 1962||impactor||failure||missed target|||
|Ranger 4||NASA||26 April 1962||impactor||failure||hit the lunar farside; no data returned|||
|Ranger 5||NASA||21 October 1962||impactor||failure||power failure, missed target|||
|Sputnik 25||USSR||5 January 1963||lander||failure||failed to escape Earth orbit|||
|Luna 1963B†||USSR||2 February 1963||lander?||failure||failed to reach Earth orbit|||
|Luna 4||USSR||5 April 1963||lander?||failure||missed target, became Earth satellite|||
|Ranger 6||NASA||2 February 1964||impactor||partial success||impacted, but no pictures returned due to power failure|||
|Luna 1964A†||USSR||21 March 1964||lander||failure||failed to reach Earth orbit|||
|Luna 1964B†||USSR||20 April 1964||lander||failure||failed to reach Earth orbit|||
|Ranger 7||NASA||31 July 1964||impactor||success||returned pictures until impact|||
|Ranger 8||NASA||20 February 1965||impactor||success||returned pictures until impact|||
|Cosmos 60||USSR||12 March 1965||lander||failure||failed to leave Earth orbit|||
|Ranger 9||NASA||24 March 1965||impactor||success||TV broadcast of live pictures until impact|||
|Luna 1965A†||USSR||10 April 1965||lander?||failure||failed to reach Earth orbit?|||
|Luna 5||USSR||12 May 1965||lander||failure||crashed into Moon|||
|Luna 6||USSR||8 June 1965||lander||failure||missed Moon|||
|Zond 3||USSR||20 July 1965||flyby||success||possibly originally intended as a Mars probe, but target changed after launch window missed|||
|Luna 7||USSR||7 October 1965||lander||failure||crashed into Moon|||
|Luna 8||USSR||6 December 1965||lander||failure||crashed into Moon|||
|Luna 9||USSR||3 February 1966 –
6 February 1966
|lander||success||first soft landing; first images from the surface|||
|Cosmos 111||USSR||1 March 1966||orbiter||failure||failed to escape Earth orbit|||
|Luna 10||USSR||3 April 1966 –
30 May 1966
|orbiter||success||first artificial satellite of the moon|||
|Luna 1966A†||USSR||30 April 1966||orbiter?||failure||failed to reach Earth orbit|||
|Surveyor 1||NASA||2 June 1966||lander||success||first US soft landing; Surveyor program performed various tests in support of forthcoming manned landings|||
|Explorer 33||NASA||1 July 1966 –
15 September 1971
|orbiter||partial success||studied interplanetary plasma, cosmic rays, magnetic fields and solar X rays; failed to attain lunar orbit as intended, but achieved mission objectives from Earth orbit|||
|Lunar Orbiter 1||NASA||14 August 1966 –
29 October 1966
|orbiter||success||photographic mapping of lunar surface; intentionally impacted after completion of mission|||
|Luna 11||USSR||28 August 1966 –
1 October 1966
|orbiter||success||gamma-ray and X-ray-based observations of Moon’s composition; gravity, radiation and meteorite studies|||
|Surveyor 2||NASA||23 September 1966||lander||failure||crashed into Moon|||
|Luna 12||USSR||25 October 1966 –
19 January 1967
|orbiter||success||lunar surface photography|||
|Lunar Orbiter 2||NASA||10 November 1966 –
11 October 1967
|orbiter||success||photographic mapping of lunar surface; intentionally impacted after completion of mission|||
|Luna 13||USSR||24 December 1966||lander||success||TV pictures of lunar landscape; soil measurements|||
|Lunar Orbiter 3||NASA||8 February 1967 –
9 October 1967
|orbiter||success||photographic mapping of lunar surface; intentionally impacted after completion of mission|||
|Surveyor 3||NASA||20 April 1967 –
4 May 1967
|lander||success||various studies, primarily in support of forthcoming manned landings|||
|Lunar Orbiter 4||NASA||May–October 1967||orbiter||success||lunar photographic survey|||
|Explorer 35||NASA||July 1967 –
24 June 1973
|orbiter||success||studies of interplanetary plasma, magnetic fields, energetic particles and solar X rays|||
|Surveyor 4||NASA||17 July 1967||lander||failure||crashed into Moon|||
|Lunar Orbiter 5||NASA||5 August 1967 –
31 January 1968
|orbiter||success||lunar photographic survey; intentionally impacted after completion of mission|||
|Surveyor 5||NASA||11 September 1967 –
17 December 1967
|lander||success||various studies, primarily in support of forthcoming manned landings|||
|Zond 1967A†||USSR||28 September 1967||failure||lunar capsule test flight; launch failure|||
|Surveyor 6||NASA||10 November 1967 –
14 December 1967
|lander||success||various studies, primarily in support of forthcoming manned landings|||
|Zond 1967B†||USSR||22 November 1967||failure||lunar capsule test flight; launch failure|||
|History of IBM mainframes, 1952–present|
IBM mainframes are large computer systems produced by IBM since 1952. During the 1960s and 1970s, the term mainframe computer was almost synonymous with IBM products due to their marketshare. Current mainframes in IBM’s line of business computers are developments of the basic design of the IBM System/360.
First and second generation
From 1952 into the late 1960s, IBM manufactured and marketed several large computer models, known as the IBM 700/7000 series. The first-generation 700s were based on vacuum tubes, while the later, second-generation 7000s used transistors. These machines established IBM’s dominance in electronic data processing (“EDP”). IBM had two model categories: one (701, 704, 709, 7090, 7040) for engineering and scientific use, and one (702, 705, 705-II, 705-III, 7080, 7070, 7010) for commercial or data processing use. The two categories, scientific and commercial, generally used common peripherals but had completely different instruction sets, and there were incompatibilities even within each category.
IBM initially sold its computers without any software, expecting customers to write their own; programs were manually initiated, one at a time. Later, IBM provided compilers for the newly developed higher-level programming languages Fortran and COBOL. The first operating systems for IBM computers were written by IBM customers who did not wish to have their very expensive machines ($2M USD in the mid-1950s) sitting idle while operators set up jobs manually. These first operating systems were essentially scheduled work queues. It is generally thought that the first operating system used for real work was GM-NAA I/O, produced by General Motors’ Research division in 1956. IBM enhanced one of GM-NAA I/O’s successors, the SHARE Operating System, and provided it to customers under the name IBSYS. As software became more complex and important, the cost of supporting it on so many different designs became burdensome, and this was one of the factors which led IBM to develop System/360 and its operating systems.
The second generation (transistor-based) products were a mainstay of IBM’s business and IBM continued to make them for several years after the introduction of the System/360. (Some IBM 7094s remained in service into the 1980s.)
Prior to System/360, IBM also sold computers smaller in scale that were not considered mainframes, though they were still bulky and expensive by modern standards. These included:
- IBM 650 (vacuum tube logic, decimal architecture, drum memory, business and scientific)
- IBM 305 RAMAC (vacuum tube logic, first computer with disk storage; see: Early IBM disk storage)
- IBM 1400 series (business data processing; very successful and many 1400 peripherals were used with the 360s)
- IBM 1620 (decimal architecture, engineering, scientific, and education)
IBM had difficulty getting customers to upgrade from the smaller machines to the mainframes because so much software had to be rewritten. The 7010 was introduced in 1962 as a mainframe-sized 1410. The later Systems 360 and 370 could emulate the 1400 machines. A desk size machine with a different instruction set, the IBM 1130, was released concurrently with the System/360 to address the niche occupied by the 1620. It used the same EBCDIC character encoding as the 360 and was mostly programmed in Fortran, which was relatively easy to adapt to larger machines when necessary.
Midrange computer is a designation used by IBM for a class of computer systems which fall in between mainframes and microcomputers.
All that changed with the announcement of the System/360 (S/360) in April, 1964. The System/360 was a single series of compatible models for both commercial and scientific use. The number “360” suggested a “360 degree,” or “all-around” computer system. System/360 incorporated features which had previously been present on only either the commercial line (such as decimal arithmetic and byte addressing) or the engineering and scientific line (such as floating point arithmetic). Some of the arithmetic units and addressing features were optional on some models of the System/360. However, models were upward compatible and most were also downward compatible. The System/360 was also the first computer in wide use to include dedicated hardware provisions for the use of operating systems. Among these were supervisor and application mode programs and instructions, as well as built-in memory protection facilities. Hardware memory protection was provided to protect the operating system from the user programs (tasks) and the user tasks from each other. The new machine also had a larger address space than the older mainframes, 24 bits addressing 8-bit bytes vs. a typical 18 bits addressing 36-bit words.
The smaller models in the System/360 line (e.g. the 360/30) were intended to replace the 1400 series while providing an easier upgrade path to the larger 360s. To smooth the transition from second generation to the new line, IBM used the 360’s microprogramming capability to emulate the more popular older models. Thus 360/30s with this added cost feature could run 1401 programs and the larger 360/65s could run 7094 programs. To run old programs, the 360 had to be halted and restarted in emulation mode. Many customers kept using their old software and one of the features of the later System/370 was the ability to switch to emulation mode and back under operating system control.
Operating systems for the System/360 family included OS/360 (with PCP, MFT, and MVT), BOS/360, TOS/360, and DOS/360.
The System/360 later evolved into the System/370, the System/390, and the 64-bit zSeries, System z, and zEnterprise machines. System/370 introduced virtual memory capabilities in all models other than the very first System/370 models; the OS/VS1 variant of OS/360 MFT, the OS/VS2 (SVS) variant of OS/360 MVT, and the DOS/VS variant of DOS/360 were introduced to use the virtual memory capabilities, followed by MVS, which, unlike the earlier virtual-memory operating systems, ran separate programs in separate address spaces, rather than running all programs in a single virtual address space. The virtual memory capabilities also allowed the system to support virtual machines; the VM/370 hypervisor would run one or more virtual machines running either standard System/360 or System/370 operating systems or the single-user Conversational Monitor System (CMS). A time-sharing VM system could run multiple virtual machines, one per user, with each virtual machine running an instance of CMS [Wikipedia_4].
[Wikipedia_5] Multiprogramming with a Variable number of Tasks (MVT) was the most sophisticated of three available configurations of OS/360’s control program, and one of two available configurations in the final releases. MVT was intended for the largest machines in the System/360 family. Introduced in 1964, it did not become available until 1967. Early versions had many problems and the simpler MFT continued to be used for many years. Experience indicated that it was not advisable to install MVT on systems with less than 512 KiB of memory [Wikipedia_5].
A guidance system is a virtual or physical device, or a group of devices implementing a guidance process used for controlling the movement of a ship, aircraft, missile, rocket, satellite, or any other moving object. Guidance is the process of calculating the changes in position, velocity, attitude, and/or rotation rates of a moving object required to follow a certain trajectory and/or attitude profile based on information about the object’s state of motion.A guidance system is usually part of a Guidance, navigation and control system, whereas navigation refers to the systems necessary to calculate the current position and orientation based on sensor data like those from compasses, GPS receivers, Loran-C, star trackers, inertial measurement units, altimeters, etc. The output of the navigation system, the navigation solution, is an input for the guidance system, among others like the environmental conditions (wind, water, temperature, etc.) and the vehicle’s characteristics (i.e. mass, control system availability, control systems correlation to vector change, etc.). In general, the guidance system computes the instructions for the control system, which comprises the object’s actuators (e.g., thrusters, reaction wheels, body flaps, etc.), which are able to manipulate the flight path and orientation of the object without direct or continuous human control.
One of the earliest examples of a true guidance system is that used in the German V-1 during World War II. The navigation system consisted of a simple gyroscope, an airspeed sensor, and an altimeter. The guidance instructions were target altitude, target velocity, cruise time, engine cut off time.
A guidance system has three major sub-sections: Inputs, Processing, and Outputs. The input section includes sensors, course data, radio and satellite links, and other information sources. The processing section, composed of one or more CPUs, integrates this data and determines what actions, if any, are necessary to maintain or achieve a proper heading. This is then fed to the outputs which can directly affect the system’s course. The outputs may control speed by interacting with devices such as turbines, and fuel pumps, or they may more directly alter course by actuating ailerons, rudders, or other devices.
Inertial guidance systems were originally developed for rockets. American rocket pioneer Robert Goddard experimented with rudimentary gyroscopic systems. Dr. Goddard’s systems were of great interest to contemporary German pioneers including Wernher von Braun. The systems entered more widespread use with the advent of spacecraft, guided missiles, and commercial airliners.
US guidance history centers around 2 distinct communities. One driven out of Caltech and NASA Jet Propulsion Laboratory, the other from the German scientists that developed the early V2 rocket guidanceand MIT. The GN&C system for V2 provided many innovations and was the most sophisticated military weapon in 1942 using self-contained closed loop guidance. Early V2s leveraged 2 gyroscopes and lateral accelerometer with a simple analog computer to adjust the azimuth for the rocket in flight. Analog computer signals were used to drive 4 external rudders on the tail fins for flight control. Von Braun engineered the surrender of 500 of his top rocket scientists, along with plans and test vehicles, to the Americans. They arrived in Fort Bliss, Texas in 1945 and were subsequently moved to Huntsville, Al in 1950 (aka Redstone arsenal). Von Braun’s passion was interplanetary space flight. However his tremendous leadership skills and experience with the V-2 program made him invaluable to the US military. In 1955 the Redstone team was selected to put America’s first satellite into orbit putting this group at the center of both military and commercial space.
The Jet Propulsion Laboratory traces its history from the 1930s, when Caltech professor Theodore von Karman conducted pioneering work in rocket propulsion. Funded by Army Ordnance in 1942, JPL’s early efforts would eventually involve technologies beyond those of aerodynamics and propellant chemistry. The result of the Army Ordnance effort was JPL’s answer to the German V-2 missile, named MGM-5 Corporal, first launched in May 1947. On December 3, 1958, two months after the National Aeronautics and Space Administration (NASA) was created by Congress, JPL was transferred from Army jurisdiction to that of this new civilian space agency. This shift was due to the creation of a military focused group derived from the German V2 team. Hence, beginning in 1958, NASA JPL and the Caltech crew became focused primarily on unmanned flight and shifted away from military applications with a few exceptions. The community surrounding JPL drove tremendous innovation in telecommunication, interplanetary exploration and earth monitoring (among other areas).
In the early 1950s, the US government wanted to insulate itself against over dependency on the Germany team for military applications. Among the areas that were domestically “developed” was missile guidance. In the early 1950s the MIT Instrumentation Laboratory (later to become the Charles Stark Draper Laboratory, Inc.) was chosen by the Air Force Western Development Division to provide a self-contained guidance system backup to Convair in San Diego for the new Atlas intercontinental ballistic missile. The technical monitor for the MIT task was a young engineer named Jim Fletcher who later served as the NASA Administrator. The Atlas guidance system was to be a combination of an on-board autonomous system, and a ground-based tracking and command system. This was the beginning of a philosophic controversy, which, in some areas, remains unresolved. The self-contained system finally prevailed in ballistic missile applications for obvious reasons. In space exploration, a mixture of the two remains.
In the summer of 1952, Dr. Richard Battin and Dr. J. Halcombe (“Hal”) Laning Jr., researched computational based solutions to guidance as computing began to step out of the analog approach. As computers of that time were very slow (and missiles very fast) it was extremely important to develop programs that were very efficient. Dr. J. Halcombe Laning, with the help of Phil Hankins and Charlie Werner, initiated work on MAC, an algebraic programming language for the IBM 650, which was completed by early spring of 1958. MAC became the work-horse of the MIT lab. MAC is an extremely readable language having a three-line format, vector-matrix notations and mnemonic and indexed subscripts. Today’s Space Shuttle (STS) language called HAL, (developed by Intermetrics, Inc.) is a direct offshoot of MAC. Since the principal architect of HAL was Jim Miller, who co-authored with Hal Laning a report on the MAC system, it is a reasonable speculation that the space shuttle language is named for Jim’s old mentor, and not, as some have suggested, for the electronic superstar of the Arthur Clarke movie “2001-A Space Odyssey.” (Richard Battin, AIAA 82-4075, April 1982)
Hal Laning and Richard Battin undertook the initial analytical work on the Atlas inertial guidance in 1954. Other key figures at Convair were Charlie Bossart, the Chief Engineer, and Walter Schweidetzky, head of the guidance group. Walter had worked with Wernher von Braun at Peenemuende during World War II.
The initial “Delta” guidance system assessed the difference in position from a reference trajectory. A velocity to be gained (VGO) calculation is made to correct the current trajectory with the objective of driving VGO to Zero. The mathematics of this approach were fundamentally valid, but dropped because of the challenges in accurate inertial navigation (e.g. IMU Accuracy) and analog computing power. The challenges faced by the “Delta” efforts were overcome by the “Q system” of guidance. The “Q” system’s revolution was to bind the challenges of missile guidance (and associated equations of motion) in the matrix Q. The Q matrix represents the partial derivatives of the velocity with respect to the position vector. A key feature of this approach allowed for the components of the vector cross product (v, xdv,/dt) to be used as the basic autopilot rate signals-a technique that became known as “cross-product steering.” The Q-system was presented at the first Technical Symposium on Ballistic Missiles held at the Ramo-Wooldridge Corporation in Los Angeles on June 21 and 22, 1956. The “Q System” was classified information through the 1960s. Derivations of this guidance are used for today’s military missiles. The CSDL team remains a leader in the military guidance and is involved in projects for most divisions of the US military.
On August 10 of 1961 NASA Awarded MIT a contract for preliminary design study of a guidance and navigation system for Apollo program. (see Apollo on-board guidance, navigation, and control system, Dave Hoag, International Space Hall of Fame Dedication Conference in Alamogordo, N.M., October 1976 ). Today’s space shuttle guidance is named PEG4 (Powered Explicit Guidance). It takes into account both the Q system and the predictor-corrector attributes of the original “Delta” System (PEG Guidance). Although many updates to the shuttles navigation system have taken place over the last 30 years (ex. GPS in the OI-22 build), the guidance core of today’s Shuttle GN&C system has evolved little. Within a manned system, there is a human interface needed for the guidance system. As Astronauts are the customer for the system, many new teams are formed that touch GN&C as it is a primary interface to “fly” the vehicle. For the Apollo and STS (Shuttle system) CSDL “designed” the guidance, McDonnell Douglas wrote the requirements and IBM programmed the requirements.
Much system complexity within manned systems is driven by “redundancy management” and the support of multiple “abort” scenarios that provide for crew safety. Manned US Lunar and Interplanetary guidance systems leverage many of the same guidance innovations (described above) developed in the 1950s. So while the core mathematical construct of guidance has remained fairly constant, the facilities surrounding GN&C continue to evolve to support new vehicles, new missions and new hardware. The center of excellence for the manned guidance remains at MIT (CSDL) as well as the former McDonnell Douglas Space Systems (in Houston).
Guidance systems consist of 3 essential parts: navigation which tracks current location, guidance which leverages navigation data and target information to direct flight control “where to go”, and control which accepts guidance commands to effect change in aerodynamic and/or engine controls.
Navigation is the art of determining where you are, a science that has seen tremendous focus in 1711 with the Longitude prize. Navigation aids either measure position from a fixed point of reference (ex. landmark, north star, LORAN Beacon), relative position to a target (ex. radar, infra-red, …) or track movement from a known position/starting point (e.g. IMU). Today’s complex systems use multiple approaches to determine current position. For example, today’s most advanced navigation systems are embodied within the Anti-ballistic missile, the RIM-161 Standard Missile 3 leverages GPS, IMU and ground segmentdata in the boost phase and relative position data for intercept targeting. Complex systems typically have multiple redundancy to address drift, improve accuracy (ex. relative to a target) and address isolated system failure. Navigation systems therefore take multiple inputs from many different sensors, both internal to the system and/or external (ex. ground based update). Kalman filter provides the most common approach to combining navigation data (from multiple sensors) to resolve current position. Example navigation approaches:
- Celestial navigation is a position fixing technique that was devised to help sailors cross the featureless oceans without having to rely on dead reckoning to enable them to strike land. Celestial navigation uses angular measurements (sights) between the horizon and a common celestial object. The Sun is most often measured. Skilled navigators can use the Moon, planets or one of 57 navigational stars whose coordinates are tabulated in nautical almanacs. Historical tools include a sextant, watch and ephemeris data. Today’s space shuttle, and most interplanetary spacecraft, use optical systems to calibrate inertial navigation systems: Crewman Optical Alignment Sight (COAS), Star Tracker.
- Inertial Measurement Units (IMUs) are the primary inertial system for maintaining current position (navigation) and orientation in missiles and aircraft. They are complex machines with one or more rotating Gyroscopes that can rotate freely in 3 degrees of motion within a complex gimbal system. IMUs are “spun up” and calibrated prior to launch. A minimum of 3 separate IMUs are in place within most complex systems. In addition to relative position, the IMUs contain accelerometers which can measure acceleration in all axes. The position data, combined with acceleration data provide the necessary inputs to “track” motion of a vehicle. IMUs have a tendency to “drift”, due to friction and accuracy. Error correction to address this drift can be provided via ground link telemetry, GPS, radar, optical celestial navigationand other navigation aids. When targeting another (moving) vehicle, relative vectors become paramount. In this situation, navigation aids which provide updates of position relative to the target are more important. In addition to the current position, inertial navigation systems also typically estimate a predicted position for future computing cycles. See also Inertial navigation system.
- Astro-inertial guidance is a sensor fusion/information fusion of the Inertial guidance and Celestial navigation.
- Long-range Navigation (LORAN) : This was the predecessor of GPS and was (and to an extent still is) used primarily in commercial sea transportation. The system works by triangulating the ship’s position based on directional reference to known transmitters.
- Global Positioning System (GPS) : GPS was designed by the US military with the primary purpose of addressing “drift” within the inertial navigation of Submarine-launched ballistic missile(SLBMs) prior to launch. GPS transmits 2 signal types: military and a commercial. The accuracy of the military signal is classified but can be assumed to be well under 0.5 meters. GPS is a system of 24 satellites orbiting in unique planes 10.9-14.4 Nautical miles above the earth. The Satellites are in well defined orbits and transmit highly accurate time information which can be used to triangulate position.
- Radar/Infrared/Laser : This form of navigation provides information to guidance relative to a known target, it has both civilian (ex rendezvous) and military applications.
- active (employs own radar to illuminate the target),
- passive (detects target’s radar emissions),
- semiactive radar homing,
- Infrared homing : This form of guidance is used exclusively for military munitions, specifically air-to-air and surface-to-air missiles. The missile’s seeker head homes in on the infrared (heat) signature from the target’s engines (hence the term “heat-seeking missile”),
- Ultraviolet homing, used in FIM-92 Stinger – more resistive to countermeasures, than IR homing system
- Laser guidance : A laser designator device calculates relative position to a highlighted target. Most are familiar with the military uses of the technology on Laser-guided bomb. The space shuttle crew leverages a hand held device to feed information into rendezvous planning. The primary limitation on this device is that it requires a line of sight between the target and the designator.
- Terrain contour matching (TERCOM). Uses a ground scanning radar to “match” topography against digital map data to fix current position. Used by cruise missiles such as the Tomahawk (missile).
Guidance is the “driver” of a vehicle. It takes input from the navigation system (where am I) and uses targeting information (where do I want to go) to send signals to the flight control system that will allow the vehicle to reach its destination (within the operating constraints of the vehicle). The “targets” for guidance systems are one or more state vectors (position and velocity) and can be inertial or relative. During powered flight, guidance is continually calculating steering directions for flight control. For example, the space shuttle targets an altitude, velocity vector, and gamma to drive main engine cut off. Similarly, an Intercontinental ballistic missile also targets a vector. The target vectors are developed to fulfill the mission and can be preplanned or dynamically created.
Control. Flight control is accomplished either aerodynamically or through powered controls such as engines. Guidance sends signals to flight control. A Digital Autopilot (DAP) is the interface between guidance and control. Guidance and the DAP are responsible for calculating the precise instruction for each flight control. The DAP provides feedback to guidance on the state of flight controls [Wikipedia_6].
Unified S-band Radio Frequency System
[Wikipedia_7] The Unified S-band (USB) system was a tracking and communication system developed for the Apollo program by NASA and the Jet Propulsion Laboratory (JPL). It operated in the S band portion of the microwave spectrum, unifying voice communications, television, telemetry, command, tracking and ranging into a single system to save size and weight and simplify operations. The USB ground network was managed by the Goddard Space Flight Center (GSFC). Commercial contractors included Collins Radio, Blaw-Knox, Motorola and Energy Systems.
The previous programs, Mercury and Gemini, had separate radio systems for voice, telemetry, and tracking. Uplink voice and command, and downlink voice and telemetry data were sent via ultra high frequency (UHF) and very high frequency (VHF) systems. The tracking capability was a C band beacon interrogated by a ground-based radar. With the much greater distance of Apollo, passive ranging was not feasible, so a new active ranging system was required. Apollo also planned to use television transmissions, which were not supported by the existing systems. Finally, the use of three different frequencies complicated the spacecraft systems and ground support. The Unified S-band (USB) system was developed to address these concerns.
The USB system did not completely replace all other radio transmitters on Apollo. While it was the sole mode for deep space communications, Apollo still used VHF for short range voice and low rate telemetry between astronauts and the Lunar Module (LM) and lunar rover during extra-vehicular activity; between the LM and Command/Service Module (CSM or CM); and between the CSM and Earth stations during the orbital and recovery phases of the mission. The CM had a backup capability to range the LM over its VHF voice links.
Apollo also carried several radars that operated independently of the USB on their own frequencies, including the landing and rendezvous radars on the LM and a C-band radar transponder on the CM.
From a NASA technical summary:
The design of the USB system is based on a coherent doppler and the pseudo-random range system which has been developed by JPL. The S-band system utilizes the same techniques as the existing systems, with the major changes being the inclusion of the voice and data channels.
A single carrier frequency is utilized in each direction for the transmission of all tracking and communications data between the spacecraft and ground. The voice and update data are modulated onto subcarriers and then combined with the ranging data […]. This composite information is used to phase-modulate the transmitted carrier frequency. The received and transmitted carrier frequencies are coherently related. This allows measurements of the carrier doppler frequency by the ground station for determination of the radial velocity of the spacecraft.
In the transponder the subcarriers are extracted from the RF carrier and detected to produce the voice and command information. The binary ranging signals, modulated directly onto the carrier, are detected by the wide-band phase detector and translated to a video signal.
The voice and telemetry data to be transmitted from the spacecraft are modulated onto subcarriers, combined with the video ranging signals, and used to phase-modulate the downlink carrier frequency. The transponder transmitter can also be frequency modulated for the transmission of television information or recorded data instead of ranging signals.
The basic USB system has the ability to provide tracking and communications data for two spacecraft simultaneously, provided they are within the beamwidth of the single antenna. The primary mode of tracking and communications is through the use of the PM mode of operation. Two sets of frequencies separated by approximately 5 megacycles are used for this purpose […]. In addition to the primary mode of communications, the USB system has the capability of receiving data on two other frequencies. These are used primarily for the transmission of FM data from the spacecraft.
The Unified S-Band System used the 2025-2120 MHz band for uplinks (earth to space transmissions) and the 2200-2290 MHz band for downlinks (space to earth transmissions). Both bands are allocated internationally for space research and operations, though by 2014 standards the ALSEP uplink was in the wrong part of the band (Deep Space instead of Near Earth).
|Apollo S-band frequency assignments|
|Spacecraft||Uplink (MHz)||Downlink (MHz)||Coherent ratio|
|Command Module PM||2106.40625||2287.5||221/240|
|Command Module FM||2272.5|
|Apollo 11 Early ALSEP||2119||2276.5|
|Apollo 12 ALSEP||2119||2278.5|
|Apollo 14 ALSEP||2119||2279.5|
|Apollo 15 ALSEP||2119||2278.0|
|Apollo 15 subsatellite||2101.802083||2282.5||221/240|
|Apollo 16 ALSEP||2119||2276.0|
|Apollo 17 ALSEP||2119||2275.5|
Each Apollo spacecraft was assigned a frequency pair. For certain phase modulation (PM) downlinks, the uplink to downlink frequency ratio was exactly 221/240. The ALSEP lunar surface experiments shared a common uplink and did not, insofar as is known, implement a coherent transponder. (The passive laser retroreflectors left by the Apollo 11, 14 and 15 missions provide much greater accuracy, and have far outlived the active electronics in the other ALSEP experiments.) The Lunar Communications Relay Unit (LCRU) on the Lunar Rover had its own downlink frequency (to avoid interference with the LM) but shared the LM’s uplink frequency as it did not implement a coherent transponder. To keep the VHF transmitters on the LM and LCRU from both trying to relay uplink voice and interfering with each other, separate voice subcarriers were used on the common S-band uplink: 30 kHz for the LM and 124 kHz for the LCRU.
The CSM had two separate transmitters, one PM and one FM. The LM had only one S-band transmitter that could operate in PM or FM, but not both simultaneously.
The S-IVB upper stage had its own USB transponder so it could be tracked independently after Apollo spacecraft separation until the stage either flew past the moon (Apollos 8, 10, 11, 12) or, starting with Apollo 13, hit the moon. This tracking data greatly aided the analysis of the impact as recorded by the seismometers left by earlier Apollo crews.
The S-IVB shared its S-band frequency pair with the LM. This created no problem in a normal mission as the LM remained dormant until lunar orbit, by which time the S-IVB had already hit the moon or flown off into orbit around the sun. However, it created an operational problem during the Apollo 13 mission when the LM had to be used as a lifeboat well before Apollo and the S-IVB reached the moon.
The LM frequency pair was also used by the subsatellites left in lunar orbit by the later J-missions. They were deployed by the CSM shortly before leaving lunar orbit and after the LM had completed its mission.
The use of two separated frequency bands made full duplex operation possible. The ground and the spacecraft transmitted continuously. Microphone audio was keyed either manually or by VOX, but unlike ordinary half duplex two-way radio both sides could talk at the same time without mutual interference.
The S-band uplinks and downlinks usually (but not always) used phase modulation (PM). PM, like FM, has a constant amplitude (envelope) regardless of modulation. This permits the use of nonlinear RF amplifiers that can be considerably more efficient than RF amplifiers that must maintain linearity.
The PM modulation index is small, on the order of 1 radian, so the modulated signal more closely resembled double sideband amplitude modulation (AM) except for the carrier phase. A PM signal can be approximated for analysis purposes as an AM signal with the carrier (and only the carrier) rotated 90 degrees from its original phase. One important difference is that in AM, the carrier component has a constant amplitude as the sidebands vary with modulation while in PM the total signal (the envelope) is constant amplitude. This means that PM shifts power from the carrier to the information-carrying sidebands with modulation, and at some modulation indices the carrier can disappear completely. This is why Apollo uses a low modulation index: to leave a strong carrier that can be used for highly accurate velocity tracking by measurement of its Doppler shift.
Coherent transponders and Doppler tracking
Allocating uplink/downlink frequency pairs in a fixed ratio of 221/240 permitted the use of coherent transponders on the spacecraft. The spacecraft tracked the uplink carrier with a phase locked loop and, with a series of frequency dividers and multipliers, multiplied the uplink carrier frequency by the ratio 240/221 to produce its own downlink carrier signal.
When no uplink was detected, the transponder downlink carrier was generated from a local oscillator at the nominal frequency.
This “two-way” technique allowed extremely precise relative velocity measurements (in centimeters/sec) by observing the Doppler shift of the downlink carrier without a high accuracy oscillator on the spacecraft, although one was still needed on the ground.
As mentioned above, the uplink and downlink carriers played a critical role in spacecraft tracking. Sidebands generated by the information also carried by the system had to be kept away from the carriers to avoid perturbing the phase locked loops used to track them. This was done through the use of various subcarriers.
The uplink had subcarriers at 30 kHz and 70 kHz. The 30 kHz subcarrier was modulated with uplink (Capcom) voice using narrowband FM (NBFM) and the 70 kHz carrier was modulated with command data for the onboard computer. This latter capability, which could be blocked by the astronauts, was used primarily to update the state vectors maintained by the computers with accurate values determined by ground tracking. It was also used to execute maneuvers in an unmanned spacecraft, e.g., deorbiting the lunar module after it had been jettisoned in lunar orbit.
Either or both subcarriers could be turned off when not needed, e.g., the voice subcarrier could be turned off during astronaut sleep periods. This improved the signal margins for the other information streams such as telemetry data.
The downlink normally had subcarriers at 1.25 MHz (NBFM voice) and 1.024 MHz (telemetry data). The telemetry could be at one of two rates, 1.6 kilobits/sec (low rate, 1/640 of the subcarrier frequency) and 51.2 kilobits/sec (high rate, 1/20 of the subcarrier frequency). High rate was used unless low rate was forced by poor link conditions, e.g., the use of a small earth receiving antenna, an omni spacecraft antenna, or the need to conserve spacecraft power by turning off its RF power amplifier. (The S-band transponder on the S-IVB had no voice subcarrier.)
A “backup voice” mode was available that shut off the 1.25 MHz NBFM voice subcarrier and transmitted voice at baseband on the main PM S-band carrier. This provided a few more dB of margin when the link was unusually degraded but worse voice quality than the normal voice mode when conditions were good.
The two modes can be easily distinguished by how they react to signal fades. In the normal (NBFM subcarrier) voice mode the audio SNR is usually very high. But as the link degrades below threshold, impulse or “popcorn” noise appears suddenly and builds up rapidly until it overwhelmed the astronauts’ voices. An excellent example occurred during the Apollo 11 lunar landing when the lunar module structure occasionally obstructed the antenna’s view of Earth.
The backup voice mode behaved more like AM; there is a constant background “hiss” and the astronauts’ voices vary with signal strength. This mode was used extensively during the Apollo 13 emergency to conserve battery power in the LM Aquarius and during Apollo 16 because of the failure of the steerable S-band antenna on the lunar module Orion.
Voice transmissions used Quindar tones for in-band signaling.
The Apollo USB downlink also provided an “emergency key” capability consisting of a manually on-off keyed subcarrier oscillator at 512 kHz. Presumably this would have been used by the crew to transmit Morse Code if the downlink were too severely degraded to support even the backup voice mode. Although this mode had been tested (on Apollo 7) and most of the astronauts were trained in its use, this mode was never actually needed during any Apollo mission. (Apollo 13 made extensive use of the “backup voice” mode, as did the Apollo 16 lunar module Orion due to a failed high gain antenna).
A similar uplink capability did not exist because the uplink budget had far more margin than the downlink. A typical Apollo S-band spacecraft transmitter produced 20 watts; a typical uplink transmitter produced 10 kW, a ratio of 27 dB.
The Apollo S-band system provided for accurate range (distance) measurements. The ground station generated a pseudorandom noise (PN) sequence at 994 kilobit/s and added it to the baseband signal going to the PM transmitter. The transponder echoed this PN signal back to earth on the downlink, and by correlating the received and transmitted versions the precise round trip light time to the spacecraft could be determined very accurately (within 15 meters).
The PN sequence, although deterministic, had the properties of a random bit stream. Although the PN sequence was periodic, its period of about 5 seconds exceeded the largest possible round trip time to the moon so there would be no ambiguity in its received timing.
Modern GPS receivers work somewhat similarly in that they also correlate a received PN bit stream (at 1.023 Mbit/s) with a local reference to measure distance. But GPS is a receive-only system that uses relative timing measurements from a set of satellites to determine receiver position while the Apollo USB is a two-way system that can only determine the instantaneous distance and relative velocity. However, an orbit determination program can find the unique spacecraft state vector or orbital element set that most closely matches (usually in a least squares sense) a set of range, range-rate (relative velocity) and antenna look angle observations made over a period of time by one or more ground stations assuming purely ballistic spacecraft motion over the observation interval.
Once the state vector has been determined, the spacecraft’s future trajectory can be fully predicted until the next propulsive event.
Transponder ranging turn-around had to be manually enabled by an astronaut. It consumed an appreciable fraction of the downlink capacity and it was only needed occasionally, typically during handover from one ground station to the next. After the new uplink station achieved a 2-way coherent transponder lock in which the spacecraft generated its carrier at 240/221 times the received uplink frequency, the new ground station ranged the spacecraft. Then the ranging signal was turned off and the range measurement was continually updated by Doppler velocity measurements.
If for some reason a ground station lost lock during a pass, it would repeat the ranging measurement after re-acquiring lock.
FM and video
The normal operating mode of an Apollo S-band downlink transmitter was PM. This mode provided for coherent Doppler tracking, uplink commands, downlink telemetry and two-way voice—but not television. Video signals, even that from the slow scan camera used during the Apollo 11 EVA, are much wider in bandwidth than the other Apollo downlink signals. The PM link margin simply could not provide an acceptable picture, even when the largest available dishes were used.
A means was also needed to transmit wideband engineering and scientific data, such as that recorded on a tape recorder and played back at high speed.
The answer to both needs was wideband FM – frequency modulation. FM with a large modulation index exhibits a capture or threshold effect. The output signal-to-noise ratio (SNR) can be significantly greater than the RF channel SNR provided that the RF SNR remains above a threshold, typically around 8-10 dB.
This enhancement comes at a price: below the FM threshold, the output SNR is worse than the RF channel SNR. Reception is “all or nothing”; a receiving antenna too small to capture the video cannot capture the narrowband elements either (e.g., voice).
The CSM carried separate FM and PM transmitters that could operate simultaneously, so voice and telemetry continued to be transmitted by PM while the video came down by FM. The LM only carried a single transmitter that could operate in either FM or PM, but not both. FM cannot be used for Doppler tracking, so the LM always transmitted PM during flight, reserving FM for when video was required (e.g., during a surface EVA).
Until the transition to digital, satellite television also used wideband FM.
It is historically understood that the USSR did intercept the Apollo missions telemetry on the territory of the USSR, but until 2005, no source in the former USSR military or intelligence services has come forth with any evidence of this happening. The USSR used different frequency bands for its own space missions, so by default its deep space network did not readily have equipment able to receive Apollo telemetry. Whether China or any other non-Western (or non-aligned) country at the time chose to intercept any of the Apollo telemetry is unclear. Amateur radio and affiliated telecommunications sector persons could listen to the Apollo telemetry the world over—provided they could afford the reception equipment.
Within the territory of the US it was legally possible for amateur radio operators to monitor the telemetry, but the FCC did issue a directive that required all disclosure of Apollo telemetry interception be cleared by NASA. Paul Wilson and Richard T. Knadle, Jr. received voice transmissions from the Command Service Module of Apollo 15 in lunar orbit on the morning of August 1, 1971. In an article for QST magazine they provide a detailed description of their work, with photographs. At least two different radio amateurs, W4HHK and K2RIW, reported reception of Apollo 16 signals with home-built equipment [Wikipedia_7].
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