Chronology: Apollo-to-Shuttle Transition 2.0

Three years ago I published on this blog the first of my "Chronology" compilations of links to posts with a common theme. That first chronological compilation brought together links to posts on the transition from Apollo to the Space Shuttle. The aim was to impose chronology on posts that do not occur in chronological order in this blog as an aid to reader understanding.

This, my fourth "Chronology" compilation, updates that first compilation. I've added links to three posts dating from 1968, 1970, and 1972; that is, near the start, at the middle, and near the end of the planning phase of the Apollo-Shuttle transition.

"A True Gateway": Robert Gilruth's June 1968 Space Station Presentation

Series Development: A 1969 Plan to Merge Shuttle and Saturn V to Spread Out Space Program Cost (December 1969)

Think Big: A 1970 Flight Schedule for NASA's 1969 Integrated Program Plan (June 1970)

McDonnell Douglas Phase B Space Station (June 1970)

From Monolithic to Modular: NASA Establishes a Baseline Configuration for a Shuttle-Launched Space Station (July 1970)

An Alternate Station/Shuttle Evolution: The Spirit of '76 (August 1970)

Apollo's End: NASA Cancels Apollo 15 & Apollo 19 to Save Station/Shuttle (August-September 1970)

The Last Days of the Nuclear Shuttle (February 1971)

A Bridge From Skylab to Station/Shuttle: Interim Space Station Program (April 1971)

Where to Launch and Land the Space Shuttle? (April 1972)

"Still Under Active Consideration": Five Proposed Earth-Orbital Apollo Missions for the 1970s (August 1972)

The First Voyager (1967)

Artist concept of a Voyager spacecraft. Although the spacecraft design is correctly depicted, some liberties are taken with the planets (Mars and Earth are never this close together). The robotic Voyager includes a main engine derived from the Apollo Lunar Module Descent Propulsion System; a body-mounted, ring-shaped solar array; a skeletal high-gain antenna (it points toward Earth); a cylindrical body with rectangular thermal-control louver ports; and a Mars landing capsule with a conical heat shield sealed inside a "sterilization canister." The spacecraft is larger than it might appear: at this point in its mission, it would weigh 10.25 tons and measure 6.1 meters in diameter. Image credit: NASA/Jet Propulsion Laboratory.

In 1961, the Pasadena, California-based Jet Propulsion Laboratory (JPL), a spaceflight engineering laboratory managed by California Institute of Technology on contract to NASA, began study of Voyager, a robotic spacecraft program for exploring Mars and Venus in the late 1960s and 1970s. NASA Headquarters formally approved Project Voyager in 1964. Cuts in NASA's space science budget and debate over how Voyager should be managed and launched delayed NASA's push for a formal "new start" until January 1967, when President Lyndon Johnson's Fiscal Year (FY) 1968 NASA budget called for $71.5 million for the new program.

In January 1967, NASA's Office of Space Science and Applications published a 26-page brochure as part of its efforts to move Voyager from planning to development. The brochure was an introduction (and sales pitch) aimed at members of Congress and other individuals who would need to support Voyager if it was to become part of NASA's approved planetary exploration program for the 1970s.

In the foreword to the brochure, Homer Newell, NASA Associate Administrator for Space Science and Applications, explained that Voyager's chosen launch vehicle was the "awe-inspiring" Saturn V. One three-stage Saturn V rocket would launch two 12-ton Voyager spacecraft to Mars. For comparison, the Mariner IV Mars flyby spacecraft, launched on an Atlas-Agena D rocket in November 1964, had weighed only 260.4 kilograms. Newell wrote that

[s]uccesses already achieved in the 1960s with unmanned spacecraft of limited weight and power. . .foretell the great work of exploration that lies ahead. . .With Voyager, the U.S. capability for planetary exploration will grow by several orders of magnitude. . .Voyager could well be the means by which man first learns of extraterrestrial life.

NASA, the brochure explained, favored Mars over Venus as Voyager's first exploration target because "the high surface temperatures on Venus make the existence of extraterrestrial life less likely than on Mars" and because "the thin, normally transparent Martian atmosphere is conducive to detailed scanning of its surface features from orbit." In addition, "manned landings on Mars will someday be possible. . .[but] they may not be possible on Venus."

The brochure placed Voyager within an evolutionary robotic exploration program designed to take advantage of low-energy Earth-Mars transfer opportunities that occur every 26 months. It retroactively made Mariner IV, which had flown by Mars on 14-15 July 1965, the first mission in its program. Inclusion of Mariner IV, the first successful Mars explorer, is somewhat ironic, for its discoveries had helped to undermine support for Voyager.

In addition to recording for slow playback 21 black-and-white images that took in about 1% of the martian surface, Mariner IV had enabled Earth-bound scientists to measure martian atmospheric pressure by transmitting its feeble radio signal through the atmosphere as it passed behind the planet as viewed from Earth. Based on the degree of refraction of the signal, scientists had determined that surface pressure on Mars is not, as expected, about 10% of Earth sea-level pressure; it is, in fact, less than 1% of Earth sea-level pressure.

The Saturn IB with Centaur third stage was the first planned Voyager launch vehicle. Never fully developed, this booster would have launched a single Voyager Orbiter/Lander combination with a mass of up to 5440 kilograms toward Mars or Venus. The third stage includes the cylindrical lower part of the Payload Shroud. The Interstage, linking the S-IB first stage and the S-IVB second stage and covering the S-IVB's single J-2 engine, is generally considered to be part of the S-IVB stage. The IU is the Instrument Unit, which houses the guidance system for the rocket's first two stages. Image credit: NASA/heroicrelics.org/DSFPortree.

The brochure acknowledged that the new atmosphere data had forced a redesign of the Voyager landing system. The new design replaced lightweight parachutes with heavier landing rockets. According to historians Edward Clinton Ezell and Linda Neumann Ezell, writing in their 1984 NASA-published history On Mars: Exploration of the Red Planet, 1958-1978, the redesign bumped Voyager's projected cost above the psychologically significant $1 billion mark. Adoption of the Saturn V launch vehicle in place of the Saturn IB with a Centaur upper stage — a move designed to justify continuation of the Saturn V assembly line after Apollo and to provide flexibility in the event that Voyager redesigns significantly boosted its mass and heat shield diameter — pushed the program's price-tag past $2 billion.

The brochure called for new Mariner Mars flybys in 1969 and 1971. In 1969, a Mariner spacecraft would photograph the entire visible disk of Mars during approach and return detailed images of 10% of the planet. During the 1971 flyby, a Mariner would release a small sterilized probe into the martian atmosphere to measure pressure, density, temperature, and composition as it plummeted toward surface impact and destruction. The flyby spacecraft would act as a relay for probe signals and would image 10% of Mars at high resolution.

The first Voyager missions would take place in 1973. A battery-powered Voyager Lander with a mass of up to 390 kilograms would seek life and observe changes at the landing site over several days, and a solar-powered Voyager Orbiter would observe seasonal changes on a planet-wide scale for months.

The Voyager 1975 orbiters and landers would rely on Radioisotope Thermoelectric Generators (RTGs) for electricity. This would allow the landers to survive on Mars for one martian year (about two Earth years); that is, long enough for them to observe seasonal changes at their landing sites. Voyager could land up to 499 kilograms on Mars in the 1975 opportunity. The 1977 and 1979 Voyager missions would see introduction of a lander-deployed Mars surface rover and biological experiments specially designed to study any living things found in 1973 and 1975. A Voyager lander could deliver up to 680.4 kilograms to the surface of Mars in 1977 and 1979.

The brochure then detailed the 1973 Voyager Mars mission, which it described as typical. Voyagers would lift off from the Kennedy Space Center Complex 39 launch pads NASA built for the Apollo Saturn V launches. The 1970s Mars launch windows would last at least 25 days and would include daily one-hour launch opportunities. Voyager Saturn V rockets would be identical to Apollo lunar Saturn Vs; that is, each would comprise an S-IC first stage with five F-1 engines, an S-II second stage with five J-2 engines, and an S-IVB third stage with one J-2.

Voyager Saturn V rockets would have included the three stages and ring-shaped Instrument Unit of the Apollo Saturn V rocket. A cylindrical segmented shroud with a conical top containing twin Voyager spacecraft would, however, have replaced the Apollo spacecraft, which included the  Command and Service Module (CSM) and Lunar Module (LM) spacecraft, the Spacecraft Launch Adapter shroud containing the LM, and the pencil-shaped Launch Escape System tower. Image credit: NASA. 

The twin Voyager lander/orbiter combinations would be stacked atop the S-IVB third stage within a protective launch shroud. The first stage would burn for 2.5 minutes and fall away at an altitude of 62.8 kilometers, then the second stage would burn for 6.5 minutes and fall away at an altitude of 182.5 kilometers. The third stage would fire briefly to place itself, the twin Voyagers, and their launch shroud into Earth parking orbit.

Voyager's launch shroud would measure 6.7 meters in diameter — the same diameter as the S-IVB stage — and would have a mass of 4.7 tons. Once in Earth orbit, the shroud's conical top section would jettison, exposing the upper Voyager to space. The S-IVB stage would then ignite a second time to push the Voyagers out of Earth orbit toward Mars. After S-IVB shutdown, the upper Voyager would separate. The shroud's cylindrical central portion would then jettison to expose the lower Voyager, which would separate from the S-IVB a short time later. In the 1973 opportunity, each Voyager would have a mass of 10.25 tons after separation.

A model made in 1967 displays how two Voyager spacecraft would be stacked within the shroud atop the Saturn V S-IVB third stage. The conical, streamlined nose cone at the top of this shroud/spacecraft stack is absent. Image credit: NASA.

During the months-long interplanetary cruise, the twin Voyagers would turn their ring-shaped body-mounted solar arrays toward the Sun. They would use course-correction engines based on the Minuteman missile second-stage engine to place themselves on precise paths to Mars. The S-IVB stage trailing them would make no course adjustments, so would miss the planet by a wide margin. Because the Voyagers would perform their course corrections at different times, they would arrive at Mars up to 10 days apart.

As each Voyager neared Mars, it would fire its main rocket engine to slow down so that the planet's gravity could capture it into an elliptical orbit. Initial orbit periapsis (low point) would be about 1127 kilometers above the planet, while apoapsis (high point) would occur beyond the orbit of Deimos, the outer martian moon, which orbits at a mean altitude of 22,660 kilometers. The brochure noted that the leading Voyager main engine candidate was a modified Apollo Lunar Module descent engine. The complete Voyager Orbiter propulsion system fully loaded with propellants would weigh 6.5 tons.

After orbit insertion, the Orbiter's instruments would be turned toward Mars to image candidate sites for the first Voyager landing. After scientists and engineers on Earth settled on a site, the 2.5-ton Voyager landing capsule would eject its sterilization canister, separate from the Orbiter beyond Deimos, and fire a 188.2-kilogram solid-propellant deorbit rocket to change its path so that at periapsis it would intersect the martian atmosphere. The deorbit rocket would then detach.

The Voyager landing capsule would enter the martian atmosphere moving at between two and three miles per second. Aerodynamic braking using the 6.1-meter-diameter conical heat shield would cut speed to between 122 and 305 meters per second by the time the capsule fell to within 4570 meters of the surface. The heat shield would eject, then the Lander would fire its descent engines and deploy a supplemental parachute.

Voyager hardware heritage. Note the Voyager capsule and Lander configurations. Image credit: NASA.

During descent, the Lander would image the surface and collect atmospheric data. It would release the parachute, then slow to a hover three meters above Mars. Its descent engines would then shut off, allowing it to drop to a gentle touchdown on three legs.

The 1973 Lander would include 136.1 kilograms of science equipment. Over several days, it would search for water and life, measure cosmic and solar radiation, and study the atmosphere — it would, for example, measure the quantity of dust in the martian air.

The 1973 Orbiter, for its part, would include 181.4 kilograms of scientific instrumentation, which it would use to map Mars in detail and search for surface changes over time, determine surface composition, and measure solar and cosmic radiation. The Orbiter would also act as a martian weather satellite. It would, the brochure explained, use its main engine to change the altitude and inclination of its orbit several times during its two-year operational lifetime, allowing detailed study of much of Mars.

Congress refused to fund Voyager in FY 1968, in part because it had come to be seen as a lead-in to a costly post-Apollo piloted Mars/Venus flyby program, and also because the Apollo 1 fire (27 January 1967) undermined confidence in NASA. The U.S. civilian space agency formally abandoned its Voyager plans in September 1967.

In 1968, however, Congress agreed to fund the Viking program in FY 1969. Like Voyager, Viking would emphasize the search for life and would use twin spacecraft, each including a lander and an orbiter. Unlike its ill-starred progenitor, however, Viking made no claim to be a precursor for a piloted Mars mission. In addition, Viking would be managed by NASA's Langley Research Center, not JPL, though the later would build the Viking orbiters. Many interpreted assignment of Viking management to Langley as a congressional rebuke to JPL for its independent mindset; efforts to preserve NASA centers as Apollo spending began to wind down probably also played a role.

Twin flyby Mariners 6 and 7 flew by Mars in 1969, and Mariner 9 orbited the planet in 1971-1972. After skipping the 1973 Mars launch opportunity, NASA launched Viking 1 on a Titan-IIIE rocket with a Centaur upper stage on 20 August 1975. Viking 1's Mariner-based, solar-powered orbiter and RTG-powered lander together weighed about 2.56 tons at launch. After deploying the lander in Mars orbit, the Viking 1 orbiter weighed about 898.1 kilograms.

The Viking 1 lander became the first spacecraft to land successfully on Mars on 20 July 1976, seven years to the day after Apollo 11 became the first manned lunar lander. The lander had a mass of about 598.7 kilograms after touchdown; of this, about 42.2 kilograms comprised scientific instrumentation. Viking 2 launched from Earth on 9 September 1975, and its lander touched down on 3 September 1976. At about three meters wide, the Viking landers were about half the size of the planned Voyager landers.

NASA and JPL recycled the Voyager name in 1977, applying it to twin Mariner-derived Jupiter-Saturn flyby spacecraft (the mission was originally called Mariner Jupiter-Saturn 77). Voyager 2 left Earth first, on 20 August 1977, atop a Titan III-E/Centaur. Voyager 1 launched 16 days later, on 5 September. Voyager 1 passed Voyager 2 on 19 December 1977, as the twin spacecraft traversed the Asteroid Belt between Mars and Jupiter.

Beyond Mars: artist concept of the flight path of a Mariner Jupiter-Saturn 77 spacecraft. Image credit: NASA.

At each planet, the Voyagers performed a gravity-assist maneuver; that is, they used the target planet's gravity and orbital momentum to speed themselves onward to their next destination. Voyager 1 flew by Jupiter at a distance of 349,000 kilometers on 5 March 1979; Voyager 2 followed on 9 July 1979, passing the giant planet at a distance of 570,000 kilometers. Voyager 1 then flew by Saturn, its last planned target, at a distance of 124,000 kilometers on 12 November 1980. Voyager 2 passed Saturn at a distance of 101,000 kilometers on 25 August 1981.

By that time, the decision had been made to add Uranus and Neptune to Voyager 2's list of targets. The intrepid spacecraft flew by the former at a distance of 81,500 kilometers on 24 January 1986, and passed the latter at a distance of just 4951 kilometers on 25 August 1989.

The Voyagers continue to transmit data on space conditions beyond the planets. At this writing, Voyager 1 is 144.7 Astronomical Units (AU) from the Sun (one AU is defined as the mean distance from Earth to the Sun, or about 149.6 million kilometers - for comparison, the most distant planet in the Solar System, Neptune, is on average 30.1 AU from the Sun). Radio signals traveling at the speed of light (299,792 kilometers per second) need more than 20 hours to reach it. Voyager 2, which dove below the plane of the Solar System after departing Neptune, is 119.8 AU from the Sun; radio signals need about 16 hours, 42 minutes to reach it.

Voyager 1 became the first spacecraft to pass beyond the heliosphere, the bubble of space where solar particles and fields are dominant, in August 2012. Voyager 2 joined it at the edge of interstellar space in November 2018.

Sources

Summary of the Voyager Program, NASA Office of Space Science and Applications, January 1967.

On Mars: Exploration of the Red Planet 1958-1978, NASA SP-4212, Edward Clinton Ezell and Linda Neumann Ezell, NASA, 1984, pp. 85-86, 101-103, 117-118.

Voyager: Mission Status (https://voyager.jpl.nasa.gov/mission/status/#where_are_they_now - accessed 4 February 2019).

Saturn IB/Centaur (http://heroicrelics.org/info/saturn-i-and-ib/saturn-ib-centaur.html - accessed 6 February 2019).

More Information

Centaurs, Soviets, and Seltzer Seas: Mariner 2's Venusian Adventure (1962)

Things To Do During a Venus/Mars/Venus Piloted Flyby Mission (1968)

Lunar Viking (1970)

The Challenge of the Planets, Part Three: Gravity

Purple Pigeon: Mars Multi-Rover Mission (1977)

Image credit: JPL/NASA.

Planetary scientist Bruce Murray became director of the Jet Propulsion Laboratory (JPL) in April 1976, just three months before Viking 1 was due to land on the northern plains of Mars. Though NASA's Langley Research Center managed Project Viking, JPL included Viking Mission Control. When Viking 1 landed, JPL could expect to play host to hundreds of journalists from all over the Earth.

According to his 1989 memoir Journey into Space: The First Thirty Years of Space Exploration, Murray saw this as an opportunity. He quickly assembled a group of six engineers to propose planetary missions that he could pitch to the journalists and, through them, to U.S. taxpayers.

The missions, which Murray dubbed "Purple Pigeons," were intended to include both "high science content" and "excitement and drama [that would] garner public support." They were called Purple Pigeons to differentiate them from "Gray Mice," unexciting and timid missions which Murray felt would help to ensure that JPL had no future in the space exploration business. By August 1976, the Purple Pigeons flock included a solar sail mission to Halley's Comet, a Mars Surface Sample Return (MSSR), a Venus radar mapper, a Saturn/Titan orbiter/lander, a Ganymede lander, an asteroid tour, and an automated lunar base.

Bruce Murray, JPL director from April 1976 until June 1982. Image creditI JPL/Caltech.

The Purple Pigeons effort continued even after Viking 2 landed (3 September 1976) and all the journalists went home. In a February 1977 JPL report, for example, JPL engineers described a Purple Pigeon mission that would explore Mars with up to four rovers simultaneously. The Viking-based multi-rover mission would include a pair of identical 4800-kilogram spacecraft, each consisting of a Viking-type orbiter and a 1578-kilogram Mars lander bearing twin 222.4-kilogram rovers. The rovers would, the report stated, perform traverses to "regions difficult to reach by direct landings." This would, it added, fill the gap between "detailed information" from MSSR missions and "global information" from Mars orbiters.

The image at the top of this post shows a somewhat different (probably later) multi-rover mission design. Its four six-wheel, multi-cab rovers (two of which are operating out of view over the horizon) rely on a single Viking orbiter-type spacecraft to relay radio signals to and from Earth. In principle, however, it is identical to the early multi-rover mission design described in this post.

Most MSSR plans of the 1970s assumed a "grab" sample; that is, that the stationary MSSR lander would return to Earth a sample of whatever rocks and soil happened to be within reach of its robotic sample scoop. The report suggested that the rovers of the multi-rover mission might enhance a follow-on MSSR mission by collecting and storing samples as they roved across the planet. After the MSSR lander arrived on Mars, the rovers would rendezvous with it and hand over their samples for return to Earth. The report contended that its multi-rover/MSSR strategy would be "an enormous advance over even multiple grab samples" collected by MSSR landers at widely scattered sites.

At the time the Purple Pigeons team proposed the multi-rover mission, NASA intended to launch all payloads, including interplanetary spacecraft, on board reusable Space Shuttles. The Shuttle orbiter would be able to climb no higher than about 500 kilometers, so launching payloads to higher Earth orbits or interplanetary destinations would demand an upper stage. The powerful liquid-propellant Centaur upper stage would not be ready in time for the opening of the Mars multi-rover launch window, which spanned from 11 December 1983 to 20 January 1984, so JPL tapped a three-stage solid-propellant Interim Upper Stage (IUS) to push its Purple Pigeon out of Earth orbit toward Mars.

After an Earth-Mars cruise lasting about nine months, the twin multi-rover spacecraft would arrive at Mars a week or two apart between 16 September and 27 October 1984. They would each fire their main engines to slow down so that Mars gravity could capture them into an elliptical orbit with a periapsis (low point) of 500 kilometers, a five-day period, and an inclination of 35° relative to the martian equator.

The multi-rover landers would then separate and each fire a solid-propellant de-orbit rocket at the apoapsis (high point) of its orbit to begin descent to the surface. Landing sites between 50° north latitude and the south pole would in theory be accessible, though the need for a direct Earth-to-rover radio link would in practice prevent landings below 55° south.

The landers would each be encased within an aeroshell with a heat shield for protection during the fiery descent through the martian atmosphere. The aeroshell would have the same 3.5-meter diameter as its Viking predecessor, though its afterbody would be modified to make room for the large cooling vanes of the twin rovers' electricity-producing Radioisotope Thermal Generators (RTGs).

JPL's dual rovers packed inside their modified Viking-type aeroshell. Image credit: JPL.

After the landers touched down, the orbiters would maneuver to areosynchronous orbit. In such an orbit, 17,058 kilometers above the martian equator, only minor orbital corrections would enable a spacecraft to "hover" indefinitely over one spot on the equator. Each orbiter would position itself over a spot on the equator near its lander's longitude so that it could relay radio signals between its rovers on Mars and operators on Earth.

The multi-rover lander, which would serve no purpose beyond rover delivery, would constitute a radical departure from the triangular Viking lander design, though it would use Viking technology where possible to save development costs. It would comprise a rectangular frame to which would be attached three uprated Viking-type terminal descent engines, two spherical propellant tanks, and three beefed-up Viking-type landing legs.

Multi-rover lander. Image credit: JPL.

The 1.5-meter-long rovers would be mounted on the lander frame with their four 0.5-meter-diameter wire wheels compressed. Releasing a latching mechanism would permit the wheels to expand, raising the rover off four stabilizing "taper pins." The pins and one terminal descent engine would then swing out of the way, ramps would deploy, and the first rover would roll onto the rocky martian surface. The second rover would then ride a motor-driven "dolly" to the first rover's initial position before unlatching and joining its twin on the ground.

JPL envisioned that its four-wheeled rovers would each deploy a one-meter-tall boom holding a still-image camera, a floodlight, a strobe light, a weather station, and a pointable horn-shaped radio antenna. The camera/antenna boom, the tallest part of the rover, would stand about two meters above the surface. Controllers on Earth would then put the rovers through an initial checkout lasting at least two weeks. The checkout would culminate in slow "manual" (Earth-controlled) and faster semi-autonomous (Earth-directed but rover-controlled) traverses.

JPL's nuclear-powered rover viewed from above (top) and from the side. Image credit: JPL.

In semi-autonomous mode, operators would plan traverse routes and science targets using stereo images from the rover camera taken from terrain "high points," then would command the rover to proceed. The rovers might assist each other in traverse planning; for example, "high point" pictures from one might fill in blind spots in the other's field of view. "After the first few kilometers of traverse," the JPL engineers assumed, operators on Earth would "begin to build an intuitive feeling for the Martian geography and its impact on the rover capabilities, allowing them to plan better paths." The rovers would also photograph each other to enhance the mission's "general public appeal."

The rover mobility system would include one electric drive motor per wheel, eight proximity sensors for obstacle detection, inclinometers to monitor rover tilt, motor temperature sensors to judge wheel traction, a gyrocompass/odometer, a laser rangefinder with a 30-meter range, and an "8-bit word, 16k active, 64k bulk, floating point arithmetic and 16-bit accuracy" computer. The JPL engineers judged that their rovers would be capable of moving at up to 50 meters per hour over terrain similar to that seen at the Viking 1 landing site.

Dusk at the Viking 1 landing site in Chryse Planitia. Image credit: NASA.

Alpha-scattering X-ray fluorescence and gamma-ray spectrometers would collect data while the rovers were in motion, but all other science, including imaging and sample collection, would occur only while they were parked. Each rover would gather samples using an "articulated arm" with an "electromechanical hand."

In order to avoid "an overabundance of data from a single track," the rovers would travel slightly different routes and rendezvous at the end of each leg of their traverse. They would, however, travel close enough together that each could aid the other in the event of trouble. If one rover became stuck in loose dirt, for example, its companion could use its articulated arm to place rocks under its wheels to improve traction. If one rover of a pair failed, the report maintained, the other would continue to yield "good, solid science."

The rovers would be designed to operate for at least one martian year (about two Earth years) to help ensure that at least one of the four could successfully rendezvous with the follow-on MSSR mission, which would leave Earth in 1986. Estimates of rover traverse distances in 1970s and 1980s studies were typically highly optimistic, and the multi-rover mission was no exception: each of the mission's four rovers was expected to travel up to 1000 kilometers. The JPL engineers concluded their report by calling for new technology development to ensure that adequate power and mobility systems would become available by the time their Purple Pigeon was due to fly.

Sources

Journey into Space: The First Thirty Years of Space Exploration, Bruce Murray, W. W. Norton & Co., 1989.

Feasibility of a Mars Multi-Rover Mission, JPL 760-160, Jet Propulsion Laboratory, 28 February 1977.

More Information

Triple-Flyby: Venus-Mars-Venus Piloted Missions in the Late 1970s/Early 1980s (1967)

Prelude to Mars Sample Return: The Mars 1984 Mission (1977)

Making Propellants from Martian Air (1978)

Exploring Mars from Pole to Pole: MESUR Network (1991)

Pioneer Venus 2 releases its three small Venus atmosphere entry probes. Through artist license, the large probe is visible against the clouds of Venus; it would not in fact have been visible at the time the small probes were released. Image credit: NASA.

On 8 August 1978, NASA launched Pioneer Venus 2 (PV2) on an Atlas-Centaur rocket. The 904-kilogram spacecraft, known also as Pioneer Venus Multiprobe, released a 1.5-meter-diameter battery-powered atmosphere entry probe on 16 November and three 76-centimeter-diameter probes on 20 November.

On 9 December 1978, the five parts of PV2 entered the thick, hot Venusian atmosphere. The drum-shaped probe carrier burned up as planned at an altitude of 110 kilometers. Sturdy conical heat shields protected the spherical instrumented probes from aerodynamic heating. As drag slowed it, the large probe deployed a parachute.

Two of the small probes, which did not include parachutes, exceeded all expectations by surviving landing and transmitting data from the hellish Venusian surface. One, the Day Probe, transmitted for 67.5 minutes before succumbing to heat, pressure, and battery failure, setting a new world record for spacecraft endurance on Venus.

PV2 was the last U.S. planetary mission launched until 1989. NASA Ames Research Center (ARC), located near San Francisco, California, managed PV2 and its sister spacecraft, PV1 (the Pioneer Venus Orbiter).

In July 1991, ARC proposed a multiprobe system outwardly not too different from PV2, but intended to create a long-lived network of low-cost science stations on Mars. According to ARC's report on the concept, its network would reflect a design philosophy with "unique characteristics . . . derived from the Pioneer Project corporate memory."

Mars networks were first proposed in the early 1970s. Scientific advisory groups endorsed the network concept repeatedly in the following two decades as the best way to obtain global-scale weather and seismic data. In the late 1980s, at the behest of the NASA Headquarters Solar System Exploration Division (SSED), the Jet Propulsion Laboratory (JPL) Precursor Task Team included a network in its program of precursor robotic missions for paving the way for astronauts on Mars. In common with previous Mars network plans, the 1989 plan invoked spear-shaped penetrators to hard-land stations at low cost.

NASA ARC's Mars Environment Survey (MESUR - pronounced "measure"), on the other hand, invoked cheap rough-landing landers, or "stations," that would deploy protective airbags seconds before landing. MESUR would build up a "pole-to-pole" network of 16 stations during the 1999, 2001, and 2003 minimum-energy Mars launch opportunities.

Each 158.5-kilogram MESUR lander would leave Earth attached toa Mars atmosphere entry deceleration system and a simple cruise stage. Upon arrival at Mars, each would cast off its cruise stage and enter the atmosphere directly from its Earth-Mars trajectory at up to seven kilometers per second. The ARC report compared this with the Viking landers, which entered from Mars orbit at only 4.4 kilometers per second. The lander's heat shield, a two-meter-diameter flattened cone, would be designed to withstand atmosphere entry during planet-wide dust storms, when suspended dust particles might exacerbate shield erosion.

Partial cutaway of a MESUR station on the surface of Mars. Image credit: NASA Ames Research Center.

The ARC report acknowledged that the disk-shaped lander might bounce to rest on Mars in either "heads" or "tails" orientation, but rejected as costly and risky a mechanical system for tipping it upright. The ARC engineers opted instead for circular ports that would enable controllers to deploy instruments from either side of the station. Instruments might include imagers, an atmospheric structure experiment, gas analyzers, a weather station, a spectrometer, and a seismometer.

The report explained that solar cells were initially ARC's preferred MESUR power system, but analysis had shown that the number of cells that could be mounted on the lander's small surface would not generate enough electricity to drive its science instruments unless landings were limited to sites within 30° of the martian equator. This limitation was deemed unacceptable by the MESUR Science Definition Team, so engineers opted for a small (nine-kilogram) General Purpose Heat Source (GPHS) Radioisotope Thermal Generator (RTG) "brick" based on Ulysses solar polar orbiter/Galileo Jupiter orbiter RTG technology.

Sixteen MESUR landers would need 16 GPHS bricks over six years. The report noted that the entire MESUR Network would need less than half as much plutonium as the Cassini Saturn orbiter, which would carry two RTGs with 18 GPHS bricks each.

Cutaway of the MESUR Network launch shroud showing four MESUR landers (one is mostly obscured behind the lander support structure) and the solid-propellant Mars transfer orbit injection stage. Image: NASA Ames Research Center.

The MESUR mission would begin in 1999 with the launch of a single Delta II 7925 rocket from Cape Canaveral, Florida, with four MESUR landers mounted on a framework within its 9.5-foot-diameter streamlined launch shroud. After a solid-propellant upper stage placed them on course for Mars, the landers would separate from the framework to travel on "independent free-flyer trajectories" that would permit precise Mars landing site targeting. Three side-mounted landers would tumble after separation, but sloshing propellants in their cruise stages would gradually damp their gyrations.

The landers would discard their cruise stages 125 kilometers above Mars. Ten kilometers above the planet, each would deploy a pilot parachute, then cast off its heat shield and open its single main parachute. The landers would image the surface and collect atmospheric structure data during the final eight kilometers of descent.

Just two meters above the landing site, each lander would release its main parachute and inflate its airbags. A small rocket on the parachute would ignite to prevent it from settling over the lander.

The MESUR lander design would permit landings at sites up to six kilometers above the base datum, the martian equivalent of Earth's sea level. The base datum, referenced to the minimum Mars atmospheric pressure required for liquid water to exist on the surface, was established after Mariner 9 mapped the planet from orbit in 1971-1972. (In 2001, a new system referenced to the mean radius of Mars as measured by Mars Global Surveyor's MOLA instrument replaced the base datum.)

Though all 16 MESUR landers would carry the same suite of instruments, their individual landing sites would be selected to cater to different science requirements. The report advised that weather stations should be spaced widely over the planet, while seismic stations should form closely spaced "triads." These conflicting requirements forced a "compromise network design."

MESUR Network Stations 1 and 2 would land near each other on the north rim of Valles Marineris to form a "seismic pair." Station 3, at the foot of Olympus Mons in Tharsis, would also emphasize seismic research. Station 4 would aim to extend the weather record for Chryse Planitia, where Viking 1 accumulated data from 1976 to 1983.

The Tharsis hemisphere of Mars showing proposed positions of MESUR stations. See text for explanation. Image credit: NASA.

In 2001, two Delta II 7925s would launch 20 days apart bearing four more MESUR landers and a communications relay orbiter, respectively. The latter payload, based on an existing Earth-orbital comsat design, would serve as radio relay for the expanding network, enabling MESUR stations to return data from sites all over the martian surface.

It would reach Mars in 10 months on a slow "Type II" trajectory to reduce the amount of propellant it would need to slow down so that the planet's gravity could capture it. Launch of the communications orbiter would be delayed until 2001 in order to spread its cost over a longer period.

With the successful arrival of the four 2001 stations, a "minimal network" would be in place on Mars. Station 5, on the Marineris north rim, would create a "seismic triad" with Stations 1 and 2, while Station 6, northwest of Olympus Mons, would create a seismic pair with Station 3. Station 7, east of Solis Planum ("a region of known dust storm activity"), and Station 8, in western Acidalia Planum, would expand martian meteorological coverage.

The final two MESUR Delta II 7925 launches in 2003 would boost four landers each on course for Mars. Stations 9 and 10 would be located near the north and south poles, respectively, while Station 11 would report weather conditions in Aonia Terra, southwest of the great Argyre basin. Stations 12 (northwest Hellas), 13 (Elysium Planitia), and 14 (Deuteronilus Mensae) would further extend martian meteorological coverage.

Station 15 (Sirenum Terra) would form a Tharsis seismic triad with Stations 3 and 6. Station 16, in Syrtis Major on the side of Mars opposite Olympus Mons, would create a seismic pair with Station 13 and, with the Tharsis triad, enable the size of Mars's core to be determined.

The Syrtis Major hemisphere of Mars showing proposed positions of MESUR stations. See text for explanation. Image credit: NASA.

The entire 16-station network and its communications orbiter would function for at least a martian year (a little more than two Earth years). This would mean that the 1999 stations would have to endure for three martian years (6.5 Earth years), while the 2001 stations and communications orbiter would need to function for two martian years (4.3 Earth years).

In its 1991 strategic plan, published the same month as ARC's MESUR report, the SSED dubbed MESUR its "baseline plan" for a Mars network mission. In November 1991, NASA elected to move MESUR Phase A development to JPL, where the project was split into two parts.

MESUR Network would be preceded by MESUR Pathfinder, a single-spacecraft mission for technology testing. Pathfinder was built larger than the the planned MESUR landers so that it could deliver to Mars a six-wheeled "microrover." JPL also opted for solar power in place of NASA ARC's RTG bricks and a petal system to permit it to flip itself upright and release the rover instead of small instrument deployment ports.

In 1994, in the wake of the Mars Observer failure, NASA funded the Mars Surveyor Program in place of MESUR Network. Work continued on Pathfinder under the auspices of NASA's low-cost Discovery Program, however, and it landed successfully on Mars on 4 July 1997.

Mars Pathfinder Lander (background) and Sojourner rover. Image credit: NASA.

Sources

Mars Environmental Survey (MESUR) Science Objectives and Mission Description, NASA Ames Research Center, 19 July 1991.

Solar System Exploration Division Strategic Plan: Preparing the Way to the New Frontier of the 21st Century, Special Studies Office, Space Telescope Science Institute, July 1991.

More Information

Pioneer Mars Orbiter with Penetrators (1974)

Prelude to Mars Sample Return: The Mars 1984 Mission (1977)

Near-Term and Long-Term Goals: Space Station and Lunar Base (1983-1984)

The Space Operations Center (SOC), a Shuttle-launched multimodular station concept NASA Johnson Space Center and its contractors studied starting in 1978, was an oft-cited exemplar of NASA space station thinking in the early 1980s, when Science Applications Incorporated performed its study for the National Science Foundation. Image credit: NASA.

In December 1983, the Division of Policy Research and Analysis of the National Science Foundation enlisted Science Applications Incorporated (SAI) of McLean, Virginia, to compare the science and technology research potential of an Earth-orbiting space station with that of a base on the Moon. In its report, which was completed on 10 January 1984, SAI cautioned that, because its study was performed "in a very short two-week period," it could offer only "a preliminary indication" of the relative merits of a space station in low-Earth orbit (LEO) and a lunar base. Though SAI did not say so, its study had a short turnaround because its results were meant to inform the White House ahead of President Ronald Reagan's planned announcement of a NASA space station program during his 25 January 1984 State of the Union Address.

SAI explained that its study used a four-step approach. First, the study team judged which science and technology disciplines could best be served by an LEO space station and which by a lunar base. Next, the team developed a lunar base conceptual design capable of serving the disciplines it identified. It then developed a transportation system concept for deploying and maintaining its base. Finally, the team estimated the cost of its lunar base.

The team identified five science and technology disciplines that would be better served by a base on the Moon than by a space station. The first was radio astronomy. Bowl-shaped radio telescopes might be built in bowl-shaped lunar craters, SAI wrote. Radio astronomers might take advantage of the Moon's Farside (the hemisphere turned permanently away from Earth), where up to 2160 miles of rock would shield their instruments from terrestrial radio interference. The 238,000-mile separation between lunar and terrestrial radio telescopes would permit Very Long Baseline Interferometry observations, enabling astronomers to map minute details of galaxies far beyond the Milky Way.

A bowl-shaped crater makes an ideal site for a bowl-shaped radio telescope. Visible stars are artist's license; the harsh glare of the Sun in lunar daylight would banish them from view. Image credit: NASA.

High-energy astrophysics and physics was SAI's second lunar base discipline. The team noted that, because the Moon offers "a large, flat area, a free vacuum, and a local source of refined material for magnets," it might become an economical site for a large particle accelerator.

Lunar geology (which SAI called "selenology") would obviously be better served by a lunar base than by a space station. SAI noted that, despite 13 successful U.S. robotic lunar missions and six successful Apollo landings, the Moon had "barely been sampled and explored." Lunar base selenological exploration would focus on "understanding better the early history and internal structure of the Moon" and "exploring for possible ore and volatile deposits." Selenologists would rove far afield from the base to measure heat flow and magnetic properties, drill deep into the surface, deploy seismographs, and collect and analyze rock samples.

SAI's fourth lunar discipline was resource utilization. The study team noted that samples returned to Earth by the Apollo astronauts contain 40% oxygen by weight, along with silicon, titanium, and other useful chemical elements. Lunar oxygen could be used as oxidizer for chemical-propulsion spacecraft traveling between Earth and Moon and from LEO to geosynchronous Earth orbit (GEO). Silicon could be used to make solar cells. (SAI pointed out, however, that the two-week lunar night would make reliance on solar arrays for electricity "somewhat difficult.") Raw lunar dirt — known as regolith — could serve as radiation shielding. If water ice were found at the lunar poles — perhaps by the automated lunar polar orbiter SAI advised should precede the lunar base program — then the Moon might supply hydrogen rocket fuel as well as oxidizer.

SAI's fifth and final lunar base science discipline was systems development. The team expected that lunar base technology development would be "devoted to improving the efficiency and capabilities of systems that support the base," such as life support, with the goal of "reduced reliance on supplies sent from Earth." Transport system development might include research aimed at developing a linear electromagnetic launcher of the kind first proposed by Arthur C. Clarke in 1950. Such a device — often called a "mass driver" or "rail gun" — might eventually launch bulk cargoes (for example, lunar regolith, liquid oxygen propellant, and refined ores) to sites all around the Earth-Moon system.

The SAI team noted that some disciplines might be served equally well by a lunar base or an Earth-orbiting space station. Large (100-meter) telescopes for optical astronomy, for example, might be equally effective on the Moon or in Earth orbit. The Moon, however, would offer a solid surface that might enable the "pointing stability and optical system coherence" such a telescope would need to perform adequately.

SAI acknowledged that its report proposed "research and development activities. . .too numerous and often too difficult for a first-generation lunar base." It thus divided activities within the five lunar base disciplines into two categories: those suitable for its first-generation base and those that would need a more elaborate second-generation facility. First-generation radio astronomy, for example, would use two small dish antennas on Nearside (the lunar hemisphere always facing Earth). In the second generation, a 100-meter-diameter antenna would operate on Farside.

Having defined its lunar base science program, the SAI team moved on to the second and third steps in its study. The team assumed that NASA's Space Shuttle, which at the time they wrote had just completed its ninth flight (STS-9/Spacelab 1, 28 November-8 December 1983), would form part of the lunar base transportation infrastructure, along with an LEO space station. The Shuttle would cheaply and reliably deliver lunar base crews, spacecraft, and cargo to the station, where they would be brought together for flight to the Moon. SAI proposed reapplying hardware developed for the LEO station — for example, pressurized modules — to the lunar base program.

An October 1984 paper by study participants Steve Hoffman and John Niehoff for the first Lunar Bases and Space Activities of the 21st Century symposium provided additional details of SAI's Earth-Moon transportation system and surface base design. Where details in the October 1984 paper conflict with those in the December 1983 report, the description that follows defaults to information contained only in the latter (mostly).

Homeward bound: an Orbital Transfer Vehicle (OTV) bearing a returning lunar base crew aerobrakes in Earth's atmosphere. After aerobraking it will rendezvous with NASA's space station. Image credit: Pat Rawlings/NASA.

SAI's lunar transportation system would include three types of spacecraft. The first, the reusable Orbital Transfer Vehicle (OTV), would be a two-stage vehicle permanently based at the LEO station. SAI assumed that NASA would develop OTVs for moving cargoes between the LEO station and higher orbits (for example, GEO) and that the basic OTV design would then be modified for lunar base use. The OTV, which would operate as a piloted spacecraft through addition of a pressurized "personnel pod," would deliver up to 16,950 kilograms of crew and cargo to lunar orbit.

An OTV-derived four-legged lunar lander would form the basis of two vehicles: the Logistics Lander and the Lunar Excursion Module (LEM). The former would include a removable subsystem module for automated lunar landings and the latter would carry a personnel pod for piloted flight. These were listed as the second and third spacecraft in SAI's lunar transportation system, though one might argue that they were actually tricked-up OTVs.

SAI's one-way cargo lunar flight mode. Please click to enlarge. Image credit: Science Applications Incorporated.

The three vehicle types would support two basic lunar flight modes. One-way cargo missions would use Direct Descent. The OTV first stage would ignite and burn nearly all of its propellants, then would separate, turn around, and fire its engines to slow down and return to the LEO station for refurbishment. The OTV second stage would then ignite, burn most of its propellants, and separate from the Logistics Lander. The second stage would swing around the Moon on a free-return trajectory, fall back to Earth, aerobrake in Earth's atmosphere, and rendezvous with the LEO station. The Logistics Lander, meanwhile, would descend directly to the lunar base site with no stop in lunar orbit.

For two-way crew sorties, the OTV first stage would operate as during a one-way cargo mission. After a three-day flight, the OTV second stage/personnel pod combination would ignite its engines to slow itself so the Moon's gravity could capture it into lunar orbit. There it would dock with a waiting LEM carrying lunar base astronauts bound for Earth, who would trade places with the new base crew. In addition to the new crew, 12,750 kilograms of propellants (sufficient for a round trip from lunar orbit to the surface base and back again) and up to 2000 kilograms of cargo would be transferred from the OTV second stage/personnel pod to the LEM.

SAI's roundtrip crew rotation lunar flight mode. Please click to enlarge. Image credit: Science Applications Incorporated.

The OTV second stage/personnel pod and the LEM would then separate. The former would fire its engines to depart lunar orbit for Earth, and the latter would descend to a landing at the lunar base. The OTV second stage/personnel pod combination would subsequently aerobrake in Earth's atmosphere and return to the LEO station for refurbishment.

SAI's base buildup sequence would begin with a pair of Site Survey Mission flights. The first would see an unpiloted LEM with empty propellant tanks placed into lunar orbit through a variant of the crew sortie mode. An automated OTV second stage bearing the LEM in place of a personnel pod would enter lunar orbit, undock from the LEM, and return to Earth.

The second Site Survey Mission flight would employ another variant of the Crew Sortie mode. Five astronauts would arrive in lunar orbit on board an OTV second stage/personnel pod and dock with the waiting LEM. The four astronauts of the base site survey team would transfer to the LEM along with propellants and supplies. They would then undock and land at the proposed base site, leaving the OTV pilot alone in lunar orbit. After completing their survey of the site, they would return to the OTV second stage/personnel pod, then would undock from the LEM and return to Earth orbit.

Assuming that the base site checked out as acceptable, Flight 3 would see the start of base deployment. A Logistics Lander would employ Direct Descent mode to deliver to the base site an Interface Module and a Power Plant. The Interface Module, which would be based on LEO space station hardware, would include a cylindrical airlock, a top-mounted observation bubble, and a cylindrical tunnel with ports for attaching other base modules. SAI's proposed Power Plant was a nuclear source capable of generating 100 kilowatts of electricity.

Flight 4 would deliver two "mass mover" rovers, two 2000-kilogram mobile laboratory trailers, and a 1000-kilogram lunar resource utilization pilot plant. The rovers would tow the mobile labs up to 200 kilometers from the base on selenologic excursions lasting up to five days. The mobile labs would carry instruments for microscopic imaging, elemental and mineral analysis, and subsurface ice detection, stereo cameras, and a soil auger or core tube for drilling up to two meters deep. The first-generation lunar resource utilization pilot plant would process 10,000 kilograms of regolith per year to yield oxygen, silicon, iron, aluminum, titanium, magnesium, and calcium.

Flight 5 would deliver the Laboratory Module, the first 14-foot-diameter, 40-foot-long cylindrical base module based on the pressurized module design used to build the LEO station. Flight 6 would deliver the Habitat Module, which would provide living quarters for the seven-person base crew, and Flight 7 would deliver the Resources Module, which would include a pressurized control center and an unpressurized section containing water and oxygen tanks and equipment for life support, power conditioning, and thermal control. The final base deployment flight, a duplicate of Flight 1, would deliver a backup LEM to lunar orbit.

Long-term occupation of the Moon would begin with Flight 9, a crew sortie mission that would deliver a four-person construction team. Flight 10 would see three more astronauts join the construction team, bringing the total base population to seven. The OTV pilots for these flights would return to Earth alone after the construction teams undocked and landed at the base in their respective LEMs.

Using the mass mover rovers, the base crew would unload the Logistics Landers and join together the base components. The completed base would provide seven astronauts with 2000 cubic feet of living space per person. They would attach the Lab, Hab, and Resource Modules to the Interface Module, then would link the resource utilization pilot plant to the Lab Module.

The Power Plant would be placed a safe distance away from the base and linked by a cable to the base power conditioning system. The crew would then use hoses to link the Power Plant and base thermal control system to a heat exchanger/heat sink. Finally, after Power Plant activation, the astronauts would use bulldozer scoops on the rovers to cover the pressurized modules with regolith radiation shielding.

Flight 11, the first base crew rotation flight, would see the four-person construction team that arrived on Flight 9 lift off in a LEM and return to lunar orbit, where they would dock with an OTV second stage/personnel pod combination just arrived from Earth. The Flight 9 lunar base team would trade places with them and, following LEM refueling and cargo loading, would descend to a landing at the base. The first construction team and the Flight 11 OTV pilot would then return to the LEO station. On Flight 12, a three-person base team would replace the Flight 10 team.

Lunar base teams of three or four astronauts would rotate every two months. The typical base complement would include a commander/LEM pilot, a LEM pilot/mechanic, a technician/mechanic, a doctor/scientist, a geologist, a chemist, and a biologist/doctor.

Mass mover rover in the field with advanced power cart and deep drill rig. Image credit: NASA.

SAI then estimated the cost of its lunar base and three years of operations based on NASA's cost estimates for the Space Shuttle and the LEO station. At the time SAI conducted its study, NASA placed the cost of its proposed LEO station at between $8 billion and $12 billion. This was in fact an underestimation calculated to make the station more politically palatable to the White House and Congress. NASA placed the total cost of LEO station Logistics, Habitat, Laboratory, and Resource Modules and other structures at $7.1 billion, so SAI estimated the total cost of the lunar base Resource, Habitat, Laboratory, and Interface Modules at $5.8 billion.

Although the OTV would find uses in LEO and GEO, SAI charged all of its development and procurement costs (a total of $7.2 billion) to the lunar base. The expendable Logistics Lander and reusable LEM would cost $6.6 billion and $4.8 billion, respectively. The LEM, though structurally beefier and more complex, would cost less because the Logistics Lander would bear the development cost of systems common to both landers.

Based on optimistic NASA pricing, the SAI team assumed that a Shuttle flight would cost $110 million in 1990. The 89 Shuttle flights in the lunar base program would thus cost a total of $9.8 billion. The LEO station, by contrast, would need only 17 Shuttle flights at a cost of $1.9 billion. SAI placed total LEO station cost plus three years of operations at $14.2 billion. Lunar base cost plus three years of operations came to $54.8 billion.

To conclude its report, SAI noted that both the LEO station and the lunar base could be completed in about a decade. The LEO station would, however, serve a broader science user community and would provide an OTV base in LEO for eventual lunar base use. The SAI team argued that the LEO station was a reasonable near-term (10-year) objective, while the lunar base would yield obvious benefits in a long-term (50 years) space program. It added that the

Space Program will function best if it has both near-term objectives and long-range goals. The near-term objectives assure [sic] that we progress with each year that passes. The long-range goals provide direction for our annual progress. The Space Station and Lunar Base appear to serve these respective roles at the present time.

Sources

A Manned Lunar Science Base: An Alternative to Space Station Science? A Brief Comparative Assessment, Report No. SAI-84/1502, Science Applications, Inc., 10 January 1984.

"Preliminary Design of a Permanently Manned Lunar Surface Research Base," S. Hoffman and J. Niehoff, Science Applications International Corporation; published in Lunar Bases and Space Activities of the 21st Century, "papers from a NASA sponsored, public symposium hosted by the National Academy of Sciences in Washington, D.C., Oct[ober] 29-31, 1984," W. W. Mendell, editor, Lunar and Planetary Institute, 1985, pp. 69-75.

More Information

Chronology: Space Station 1.0

As Gemini Was to an Apollo Lunar Landing by 1970, So Apollo Would Be to a Permanent Lunar Base by 1980 (1968)

"A Vision of the Future": Military Uses of the Moon and Asteroids (1983)

Another Look at Staged Reentry: Janus (1962-1966)

The M2-F1 lifting-body glider (left) and its successor, the M2-F2. Of the experimental lifting bodies NASA built and flew, the Janus spacecraft would have most resembled these pioneering aircraft. Image credit: NASA.

In 2013, while spending a gleeful Sunday afternoon searching through old patent applications (don't judge me), I stumbled upon an intriguing design for a piloted spacecraft using "staged reentry." I wrote about it on my old Beyond Apollo blog on the WIRED website.

In 2017, I expanded that post with more context details on the history of lifting body research and better illustrations and posted it on this blog (see the link at the end of this post). At the time, the patent application, filed in January 1964 by TRW engineers C. Cohen, J. Schetzer, and J. Sellars and granted in December 1966, remained my only source of information on the staged reentry concept.

No longer. One benefit of working at a university is that journal articles formerly locked up behind paywalls, out of reach of independent scholars on a budget, are now readily accessible. Last month, while spending a gleeful Sunday afternoon searching through the 1965 volume of The Journal of Spacecraft & Rockets, I stumbled upon a staged reentry design named for Janus, the two-faced Roman god of endings and beginnings. Closer examination confirmed that the Janus spacecraft was indeed the unnamed spacecraft of the 1966 patent.

Janus is an apt name for the proposed spacecraft design, because its most unique features are related to launch and (especially) landing - that is, the beginning and ending of its mission. The name was first used in a confidential May 1962 TRW Space Technology Labs report by I. Spielberg and C. Cohen.

Spielberg, whose name does not appear on the patent application, presented the staged reentry concept at the first conference of the American Institute of Aeronautics and Astronautics in Washington, DC (29 June-2 July 1964) along with Cohen, whose name was the only one to appear on the 1962 report, the 1964 presentation, the 1965 Journal of Spacecraft & Rockets paper based on the presentation, and the 1966 patent. It seems likely, given his continuous involvement, that Cohen originated and championed the Janus staged reentry concept.

Patent applications are not engineering papers; or, perhaps, one may say that lousy is the engineering paper that reads like a patent application. In addition to being more readable, the 1965 Spielberg and Cohen paper provides considerably more detail than the patent application.

The TRW engineers explained the rationale behind the staged reentry concept:

A manned system should provide precision and flexibility in its landing characteristics. It should be capable of routine launch and routine return without a large recovery task force. Moreover, these criteria must be satisfied without curtailing payload volume or weight or reducing the reliability of reentry protection. In general, these requirements conflict, since efficient entry vehicles (e.g., blunt lifting bodies) have poor landing characteristics, whereas vehicles that land well (winged configurations) tend to have low volumetric efficiency and serious reentry design problems. The staged reentry concept. . . circumvents the difficult design compromises that otherwise must be made to ensure good landing qualities, high volumetric efficiency, and desirable reentry characteristics.

The Janus spacecraft comprised two parts that would separate in flight. The largest part was a 26.8-foot-long, 16-foot-wide, 10-foot-deep "pod." Designed to carry three astronauts, it was an 11,660-pound half-cone lifting body with flat aft and top surfaces and a curved, blunt nose.

The TRW engineers described the pod's double-walled structure. Its inner hull, the pressure vessel, would be manufactured from aluminum sheet. The outer hull would be made of aluminum honeycomb with aluminum alloy plates for added strength. Aluminum frames with "I" and "Z" cross-sections would link the two hulls. An ablative heat shield (that is, one that chars and erodes to carry away heat) would cover the aluminum honeycomb, and low-density insulation would fill the space between the inner and outer hulls.

Cutaway view of the Janus spacecraft. Image credit: U.S.Patent Office.

The other part of the Janus spacecraft was a 4000-pound delta-wing jet aircraft measuring 21 feet long, 13.3 feet across its wings, and 5.33 feet tall. It would include twin downward facing rudder fins and a belly-mounted air intake feeding a Continental 356-23 turbojet engine. The engine could be started at 18,000 feet of altitude using ambient air or at up to 45,000 feet with supplemental oxygen. Cruise speed at 30,000 feet was about Mach 0.6 (370 knots) and range with a full load of 77 gallons (500 pounds) of jet fuel was 200 nautical miles.

The flat top of the small jet would form the largest part of the top of the lifting body. The jet's underside would form the "ceiling" of the lifting body's 860-cubic-foot pressurized internal volume; that is, the plane's belly, including its air intake, would protrude into the main crew living and working space. Ceiling height, though variable, would measure no less than seven feet.

The jet would ride on three rod-like "pneumatic/explosive actuators" attached to the pod. Latches would link the actuators to holes in the plane's nose and on the underside of its wings. Other latches would anchor the jet's wing leading edges.

Spielberg and Cohen recognized that creating an air-tight seal between jet and pod would pose significant design challenges. They proposed an inflatable or "fluted" (grooved) gasket, presumably made of a rubberized fabric. They admitted that their seal system, though "feasible," was not yet "optimized."

Atop a booster on the launch pad, jet and lifting body would point their noses at the sky. Spielberg and Cohen envisioned that the flat aft surface of the pod would sit atop a launch vehicle adapter that would measure 10 feet in diameter where it linked to the pod. The bottom of the adapter would match the larger diameter of the launch vehicle upper stage.

Just before launch, the astronauts would pass through a hatch in the side of the adapter. Overhead they would see the flat aft surface of the pod, which would include a round hatchway. The hatchway would lead into a cylindrical airlock just large enough to hold one space-suited astronaut. A round hatch in the airlock would in turn lead into the pod. In the near-vacuum of low-Earth orbit, the airlock would permit astronauts to spacewalk without depressurizing the pod.

Forward-facing crew couches would be arranged single-file, one behind the other, in a line beneath the jet fuselage. This would place the astronauts one above the other on the launch pad.

The pod would contain the Janus spacecraft main control console. Intended for use in orbit, it would be mounted on the pod's aft interior wall next to the inner airlock hatch. This would place it out of reach of the reclining astronauts. Critically important controls would be mounted on couch arms.

The patent application said nothing about possible launch vehicles, but in their paper Spielberg and Cohen specified two candidates: Titan III (probably the Titan IIIC variant) and Saturn C-1 (otherwise known as Saturn I). The former could boost 28,000 pounds into the 140-nautical-mile orbit required to forestall orbital decay long enough to carry out a two-week Janus mission; the latter, 20,000 pounds. The total weight of the Janus spacecraft (crew, pod, and jet) was 15,660 pounds, so in theory it could transport 12,340 pounds of unspecified payload if launched on a Titan III and 4340 pounds if launched on a Saturn C-1.

It is worth noting that Janus included no docking mechanism, and that was it not designed to perform significant maneuvers in space (apart from a deorbit burn). This ran against the grain of NASA requirements in the first half of the 1960s, when both Gemini and Apollo were under development. Though it could carry a hefty payload, it could not deliver it anywhere. Presumably, this meant that its payload would always take the form of equipment that would remain inside the pod. It is conceivable, however, that small payloads could be tossed out its airlock and larger ones assembled outside by spacewalkers — Spielberg and Cohen did not, however, suggest these possibilities.

A successful mission would begin with launch from Cape Kennedy on Florida's east coast. The launch vehicle would ascend vertically, then roll toward the southeast on a course that would avoid Caribbean islands and South America. About 10 minutes after liftoff, Janus would reach its operational orbit and separate from the upper stage of its launch vehicle. The crew would then unstrap from their couches and begin work in the pod's large pressurized volume.

They would also work in the jet cockpit. The jet's glass canopy, which would stand higher than the rest of the Janus spacecraft's mostly flat top, would make the cockpit the prime spot for conducting Earth and astronomy observations.

Spielberg and Cohen proposed a novel method for entering and leaving the cockpit. The crew couches would each be mounted on a pair of rails, and the underside of the jet's fuselage would include automatic doors. Operating controls on the couch arms would cause the doors to open and the couch to ride the rails from pod to cockpit and vice versa. The TRW engineers explained that a single set of couches shared between the pod and the jet would save weight, though with the large Janus payload capability this would probably have been a minor concern.

The crew would breathe a 47% oxygen/53% nitrogen air mix at a pressure of 7.5 pounds per square inch. Water for crew needs would come from fuel cells, the primary task of which would be to generate 2.5 kilowatts of continuous electricity by combining liquid hydrogen and liquid oxygen. Fluid circulating in pipes in the pod walls would gather and carry waste heat from the pressurized volume and the fuel cells to a radiator mounted on the pod's aft surface.

For return to Earth, the astronauts would sit in their couches in the pod, turn the Janus spacecraft using small thrusters so that its aft end pointed in its direction of motion, and ignite its 1100-pound solid-propellant retrorocket. After burnout, the retrorocket casing would be cast off and Janus reoriented with its nose aimed forward. Descent toward 400,000-foot reentry altitude would last 14 minutes. At start of reentry, the Janus spacecraft would be moving at about 250 feet per second (fps).

Reentry would be a balancing act. The lifting-body pod would need trim flaps for stability and steering; however, four trim flaps attached in pairs to the bottom edge of its flat aft surface would tend to tip its nose down (that is, give it a negative angle of attack). This would permit hot reentry plasma to course over the pod's top surface, destroying the jet canopy. At the same time, the pod would be tail-heavy, raising its nose and making it aerodynamically unstable.

Spielberg and Cohen proposed a two-part solution: cautiously reshaping the pod's nose and packing its triangular nose volume with heavy subsystems (for example, the fuel cells and their reactants). The former would tend to level its angle of attack and the latter, they calculated, would shift its center of gravity forward to a point 54% of its length (about 11 feet) aft of the pod's nose, yielding a slightly "nose up" angle of attack. The pod's nose would thus bear the brunt of reentry heating, and no reentry plasma would reach the jet canopy.

The Janus spacecraft would reenter at a very shallow angle (just 2°). It would thus shed speed gradually in a low-density atmosphere, preventing maximum deceleration from exceeding 1.9 gravities. An automated attitude control system would operate the trim flaps and small thrusters to maintain stability as the pod descended.

During reentry, the outer hull, safe behind its heat shield, would maintain a temperature below 600° Fahrenheit (F). The inner hull would remain at 70° F throughout the mission. The hot outer hull would tend to expand. If the aluminum frames linking the inner and outer hulls were rigidly attached at both ends, differential expansion would tear them apart. To avoid this, Spielberg and Cohen proposed that the frames be attached to the outer hull by flexible connections and to the inner hull by rigid ones.

A little less than 12 minutes after reentry start, at an altitude of about 120,000 feet, the Janus spacecraft would slow to a velocity of about 50 fps. Deprived of lift, its angle of descent would increase in a little over a minute to about 55°.

At 50,000 feet of altitude, the Janus spacecraft would slow to subsonic speed and begin to lose stability. The mission commander would activate the motors that would raise the three couches into the jet cockpit. Beneath the astronauts' feet, the fuselage doors would close and seal. At 45,000 feet, the spacecraft would slow to Mach 0.9, and jet separation from the pod could occur.

Separation would begin with a command to fire explosive bolts. This would release the latches linking the jet to the pod so that the three rod-like pneumatic actuators could extend, pushing the jet away from the pod with a jolt. The pressure seal would be breached, exposing the pod's interior to the outside environment.

The commander would ignite the jet's engine and fly at a cruise altitude of 30,000 feet to a waiting airfield up to 200 nautical miles away. The jet would land on a nose wheel and skids attached to the ends of its rudder fins. The pod, meanwhile, would deploy parachutes from its aft surface and descend to a landing on its nose.

In the event of an abort on the launch pad or during first-stage operation, a pair of solid-propellant abort rocket motors mounted on the pod's aft surface outside the adapter linking it to the launch vehicle would ignite to boost the Janus spacecraft up and away. The motors would propel it to an altitude of 6600 feet in 19 seconds. If no first-stage abort took place, the abort motors would eject after second-stage ignition so that the launch vehicle would not need to carry their weight to orbit.

The deorbit rocket motor would play two possible abort roles: in an abort off the launch pad, it could be ignited after the twin abort rocket motors burned out to boost the Janus spacecraft higher and farther downrange, providing more time for successful jet separation; it would also become the primary abort rocket motor after the twin abort motors ejected.

An abort within 200 nautical miles of Cape Kennedy would see the commander separate the jet from the pod as during a normal descent, then fly back to the launch site. The jet could also remain attached to the pod throughout the abort, in which case the entire Janus spacecraft would descend nose down on parachutes to a landing or splashdown at 25 feet per second. Spielberg and Cohen included 1030 pounds of recovery gear in the Janus spacecraft mass budget.

Down-range aborts — for example, during second stage flight — would occur over open ocean, placing land — never mind suitable airports — outside the jet's 200-nautical-mile range. Spielberg and Cohen noted that the lifting body would during second-stage flight be high enough to use its trim flaps and steering thrusters to maneuver closer to land. This would, they judged, permit jet separation within 200 miles of airfields on Caribbean islands or in northeastern South America.

Here is the link to my staged reentry post based only on the Cohen, Schetzer, and Sellars patent of December 1966. In addition to a summary history of lifting body development in the United States, the post contains detailed labeled drawings from the patent application.

Sources

"Janus: A Manned Orbital Spacecraft with Staged Re-Entry," I. N. Spielberg and C. B. Cohen, The Journal of Spacecraft & Rockets, Volume 2, Number 4, July-August 1965, pp. 531-536.

Patent No. 3,289,974, "Manned Spacecraft With Staged Re-Entry," C. Cohen, J. Schetzer, and J. Sellars, TRW, 6 December 1966.

Related Links

X-15: Lessons for Reusable Winged Spaceflight (1966)

Where to Launch and Land the Space Shuttle? (1971-1972)

What if a Shuttle Orbiter Struck a Bird? (1988)

NASA Johnson Space Center's Shuttle II (1988)