Star-Raker (1978)

Star-Raker (right), a single-stage-to-orbit space plane, parks next to a 747 at a conventional airport. Image credit: M. Alvarez/Rockwell International.

Elsewhere in this blog, I have described the 1970s joint NASA/Department of Energy Solar Power Satellite (SPS) studies (see "More Information" below). Had even a single SPS been assembled, it would have been by far the largest human construction project in space; it would have weighed more than 100 times as much as the 420-metric-ton (460-U.S.-ton) International Space Station. The SPS studies envisioned assembly of two such satellites per year between 2000 and 2030, bringing the total number in the SPS constellation to sixty. 

NASA envisioned boosting SPS components to low-Earth orbit (LEO) in the payload bays of massive reusable launch vehicles. One such launcher, Boeing's winged, two-stage Space Freighter, would have weighed about 11,000 metric tons (12,125 U.S. tons) at liftoff and delivered about 420 metric tons (463 U.S. tons) to LEO. For comparison, the two-stage Saturn V rocket used to place 77-metric-ton (85-U.S.-ton) Skylab into LEO weighed about 2800 metric tons (3086 U.S. tons) at liftoff.

The Space Freighter would have risen vertically from a launch pad and pointed itself generally toward the east. As its first stage, the Booster, expended its propellants, it would have separated. The second stage, the Orbiter, would then have ignited its engines to complete its climb to LEO. In orbit, it would have maneuvered to rendezvous and dock with a large space station designed specifically for handling SPS cargo modules.

The Space Freighter Booster would have been a fully reusable winged vehicle closely resembling the Space Freighter Orbiter. After Space Freighter Orbiter separation, the Space Freighter Booster would have turned, deployed jet engines, and flown to a long, wide runway at its launch site. 

To begin return to Earth, the Space Freighter Orbiter in LEO would have separated from the cargo-handling space station, then would have turned its tail forward and ignited rocket motors to slow down, lowering its orbit so that it intersected Earth's atmosphere. Following a fiery reentry, it would have landed on the runway near its launch pad. 

After launch pad, Orbiter, and Booster refurbishment, the two Space Freighter stages would have been hoisted vertical. After the Orbiter was placed atop the Booster's nose, a cargo module would have been loaded into its payload bay. The Space Freighter would then have been moved to a launch pad to begin another flight. Launching parts for two SPS into LEO in a year would have required about 240 Space Freighter launches, or about one launch every 36 hours.

In October 1977, a team of 14 Rockwell International engineers studied a Space Freighter alternative. The Star-Raker space plane, 103 meters (310 feet) long with a wing span of about 93 meters (280 feet), would have carried a maximum of 89.2 metric tons (98.3 U.S. tons) of cargo into LEO. More than 1100 flights would have been required each year to support the SPS program, or about one launch every eight hours.

In its fully developed form, however, Star-Raker would have had important advantages over Space Freighter which might have made its required high flight rate feasible. For example, it would have begun its flights to LEO by taking off horizontally from a conventional 2670-to-4670-meter-long (8000-to-14,000-foot-long) runway at virtually any civilian or military airport capable of supporting 747 or C-5A Galaxy cargo planes. No specialized launch and landing site would have been required.

Every bit as important, Star-Raker would have been capable of flying routinely between such airports. The Rockwell team explained that this would "reduce the number of operations required to transport material and equipment from their place of manufacture on Earth to [LEO]." For example, rolls of solar cell blankets would not need to be shipped by train, barge, or plane to a specialized launch and landing site; they would, potentially, need only be transported to a local airport for Star-Raker pickup.

Though the 1977-1978 Star-Raker study focused on its possible use in the Department of Energy/NASA Solar Power Satellite program, Star-Raker would have had potential as a general-purpose space cargo plane. In the image above, three Star-Rakers, their nose sections hinged back to expose their cargo bays, take on payloads bound for destinations ranging from low-Earth orbit to deep space. Image credit: M. Alvarez/Rockwell International.

David Reed, an engineer at North American Rockwell (NAR), as the company was then known, originated the Star-Raker concept in 1968, as NASA began earnest efforts to develop a reusable Space Shuttle. Key elements of the concept had been proposed — and rejected — earlier in the 1960s decade. These included wings packed with lightweight structurally integral tanks holding liquid hydrogen fuel and liquid oxygen oxidizer and a complex jet engine/rocket engine propulsion system.

The 1968-1969 study determined that, as it burned the propellants in its wings and maneuvered through ascent from subsonic speed to Mach 6 (six times the speed of sound), aerodynamic pressure on its structure would become excessive. This led NAR to examine wing designs developed in 1970 for the proposed (and subsequently abandoned) U.S. Supersonic Transport program. 

A "tridelta flying wing" design appeared to solve the pressure problem; by then, however, NASA had narrowed its Shuttle design requirements, excluding Star-Raker from consideration. NAR continued Shuttle studies and became Shuttle prime contractor in July 1972. 

Rockwell revived study of the tridelta flying wing Star-Raker as SPS studies ramped up in 1976. The Star-Raker study that began in October 1977, led by Reed and performed for NASA Marshall Space Flight Center (MSFC) in Huntsville, Alabama, continued into late 1978, yielding the design described in this post.  

The 1977-1978 study benefited from computer modeling that enabled Rockwell to further refine Star-Raker wing shape and flight profile. It also allowed Reed's team to take more fully into account the benefits of propellant-saving "lifting ascent." 

Star-Raker's propellants, liquid hydrogen and liquid oxygen, were not typically found at airports in 1977-1978; this remains true in 2020. The Star-Raker study team might have assumed that airports would evolve to provide them by the time SPS cargo flights began in 2000. This would, perhaps, not have been an unreasonable assumption, given that the 30-year SPS program was expected to create a lucrative new industry spanning the continental United States. 

One Star-Raker takes off as another undergoes airport servicing. With its landing gear extended, Star-Raker ground clearance would have been 1.52 meters (five feet). Image credit: M. Alvarez/Rockwell International.

For the 1977-1978 study, however, they hedged their bets by assuming that liquid hydrogen fuel would be available at airports only in sufficient quantities for airport-to-airport subsonic air-breathing jet engine Star-Raker flights. Liquid oxygen would, of course, not have been required. Flights to LEO, which would have needed both propellants in large quantities, would have begun on a runway at NASA's Kennedy Space Center (KSC) in Florida, at Vandenberg Air Force Base, California, or at any other launch sites the U.S. might have deigned to establish. 

The propellant tanks in Star-Raker's wings would have been approximately conical in shape. They would have extended from the space plane's body to its wing tips and been designed to strengthen the wings with minimal weight penalty. They would have been reinforced with regularly spaced "cell web" walls. Foam-filled glass-fiber honeycomb would have surrounded the tanks, defining Star-Raker's shape.

The Rockwell team described in detail a Star-Raker flight from KSC to 556-kilometer-high (345-mile-high) LEO and back to a U.S. airport. It would have begun with arrival at KSC of a Star-Raker space plane loaded with cargo bound for LEO at the end of a subsonic flight from a conventional airport. 

Following a limited airplane-type checkout, crews would have installed three sets of jettisonable orbital-takeoff main landing gear, each with eight wheels, and pumped liquid hydrogen and liquid oxygen propellants into Star-Raker's tanks. Fully loaded with propellants and cargo and with its orbital-takeoff gear attached, Star-Raker would have weighed about 1935 metric tons (2130 U.S. tons). 

Star-Raker would have lifted off from the runway at a speed of 415 kilometers per hour (260 miles per hour) under "supercharged afterburner" power from its 10 multicycle jet engines. The Rockwell team explained that it had consulted with leading jet engine manufacturers to arrive at its jet engine design; these included General Electric, Pratt & Whitney, Aerojet, Marquardt, and Rocketdyne. The resulting engine was more a wish list than a firm design, though it was an informed wish list. 

The Rockwell team envisioned four operational cycles for its jet engine ranging from conventional turbofan to ramjet. Liquid hydrogen would have been used to cool the engine and then burned as fuel. Large, slot-shaped inlets on the underside of Star-Raker's wings, arranged in two groups of five on either side of the space plane's body, would have funneled air to the engines, which would have been mounted at the wing trailing edge. The inlets would have been equipped with "ramp" doors that could close partially or fully to moderate or halt airflow.

Shortly after leaving the ground, the space plane's crew would have dropped the three sets of orbital-takeoff landing gear (they would have lowered to the ground on parachutes for recovery and reuse), then would have retracted its nose and main landing gear. The space plane would then have switched its jet engines to turbofan power, climbed to 6100-meter (20,000-foot) cruise altitude, and increased its speed to Mach 0.85. It would have turned due south and, over the next hour and fifty minutes, flown directly to Earth's equator.

Star-Raker would have flown to the equator and turned east so that it could get a boost from Earth's rotational velocity, which at our planet's midriff can, in theory, add about 1600 kilometers (1000 miles) per hour to the orbital velocity of ascending launch vehicles. 

In addition, and more importantly, the turbofan flight to the equator would have amounted to a plane-change maneuver; that is, it would have enabled Star-Raker to reach equatorial LEO without performing the rocket-propelled plane-change maneuver in LEO required if Star-Raker flew directly to orbit from a non-equatorial launch site, such as KSC. The Rockwell team hoped that this would save propellants, enabling an increase in cargo weight.

Following the eastward turn, the space plane would have climbed to 13,710 meters (45,000 feet) under supercharged afterburner power, then would have begun a shallow dive to 11,280 meters (37,000 feet). During the powered dive, a propellant-saving maneuver, Earth's gravity would have helped it to break the sound barrier and accelerate to Mach 1.2. 

Go for orbit: the Star-Raker space plane design included 10 multicycle air-breathing jet engines, three high-pressure rocket engines akin to the Space Shuttle Main Engine, and two advanced Orbital Maneuvering System rocket engines. In the image above, the 10 jet engines are throttling up to begin the transition to supersonic flight. Image credit: M. Alvarez/Rockwell International.

Star-Raker would then have begun ascent to orbit in earnest, with a supersonic climb to 29 kilometers (18 miles). During this phase, the space plane's jet engines would have throttled up to "full ramjet" power, accelerating it to Mach 6.2. Throughout its climb to orbit, Star-Raker would have maneuvered to put to good use lift provided by its wings. 

Upon reaching Mach 6.2, the three rocket motors in Star-Raker's tail would have ignited, adding rocket power to ramjet power. The three engines, with a combined thrust of 1.45 million kilograms (3.2 million pounds), would have drawn liquid hydrogen from a sturdy tank located at the aft end of the long, narrow Star-Raker cargo bay. The tank, to which the engines would have been mounted, would have served as the load path that would have distributed their thrust to the space plane's structure.

At Mach 7.2, Star-Raker would have switched to full rocket power. As it throttled up the rocket motors to full thrust, it would have shut down the jet engines and closed completely their air inlet doors. 

When Star-Raker reached a 51-kilometer-by-556-kilometer (32-mile-by-345-mile) equatorial orbit, the main rocket motors would have shut down. At apogee, the high point in its orbit, the crew would have ignited the twin advanced Orbital Maneuvering System (OMS) engines at the base of its tail to raise its perigee (orbit low point) and circularize its orbit. Upon attainment of circular equatorial orbit, Star-Raker would have used the OMS to maneuver to a rendezvous with the SPS cargo-handling space station.

Star-Raker in low-Earth orbit. Image credit: M. Alvarez/Rockwell International.

The weight of cargo Star-Raker could carry would depend on its mission profile. For the profile described here, cargo weight delivered to orbit would have amounted to only about 48.6 metric tons (53.6 U.S. tons). The aerodynamic flight to the equator under jet power, meant to steal some of the Earth's rotational energy and avoid a plane change maneuver in LEO, had under close examination turned out to be expensive. 

The Rockwell team proposed improving the equatorial profile's payload performance by loading liquid oxygen at the equator, either during flight using a new-design tanker aircraft, or after a landing at an equatorial facility with an adequate runway, orbital-takeoff gear attachment and recovery capability, and ability to provide liquid oxygen. Either approach would, however, have complicated Star-Raker operations.

To unload cargo, Star-Raker would have swung its nose, which would have contained its crew compartment, sideways out of the way, exposing one end of its six-meter-high-by-six-meter-wide-by-43-meter-long (20-foot-high-by-20-foot-wide-by-141.5-foot-long) cargo bay. The bay's arched ceiling would have made it a point of structural strength, not weakness, in the Star-Raker design.

The crew would have moved to the rear of the crew compartment to assist with cargo transfer. Windows at the rear of the two-deck crew compartment would have provided a 121° field of visibility. 

The Rockwell team did not describe its cargo transfer system in any detail, though it is clear that Star-Raker would not have docked in the conventional sense. Brief mention was made of a transfer rail system in the cargo bay that would have linked to equivalent rails on the space station.

Return to Earth would have begun with cargo bay closure. After moving away from the space station, the crew would have turned Star-Raker so that its tail faced in its direction of orbital motion, then would have fired its OMS engines to slow down. 

Maximum deceleration during the unhurried shallow-angle reentry would have reached no more than 2.3 gravities. Star-Raker would, in general, have experienced reentry temperatures lower than the Space Shuttle Orbiter, though nose and wing leading-edge temperatures were expected be somewhat higher. The higher leading-edge temperature was attributable to its relatively blunt shape. 

The Rockwell team proposed two types of reusable Thermal Protection System (TPS) for Star-Raker. Both would have been mounted on an outer facing sheet covering a honeycomb layer. The honeycomb layer would in turn have been attached to an inner facing sheet covering the honeycomb core that surrounded the propellant tanks.

The first TPS design closely resembled that baselined for the Space Shuttle Orbiter. Ceramic tiles individually molded and milled to match Star-Raker's curves would have been glued to fabric strain-isolator pads affixed to the outer facing sheet. 

The second TPS design, similar to one developed for the B-1 Bomber, was more complex. Metal panels — titanium-aluminum for low-temperature areas and "superalloy" for high-temperature areas — would have been attached to the outer facing sheet using flexible standoffs. The standoffs would have permitted the overlapping panel edges to slide over each other as they grew hot and expanded or cooled and contracted. Foil-wrapped thermal insulation blankets affixed to the outer facing sheet would have provided additional thermal protection.

Both TPS designs would have included a system for detecting breaches in the TPS. The Rockwell team provided no details of its design and did not describe what the crew might do if a breach were detected.

Star-Raker on approach. Image credit: M. Alvarez/Rockwell International.

When Star-Raker slowed to Mach 6, it would have begun cross-range maneuvers designed to shed energy and slow it to Mach 0.85. The crew would then have opened the inlet ramps and started "some" of its jet engines. 

The Rockwell team provided the space plane with enough liquid hydrogen to permit a 556-kilometer (345-mile) subsonic cruise and two powered landing attempts. Landing velocity would have been about 215 kilometers per hour (135 miles per hour). At wheels stop at an airport capable of supporting a cargo 747 or a C-5A Galaxy, Star-Raker would have weighed about 281 metric tons (310 U.S. tons).

Star-Raker weights given in this flight description are based on data the Rockwell team generated in the period spanning December 1977-January 1978. In February-March 1978, NASA MSFC and NASA Langley Research Center (LaRC) in Hampton, Virginia, reviewed the Rockwell team's Star-Raker weight numbers. 

The NASA centers found that Rockwell's estimates were low if "normal" technology were assumed and high if "acceleration" (advanced) technology were assumed. Whereas Rockwell had placed Star-Raker's "dry" weight with orbital-takeoff gear at 293.5 metric tons (323.5 U.S. tons), MSFC/LaRC determined that, with normal technology and a 10% cushion for weight growth during development, Star-Raker would weigh 407.6 metric tons (449.3 U.S. tons) without propellants; with advanced technology and the cushion, it would weigh only 257.6 metric tons (284 U.S. tons). 

The Rockwell team and NASA MSFC engineers met in May 1978 to try to reconcile the weight estimates. They made one important change in Star-Raker's flight profile: they abandoned the subsonic flight to the equator in favor of a KSC launch and direct climb to a 556-kilometer (345-mile) LEO inclined 28.5° relative to Earth's equator (that is, the latitude of KSC). 

The NASA and Rockwell teams settled on a Star-Raker weight without propellants (but with orbital-takeoff gear and 10% cushion) of 330.4 metric tons (364.2 U.S. tons). As it began ascent to orbit on a KSC runway, the space plane would have weighed 2280.5 metric tons (2514 U.S. tons). Of this, Star-Raker's maximum weight, 89.2 metric tons (98.3 U.S. tons) would have comprised cargo for the SPS project.

Sources

Independent Research and Development Data Sheet, Project Title: Earth-to-LEO Transportation System for SPS, Rockwell International, 15 December 1978.

"Star-Raker: An Airbreather/Rocket-Powered, Horizontal Takeoff Tridelta Flying Wing, Single-Stage-to-Orbit Transportation System," SSD 79-0082, D. Reed, H. Ikawa, and J. Sadunas, North American Rockwell Space Systems Division; paper presented at the American Institute of Aeronautics & Astronautics Conference on Advanced Technology for Future Space Systems in Hampton, Virginia, 8-11 May 1979.

More Information

Electricity from Space: The 1970s DOE/NASA Solar Power Satellite Studies

NASA Johnson Space Center's Shuttle II (1988)

Venus is the Best Place in the Solar System to Establish a Human Settlement (2003)

A dirigible approaches an outpost in the atmosphere of Venus. Image credit: NASA.

There's no award for "Most Imaginative Space Engineer," but if there were, Geoffrey Landis would certainly be a top contender. In fact, if such an award is ever created, it should perhaps be named the Geoffrey, in parallel with science fiction's Hugo Award, which owes its name to pioneering author, editor, and publisher Hugo Gernsback. Not incidentally, Landis owns a pair of Hugos; he received his first in 1992 for "A Walk in the Sun," a short story set on the Moon, and his second in 2003 for his story "Falling Onto Mars."

Landis is an engineer at NASA's Glenn Research Center (GRC) in Cleveland, Ohio. Much of his NASA work has centered on energy systems, with an emphasis on solar photovoltaic power. 

In a brief paper prepared for the February 2003 Space Technology and Applications International Forum in Albuquerque, New Mexico, Landis made a compelling case for Venus, not the Moon, nor Mars, nor a twirling sphere, torus, or tube in open space, as the ideal place to establish an off-Earth human settlement. Specifically, he set his sights on the Venusian atmosphere just above the dense sulfuric-acid clouds. Landis called it "the most earth-like environment (other than the Earth itself) in the Solar System." 

Most people think of Venus as a hell planet because they think only of its surface. By about 1960, scientists using Earth-based instruments had determined that Venus had a temperature of 342° C (648° F). Many, however, refused to believe that Venus could be so hot. Some tried to find a loophole: they hypothesized that the Venusian atmosphere was hot while its surface was cool enough for liquid water and life.

Mariner 2, the first successful interplanetary spacecraft, flew past Venus in December 1962. Its crude scanning radiometer found a lower temperature — around 230° C (450° F) — though one still much higher than most planetary scientists expected. Mariner 2 also determined that air pressure at the Venusian surface was at least 20 times Earth sea-level pressure.

For more than two decades, Venus was the Soviet Union's favorite Solar System exploration target. The Venera landers determined that its surface is made of basalt, a volcanic rock. They also found that the mean atmospheric pressure at the surface is 96 times Earth sea-level pressure and that the surface temperature averages about 462° C (863° F) with relatively modest day/night, latitude, and altitude variations

The Venusian atmospheric temperature, on the other hand, was found to vary significantly with altitude, a fact that the Soviet Union would put to good use. In June 1985, the Vega 1 and Vega 2 spacecraft released armored landers and lightly constructed rubber balloons as they flew past Venus on their way to Comet Halley. The Vega 1 lander touched down but returned minimal data. Vega 2 landed successfully and survived the hellish surface conditions for 56 minutes. 

The twin three-meter-diameter, helium-filled balloons deployed between 50 and 55 kilometers (34 and 31 miles) above the Venusian surface — that is, just above the cloud-tops, in the zone Landis saw as promising for human settlement. Their small instrument payloads transmitted data for approximately two days — until they exhausted their chemical batteries. 

In that time, the balloons rode the carbon dioxide winds from their deployment points over the nightside into bright Venusian daylight. The Vega 2 balloon travelled about 11,100 kilometers (6900 miles) and the Vega 1 balloon travelled 11,600 kilometers (7210 miles). When their instrument payloads exhausted their batteries, the balloons carrying them showed no sign of imminent failure. They might have lasted for months or even years.

Vega-type balloon on display at the National Air and Space Museum's Udvar-Hazy Center in northern Virginia, just outside Washington, DC. Image credit: Geoffrey A. Landis. 

The fragile balloons could last so long because 50 kilometers above Venus, just above the cloud tops, the temperature ranges from between 0° C to 50° C (32° F to 122° F) and the atmospheric pressure approximates Earth sea-level pressure. A thin fabric cover was sufficient to shield each balloon from sulfuric acid droplets drifting up from the cloud layer.

Venus settlers would float where Vega 1 and Vega 2 floated, but Landis rejected helium balloons. He noted that, on Venus, a human-breathable nitrogen/oxygen air mix is a lifting gas. A balloon containing a cubic meter of breathable air would be capable of hoisting about half a kilogram, or about half as much weight as a balloon containing a cubic meter of helium. A kilometer-wide spherical balloon filled only with breathable air could in the Venusian atmosphere lift 700,000 tons, or roughly the weight of 230 fully-fueled Saturn V rockets. Settlers could build and live inside the air envelope. 

The air envelope supporting a settlement would not necessarily maintain a spherical form. Lack of any pressure differential would allow the gas envelope to change shape fluidly over time. It would also limit the danger should the envelope tear. The internal and external atmospheres would mix slowly, so the settlement atmosphere would not suddenly turn poisonous, nor would the settlement rapidly lose altitude. 

A repair crew would not require pressure suits, Landis explained. They would, of course, need air-tight face masks to provide them with oxygen and keep out carbon dioxide; adding goggles and unpressurized protective garments would keep them safe from acid droplets.

Acid droplets in the Venusian atmosphere would no doubt be annoying, but Venus would lack the frequent toxic dust storms of Mars. Orbiting nearly twice as close to the Sun as does Mars, a Venusian solar farm would have available four times as much solar energy at all times — and with no dust storms to get in the way. Landis noted that solar panels could collect almost as much sunlight reflected off the bright Venusian clouds as they could from the Sun itself. 

Mars, the Moon, and free-space habitats all must contend with solar and galactic-cosmic ionizing radiation. A settlement 50 kilometers above Venus, by contrast, could rely on the Venusian atmosphere to ward off dangerous radiation. Radiation exposure would be virtually identical to that experienced at sea level on Earth.

Many aspiring space settlers assume that humans and the plants and animals they rely on (or simply like to have around) will be able to live in one-sixth or one-third Earth gravity — the gravitational pull felt on the Moon and on Mars, respectively — with no ill effects. The hard reality, however, is that no one knows if this is true. It is possible that astronauts living in hypogravity — that is, gravity less than one Earth gravity — will experience health effects similar to those they experience during long stays in microgravity (for example, on board the International Space Station). 

Venus is nearly as dense and as large as Earth, so its gravitational pull is about 90% that of humankind's homeworld. The likelihood that hypogravity will make long-term occupancy unhealthful might thus be reduced. 

The Venusian atmosphere is rich in resources needed for life and the Venusian surface, while hellish, would lay only 50 kilometers away from the settlement at all times. Landis suggested that Venus settlers might use a suspended super-strong cable to lift silicon, iron, aluminum, magnesium, potassium, calcium, and other essential chemical elements to the floating settlement. He noted that laboratory experiments aimed at producing robots hardy enough to function on Venus for long periods had already begun; operators might use such rovers to remotely mine the surface from the comfort of the floating settlement.

Landis pointed to the Main Asteroid Belt between Mars and Jupiter as a potential source of resources for Venus.  He noted that any given asteroid in the Main Belt is easier to reach from Venus than from the Earth or Mars. A spacecraft launched from Venus on a minimum-energy trajectory can, for example, reach resource-rich 1 Ceres, the largest asteroid, in a little less than an Earth year; a minimum-energy trip from Earth to 1 Ceres would need a little more than an Earth year. 

Image credit: NASA.

The large Main Belt asteroids are in fact generally located farther away from each other than they are from Venus. They also orbit the Sun much more slowly: 3 Vesta needs 1325 Earth days to circle the Sun once; 1 Ceres needs 1682 Earth days; 2 Pallas, 1686 Earth days; and 10 Hygeia, in the outer part of the Main Belt, 2035 Earth days. This means that minimum-energy transfer opportunities between Main Belt asteroids occur years or even decades apart. Opportunities for minimum-energy transfers between Venus and any Main Belt asteroid, on the other hand, occur about once per Venus year (that is, about once every 225 Earth days).

As the journeys of the twin Vega balloons illustrate, Venus atmosphere settlements would ride fast winds. Those near the equator would circle the planet every four days. This would mean, Landis explained, that they would experience a day/night pattern of two days of darkness followed by two days of light. He expected that settlements eager for a more Earth-like lighting pattern could migrate to the Venusian circumpolar regions, where a circuit around the planet would be shorter. 

If many "cloud cities" were eventually established in the atmosphere of Venus, then a preference for the poles might lead to crowding. If, on the other hand, any latitude were fair game, then Venus would offer for settlement a total area 3.1 times Earth's land area — that is, more than three times greater than the surface area of Mars. Landis wrote that, eventually, a "billion habitats, each one with a population of hundreds of thousands of humans, could. . . float in the Venus atmosphere."

Sources

Mariner Venus 1962 — Final Project Report, NASA SP-59, NASA Jet Propulsion Laboratory, 1965.

Soviet Space Programs 1980-1985, Nicholas L. Johnson, Volume 66, Science and Technology Series, American Astronautical Society, 1987, pp. 186-188.

"Colonization of Venus," Geoffrey A. Landis, Space Technology and Applications International Forum (STAIF) 2003, Albuquerque, New Mexico, 2-5 February 2003; American Institute of Physics Proceedings 654, Mohamed S. El-Genk, editor, 2003, pp. 1193-1198.

More Information

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

Venus as Proving Ground: A 1967 Proposal for a Piloted Venus Orbiter

Floaters, Armored Landers, Radar Orbiters, and Drop Sondes: Automated Probes for Piloted Venus Flybys (1967-1968)

Two for the Price of One: 1980s Piloted Missions with Stopovers at Mars and Venus (1968)

Multiple Asteroid Flyby Missions (1971)

Footsteps to Mars (1993)

Mars Polar Ice Sample Return (1976-1978)

This oblique view of the southern polar ice cap of Mars (bottom) includes the entire permanent cap and a small portion of the adjoining temporary cap. Much of the southern hemisphere is in view; for example, the large Hellas impact basin is visible at center left. At its fullest extent in southern hemisphere midwinter, the seasonal cap expands to touch the southern edge of Hellas. Image credit: NASA.

Mars, like its neighbor Earth, has ice caps at its north and south poles. On both worlds, the polar caps are dynamic; for example, they expand and contract with the passage of the seasons. On Earth, both the permanent and seasonal polar caps are made up entirely of water ice; on Mars, which is generally colder, temperatures fall low enough that carbon dioxide condenses out of the atmosphere at the winter pole, depositing a frost layer about a meter thick on the permanent water ice polar cap and surrounding terrain. 

The three-kilometer-thick permanent caps cover a little more than one percent of the martian surface. In northern hemisphere midwinter, the seasonal carbon dioxide cap expands to about 60° north latitude. Roughly 13 Earth months later, in southern midwinter, carbon dioxide ice covers the cratered landscape to about 60° south latitude.

Confirmation that the permanent polar caps are made up mainly of water ice did not come easily. The polar caps were first glimpsed using crude telescopes during the 17th century, and were widely believed to be made of water ice by the end of the 18th. In 1965, however, data from Mariner IV, the first spacecraft to fly past Mars, indicated that the permanent caps were made of frozen carbon dioxide, an interpretation the Mariner 6 and 7 flybys (1969) and the Mariner 9 orbiter (1971-1972) did little to contradict.

In the late 1970s, however, the twin Viking Orbiters revealed that the northern permanent cap is made of water ice. Confirmation that the southern permanent cap is also made of frozen water had to wait, however, until 2003, when data from the Mars Global Surveyor and Mars Odyssey orbiters had become available.
Close-up of the southern permanent water ice cap of Mars in southern hemisphere summer. In winter, the entire image would be cloaked in red dust and carbon ice and frost. Image credit: NASA. 

In 1976-1977, before the composition of either of the permanent caps was known with certainty, a team of students in the Purdue University School of Aeronautics and Astronautics studied a Mars Polar Ice Sample Return (MPISR) mission. Its primary goal was to collect and return to Earth a 50-meter-long, five-millimeter-diameter ice core extracted from the planet's southern permanent cap.

The Purdue students assumed that the permanent ice caps of Mars are, as on Earth, built up of layers of snow or frost deposited annually. They anticipated that each layer would contain a sample of the dust and gases in the atmosphere at the time it was laid down, making it a record of atmospheric particulates and climate conditions. The range of materials captured in the core would enable multiple methods of age determination.

On Earth, ice cores from Greenland record lead smelting in the Roman Empire and vegetation changes in Ice Age Europe. A martian polar ice core, the students believed, might yield a planet-wide record of dust storms, asteroid impacts, volcanic eruptions, flowing surface water, and, quite possibly, the existence of microbial life. 
Section of an ice core collected from deep beneath the Greenland ice cap. Image credit: Greenland Ice Sheet Project.

As the Purdue students carried out their study, the twin Viking spacecraft were en route to Mars. Viking 1 left Cape Canaveral, Florida, on 20 August 1975, and Viking 2 lifted off about three weeks later (9 September 1975). The Vikings were two-part spacecraft — each comprised a Martin Marietta-built 571-kilogram Lander and a 2336-kilogram Orbiter built by the Jet Propulsion Laboratory (JPL).

Viking 1 fired its Orbiter-mounted rocket motor on 19 June 1976 to slow down so that the red planet's gravity could capture it into orbit. The Viking 1 Lander was scheduled to land on the American Bicentennial (4 July), but the landing was postponed after Viking Orbiter images of its prime and backup landing sites, which had been selected using Mariner 9 data, showed them to be too rough. The Viking 1 Lander separated from its Orbiter and performed the first successful Mars landing in Chryse Planitia on 20 July 1976.

Viking 2 reached Mars orbit on 7 August. Its pre-selected landing sites were also found to be too rugged, so touchdown in Utopia Planitia did not take place until 3 September 1976. 

Viking development cost close to a billion U.S. dollars, making it the most expensive automated exploration program of its time. For some planners — possibly unacquainted with the untimely end of the Apollo program — it seemed reasonable to assume that NASA would exploit Viking hardware to the fullest to cash in on its investment. JPL planners, for example, widely expected that a third Viking spacecraft — probably including a rover — would depart for Mars in 1979. For this reason, the Purdue students assumed that Viking hardware would continue to be manufactured into the 1980s and that their MPISR spacecraft could be derived from it. 
Schematic of Viking Orbiter with attached Viking Lander inside protective aeroshell. Image credit: NASA.
Schematic of a Viking Lander deployed on Mars. Image credit: NASA.

The MPISR mission would employ a Mars Orbit Rendezvous (MOR) mission plan equivalent to the Lunar Orbit Rendezvous plan used to carry out Apollo Moon landings. A 5652-kilogram MPISR Orbiter would carry to Mars a 946-kilogram Lander and a 490-kilogram Earth-Return Vehicle/Earth Orbit Vehicle (ERV/EOV). The MPISR Lander would in turn carry a 327-kilogram Ascent Vehicle (AV) for launching the polar ice sample to Mars orbit.

The need for a short-duration flight from Mars to Earth and for south pole conditions safe for landing dictated the MPISR mission's Earth departure date. A long flight back to Earth would place great demands on sample refrigeration equipment, so the Purdue students sought the shortest return opportunity they could find.

Data from the Viking Orbiters had shown the south pole ice cap to be too unstable for landing and sample collection in the spring and summer, when the temperature climbs too high for carbon dioxide to remain solid. At mid-winter, on the other hand, snow and frost accumulation might bury the MPISR Lander. The team proposed, therefore, that the Lander should set down in late summer, about 75 days before southern hemisphere autumnal equinox. 
Schematic of the MPISR spacecraft after Lander separation but before AV third stage arrival. The Earth Return Vehicle near the bottom of the image — which would carry within it the Earth Orbit Vehicle — would be based on the Pioneer Jupiter/Saturn bus design. Image credit: Purdue University.

The MPISR spacecraft would lift off from Cape Canaveral, Florida, on 29 April 1986, in the 15-foot-by-60-foot payload bay of a delta-winged, piloted Space Shuttle Orbiter. It would reach Earth orbit attached to an expendable Tug derived from the U.S. Air Force/NASA Centaur G' upper stage. The Purdue students calculated that the proposed Tug could launch up to 9000 kilograms out of Earth orbit toward Mars during the favorable 1986 Earth-Mars transfer opportunity.

On 16 November 1986, after a flight lasting nearly seven months, the MPISR Orbiter propulsion system would slow the spacecraft so that martian gravity could capture it into a polar orbit. It would then begin a 14-month orbital survey of the martian poles. 

The MPISR Orbiter would map the poles using Viking-type cameras, a Viking-type thermal mapper, and a new-design Radar Ice Sounder for determining ice depth. The sounder, which is not depicted in the MPISR Orbiter image above, would employ an 11.47-meter-diameter dish antenna that would unfold from the Orbiter soon after Mars orbit arrival. Scientists would use data from the Orbiter's instruments to select a safe and scientifically interesting south pole landing site for the MPISR Lander.

On 3 February 1988, the Lander would separate from the Orbiter, ignite solid-propellant rockets to slow down and drop from Mars orbit, then descend through the planet's thin atmosphere to the selected landing site. Because it would have nearly twice the mass of the Viking Lander from which it was derived, the MPISR Lander would lower on six parachutes and six terminal descent rocket engines (in each case, twice as many as Viking). The engines would be arranged in three clusters of two engines each.

Extra engines would complicate deployment of the Lander's most important science system, the 16.3-kilogram Ice Core Drill (ICD). Soon after touchdown, the MPISR Lander would reach out with its modified Viking sampler arm to detach one of the three descent engine clusters to clear the way for ICD deployment.

Sixty-seven times over the next 90 days, the ICD would collect a 75-centimeter-long ice core, raise it to the surface, and deposit it in an insulated 12-kilogram sample container. The final core would sample ice and dust layers hidden 50 meters below the surface. 

The students did not describe ICD operation in detail. No doubt the drilling and core acquisition process would face many challenges. The slender drill might, for example, encounter a patch of compressed crystallized ice and dust at depth and need to start again at a new place within its drill site, which would measure at most two or three square meters in area.

The MPISR Lander's south pole landing site would mean that it could not transmit radio signals directly to Earth. The MPISR Orbiter, for its part, would be able to keep Lander and Earth in view simultaneously for at most 25 minutes per day, sharply restricting radio relay time. This meant that the continuous drilling operation would need to take place autonomously.

Communication limitations, combined with slowly changing, generally unfavorable polar lighting conditions, led the Purdue students to replace the twin Viking scanning facsimile cameras with a simpler vidicon camera akin to that carried on the 1960s Surveyor lunar landers. The camera package would include three strobe lights. This would, they explained, permit the MPISR Lander to capture "snapshots" of the drill site and its novel frosty surroundings for transmission to the Orbiter.

Radioisotope Thermal Generators (RTGs) would power and warm MPISR Lander systems. The Lander's three footpads and underside would be insulated to prevent heat transmitted through its structure from melting the ice, helping to ensure that it would not sink from sight during the three-month sample-collection period. 
Near midwinter: the martian south polar ice cap near maximum extent as viewed from Earth orbit by the Hubble Space Telescope. Image credit: NASA. 

The Mars southern hemisphere autumnal equinox would occur on 17 April 1988. On 2 May 1988, with winter gradually settling in at the martian south pole, the first of the AV's three rocket stages would ignite to blast the ice core samples to Mars orbit. The AV third stage would provide refrigeration in the sample container to keep the ice core sections pristine.

The AV first stage and second stage would burn solid propellants. The liquid-propellant third stage would boost the sample container into a 2200-kilometer circular orbit about Mars, then would commence active maneuvers to perform a rendezvous with the MPISR Orbiter.

On 17 May 1988, the MPISR Orbiter would maneuver to a docking with the AV third stage. A docking collar on the ERV/EOV would dock with the third stage, then the sample container would automatically transfer to the ERV/EOV and the third stage would be cast off.

On 27 July 1988, the ERV/EOV would separate from the MPISR Orbiter and fire its engine to leave Mars orbit for Earth. To reduce the period of time during which the ice core would need refrigeration, the ERV/EOV would expend propellants to speed Earth return. A minimum-energy transfer in the 1988 Mars-Earth opportunity would last 122 days; the ERV/EOV's energetic Mars-departure burn would slash this to as little as 98 days. Arrival in Earth orbit would take place between 2 November and 14 November 1988.

Nearing Earth, the cylindrical 1.5-meter-long EOV would separate from the ERV and fire a solid-propellant rocket motor to slow down so that Earth's gravity could capture it into a 42,200-kilometer circular orbit. The ERV, meanwhile, would speed past Earth into solar orbit.

Discarding the ERV ahead of Earth-orbit capture would slash Earth-orbit insertion mass, dramatically reducing the quantity of propellants needed to place the Mars ice sample into Earth orbit. The Purdue team found that this approach would have mass-saving knock-on effects throughout the MPISR mission design, yielding a 6% reduction in total spacecraft mass at Earth launch.
 
The Purdue students described an EOV with enough refrigerant on board to passively cool the ice sample for up 28 days in Earth orbit, meaning that the it would need to be retrieved no later than 30 November, 1988 if Earth-orbit capture took place on 2 November and no later than 12 December if it occurred on 14 November. During the 28-day period, an automated Tug would climb from low-Earth orbit to retrieve the EOV.

Following retrieval, the Tug would convey the sample container to a waiting Shuttle Orbiter or to a laboratory on board an Earth-orbiting space station. Concerns about planetary protection would drive the selection. If risk of terrestrial contamination were judged to be acceptable, then the Shuttle Orbiter would deorbit and transport the sample container to Earth's surface. If, on the other hand, a more conservative approach seemed warranted, then the Mars sample would be subjected to initial examination on the Station.

Purdue's MPISR concept generated considerable interest and demonstrated surprising longevity for a student project. After a summary of the study appeared in the pages of the British Interplanetary Society publication Spaceflight, two of its authors (Staehle and Skinner) briefed JPL engineers on the concept. They discussed the possibilities of "life indigenous to polar ice" on Mars and the significance of detection of "alternate chemistries" of life.

They also adjusted some of the dates of critical MPISR mission events. Departure from Earth orbit, arrival in Mars orbit, and Mars landing would take place as in the original study, but Mars surface liftoff would take place about two weeks early (24 April 1988). Subsequent dates were adjusted accordingly: the MPISR Orbiter/ERV/EOV combination would dock with the third stage of the AV on 9 May 1988; the ERV/EOV would depart Mars orbit on 20 July 1988; the EOV would capture into Earth orbit on 8 November 1988; and 6 December 1988 was the latest sample retrieval date.

In January 1978, JPL new-hire Staehle pitched the scientific benefits of the MPISR plan at the Lunar and Planetary Institute in Houston, Texas. In his presentation, he called "reasonable" the automated acquisition of an ice core up to 200 meters long made up of segments up to 15 millimeters in diameter.

Sources

"Mars Polar Ice Sample Return Mission - 1," Robert L. Staehle, Spaceflight, November 1976, pp. 383-390

"Mars Polar Ice Sample Return Mission, Part 2," Robert L. Staehle, Sheryl A. Fine, Andrew Roberts, Carl R. Schulenburg, and David L. Skinner, Spaceflight, November 1977, pp. 399-409

"Mars Polar Ice Sample Return Mission, Part 3," Robert L. Staehle, Sheryl A. Fine, Andrew Roberts, Carl R. Schulenburg, and David L. Skinner, Spaceflight, December 1977, pp. 441-445

Mars Polar Ice Sample Return Mission, R. Staehle and D. Skinner, Jet Propulsion Laboratory, September-October 1977

Mars Polar Ice Sample Return Mission - Overview, R, Staehle, Jet Propulsion Laboratory, January 1978

More Information






Chronology: Apollo X, Apollo Extension System, and Apollo Applications Program (AAP) 1.0

Repurposing Apollo: a modified Apollo Command and Service Module (CSM) (upper right) spacecraft moves through space docked with an Apollo Telescope Mount (ATM) derived from the Lunar Module (LM) lander. NASA's Apollo Applications Program (AAP) would have seen ATMs operating alone, with docked CSMs, and docked with AAP Orbital Workshops. Image credit: uncertain, but probably Grumman, makers of the LM.
This is the latest in a series of chronology posts in this blog. I usually write posts separately, with little regard for how they fit with others; these posts enable me to preserve that approach, which I find productive, while also linking separate posts to tell a larger story. This chronology post focuses on the Apollo Applications Program (AAP).

Begun formally in 1965, AAP grew from the Apollo Extension System, Apollo X, Manned Orbital Research Laboratory, and related proposals of the first half of the 1960s. Though backed by President Lyndon Baines Johnson, who saw it as a logical Apollo successor program, AAP suffered from repeated funding shortfalls and internal NASA squabbling.

The Apollo 1 fire (27 January 1967) took place within the main Apollo Program, but it was the final straw for AAP. The program was not formally ended, however, until the Skylab Program took over some of its Earth-orbital objectives in 1970.

AAP evolved into the three J-class Apollo missions (1971-1972) and four Skylab missions (1973-1974). Some have sought to portray the 1975 Apollo-Soyuz Test Project (ASTP) flight as an AAP successor; it is, however, better seen as a spinoff of Integrated Program Plan (IPP) space rescue planning. Nixon-era politics obscured ASTP's link with the IPP.

Apollo Extension System Flight Mission Assignment Plan (1965)

Relighting the FIRE: A 1966 Proposal for Piloted Interplanetary Mission Reentry Tests

Saturn-Apollo Applications: Combining Missions to Save Rockets, Spacecraft, and Money (1966)

"Without Hiatus": The Apollo Applications Program In June 1966

Apollo Applications Program: Lunar Module Relay Experiment Laboratory (1966)

"Assuming That Everything Goes Perfectly Well in the Apollo Program. . ." (1967)

Apollo Ends at Venus: A 1967 Proposal for Single-Launch Piloted Venus Flybys in 1972, 1973, and 1975

To "G" Or Not To "G" (1968)

A Forgotten Rocket: The Saturn IB

Rocket Belts and Rocket Chairs: Lunar Flying Units

As originally conceived, AAP would have included many lunar-orbital and lunar-surface missions. Pictured here is an LM-derived, automatically landed LM Shelter designed to support two astronauts during a surface stay lasting 14 days (one lunar daylight period). The astronauts would have arrived separately in an LM-derived "Taxi" spacecraft. Image credit: Grumman.

Chronology: Asteroids, Comets, and Other Small Bodies of the Solar System 1.0

1 Ceres is complex and shows signs of ongoing surface activity. Image credit: NASA.
Chronology is essential to understanding history, yet in this blog I write posts about planned space missions more or less at random, with little regard for the order in which they occurred. Because of this, I occasionally feel moved to publish omnibus chronological posts like this one. So far, I've applied the chronological treatment to groups of posts on Space Stations, catastrophic failure during space missions, missions to Venus, and the Apollo-to-Shuttle Transition

This post's topic is tied to Asteroid Day 2020. It establishes chronology for posts related to some of the Sun-orbiting small bodies of the Solar System: specifically, asteroids, comets, dwarf planets, and Kuiper Belt Objects (KBOs). In this introductory essay, I'll start with the largest members of these four broad classes. 

1 Ceres is an asteroid and a dwarf planet, much as 134340 Pluto is a KBO and a dwarf planet. Ceres, discovered on the first day of the 19th century, is the queen of the Main Belt between Mars and Jupiter, much as Pluto is the king of the Kuiper Belt, which begins just inside the orbit of Neptune. Clyde Tombaugh discovered Pluto on 18 February 1930, at Lowell Observatory.

Ceres is the largest and most massive asteroid. Pluto remains the largest known KBO, though new discoveries could nudge it from the top spot. Pluto is not the most massive Solar System body known beyond Neptune; that honor presently belongs to 136199 Eris, another KBO and dwarf planet, which for a time was thought to be larger than Pluto.

Ceres was not immediately classified as an asteroid when it was discovered. It was widely considered to be a planet until the 1850s, by which time new data — the discovery of more than a dozen other bodies orbiting with it between Mars and Jupiter — had made clear to everyone that it should be classified as the first known example of a new class of small Solar System body. Ceres pro forma became the first asteroid.

In similar fashion, Pluto was widely considered to be a planet until the early 2000s. Beginning in 1992, space scientists discovered that Pluto has siblings. This confirmed the existence of the long-hypothesized Kuiper Belt. The parallel with Ceres was not lost on scientists. Pluto became pro forma the first KBO.

In science, classification is fundamentally about clear communication, which is essential for collaborative research. Classification is not treated as a frivolous matter by most scientists. Only after sufficient data has been obtained, exchanged, and debated is an initial classification changed. 

Since the 1990s, scientific debate has taken place among space scientists via digital communication, enabling far more participation than in the past. The formal in-person poll that reclassified Pluto as a dwarf planet on 24 August 2006 included only a small percentage of the tens of thousands of space scientists scattered around the world; the matter of Pluto's classification had, however, already been widely debated. 

In fact, the vote marked the end of a 76-year-long scientific process. When first discovered, Pluto was assumed to have a mass about six times that of Earth. It had to be that massive to have enough gravitational pull to account for observed deviations in the orbit of Neptune, which is another story (you can read about it among the posts linked below). Pluto did not, however, show a disk, which implied that it was very dark, very dense, or both. 

Pluto's orbit also crossed that of Neptune, which made it unique among the planets. Planet-crossing is common among small bodies such as asteroids, but who ever heard of an asteroid with six times the mass of Earth?

Discovery in 1978 of Charon, Pluto's largest moon, enabled scientists to calculate Pluto's mass accurately for the first time. It has just one-fifth of 1% of Earth's mass, or less than 20% of the mass of Earth's Moon. They then determined Pluto's diameter; it measures less than three times the diameter of Ceres, or about two-thirds the diameter of Earth's Moon. It is astonishing that Tombaugh was able to spot Pluto using the crude astronomical tools available in 1930.

This is as good a place as any to express my view that the term "dwarf planet" should be retired. It is not especially useful to scientists, does not enhance public understanding so is worse than useless for science education, and appears to be moribund. Though perhaps a dozen KBOs discovered since 2006 appear to qualify for the label, none have been added to the initial list of five (in addition to the three I have already mentioned, they include Haumea and Makemake).

Asteroid exploration has advanced rapidly since the 1990s, in part because missions bound for other worlds often can find one or more asteroids to visit along their flight path. Galileo, bound for Jupiter orbit, became the first spacecraft to fly past an asteroid, 951 Gaspra, on 29 October 1991. Two years later, it flew past 243 Ida, in the process imaging Dactyl, the first asteroid moon to be found.

Dedicated asteroid missions began in February 1999 with a bit of a flub; the NEAR Shoemaker spacecraft suffered a computer glitch and missed its first opportunity to enter orbit about the near-Earth asteroid 433 Eros. A year later, NEAR Shoemaker fired its engines to slow itself so that Eros could capture it, making it the first asteroid orbiter. On 12 February 2001, it ended its mission with a bonus rough landing on Eros — the first asteroid landing.

The Dawn spacecraft entered orbit around 4 Vesta in July 2011, thus becoming the first spacecraft to orbit a Main Belt asteroid. It moved on to Ceres, achieving orbit around the largest asteroid in March 2015. 

2015 was a hot year for small-body exploration. NASA's New Horizons spacecraft performed a Pluto fast flyby in July of that year, making it the first spacecraft to visit a KBO. New Horizons flew past a second, smaller KBO, 486958 Arrokoth, in January 2019. Arrokoth is the most distant Solar System body yet explored by a spacecraft.

Dedicated comet missions began in 1985-1986, when a four-spacecraft European-Japanese-Soviet "armada" explorer 1P/Halley, the most famous of the comets. The spacecraft did not try to match orbits with Halley, which revolves around the Sun "backwards" relative to the planets; instead, they carried out fast flybys. In March 1986, Europe's Giotto spacecraft raced past Halley's dark nucleus at a relative velocity of 68 kilometers per second.

Europe's Rosetta spacecraft orbited 67P/Churyumov-Gerasimenko from August 2014 to September 2016. It was the first comet orbiter. Rosetta's time-at-target bracketed the comet's closest approach to the Sun, enabling unprecedented close-up observations of activity triggered by solar heating. Rosetta released the Philae lander on 12 November 2015; though it did not land properly, Philae returned images and other data from the surface for about three days.

An exciting new frontier in small body exploration is now opening. In October 2017, the first asteroid known to have originated outside the Solar System, 1I/'Oumuamua, was discovered. We know that it originated elsewhere in the Milky Way because it is moving too quickly for the Sun's gravity to do more than bend its course before it returns to interstellar space. The first interstellar comet, 2I/Borisov, was found in August 2019.

These new discoveries have inspired proposals for intercept missions. None has so far advanced to the point of serious consideration. Both bodies will, however, remain within range of expected human spaceflight technology for a few decades at least, and the list of known interstellar visitors seems likely to grow, providing new candidate star-roving small bodies for exploration.

The links below lead to posts related to small Solar System bodies dated from 1962 through 2005. In addition, three posts not firmly linked to specific years are included at the bottom of the list.

Pluto, Doorway to the Stars (1962)

To Mars by Way of Eros (1966)

Missions to Comet d'Arrest and Asteroid Eros in the 1970s (1966)

MIT Saves the World: Project Icarus (1967)

Things to Do During a Venus-Mars-Venus Piloted Flyby Mission (1967)

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

Multiple Asteroid Flyby Missions (1971)

Cometary Explorer (1973)

A 1974 Plan for the Slow Flyby of Comet Encke

Earth-Approaching Asteroids as Targets for Exploration (1978)

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

Visions of Spaceflight, c. 2001 (1984)

Catching Some Comet Dust: Giotto II (1985)

New Horizons II (2004-2005)

The Challenge of the Planets, Part Two: High Energy

The Challenge of the Planets, Part Three: Gravity

Pluto: An Alternate History

"Wobble Angle": Characteristics of 11 Apollo-derived Artificial-Gravity Space Station Designs (1963)

"Zero-G and I feel fine" — astronaut John Glenn, the first American to reach Earth orbit, during his five-hour flight on board Mercury-Atlas 6 spacecraft Friendship 7, 20 February 1962. Image credit: NASA.
In early May 1963, Robert Mason and William Ferguson, engineers at the NASA Manned Spacecraft Center (MSC) in Houston, Texas, completed a study of 11 artificial-gravity Earth-orbital laboratory designs. Some might have argued that NASA engineers had better things to do. After all, for two years the space agency's main goal had been to land a man on the Moon and return him safely to the Earth before the Soviet Union did, and the U.S. program still lagged behind its Soviet counterpart.

When the MSC engineers completed their study, the U.S. record for weightless space endurance was held by Wally Schirra, the third American to reach Earth orbit. During the Mercury-Atlas 8 mission (3 October 1962), he racked up a little less than nine hours of weightless experience. About a week after Mason and Ferguson completed their study, Gordon Cooper would set a new record by orbiting the Earth for about 34 hours during the Mercury-Atlas 9 mission (15-16 May 1963).

The world record for weightless space endurance at the time was, however, held by cosmonaut Andriyan Nikolayev, whose Vostok 3 spacecraft lifted off from Baikonur Cosmodrome on 11 August 1962. He orbited the Earth 64 times in 3 days, 22 hours, and 28 minutes, and landed on 15 August 1962. Apart from assurances that Nikolayev was in good health, the Soviet Union shared little information about his physical condition during or after his flight.

Lack of data on human responses to continuous weightlessness goes a long way toward explaining why NASA continued to study Earth-orbiting laboratories two years after President John F. Kennedy made the Moon a major U.S. goal on 25 May 1961. It seemed prudent to some to retain the option to launch a laboratory for studies of human health in weightlessness at least until astronauts could live in space for a period of time equal to the duration of an Apollo lunar landing mission.

Lack of data also explains why Mason and Ferguson studied artificial-gravity laboratory designs. If it were found that humans could not withstand weightlessness for long periods, then it would become necessary to establish a lab in space where the human health effects and engineering requirements of spin-induced acceleration — which is what "artificial gravity" is — could be examined.

There were also policy reasons for studying Earth-orbital laboratories. Before President Kennedy put NASA on course for the Moon, an Earth-orbiting lab had been central to the agency's plans for the 1960s. Some engineers believed that the laboratory should have remained NASA's first priority after Project Mercury, and they looked for opportunities to turn back the clock.

By the end of 1962, the probable cost of the lunar program had become increasingly clear. Grumbling had begun in Congress, placing pressure on Kennedy, who in turn placed pressure on NASA brass to contain space program costs. It seemed possible that the Apollo lunar goal might be found wanting by either Kennedy or, if he lost his bid for reelection in November 1964, by his successor. If so, the reasoning went, NASA might do well to have on hand a plan for an Apollo-derived Earth-orbiting laboratory as a cheap replacement for the lunar program.

In all but one of their 11 designs, Mason and Ferguson had the laboratory and crew reach orbit together; the astronauts would ride in a modified Apollo Command and Service Module (CSM) spacecraft atop the lab's drum-shaped Mission Module (MM). CSM modifications included a much-shortened Service Module (SM) with only enough propulsion, power, and life-support capability for the trip to the lab's 300-mile-high operational orbit and return to Earth.

Mason and Ferguson focused their study on the extent of the shift in the laboratory spin axis that astronaut movement parallel to the spin axis would produce. They called that shift the "wobble angle."

This illustration from Mason and Ferguson's paper depicts the "wobble angle." The line marked "Z" corresponds to the spin axis, which passes through the center of gravity of the orbiting laboratory. The Z at the top would, if the laboratory's spin remained entirely stable, always point directly at the Sun. Astronaut movement parallel to the Z line would, however, cause the spin axis to shift along the curving line labeled "Spin-axis trace." In this design, which corresponds to Laboratory Design 1 below, astronauts would need to contend with a wobble angle of up to 43°. Mason and Ferguson likened this motion to the "rolling of a ship."
The MSC engineers assumed that the orbiting laboratory MM and other structure, habitation and science equipment, and the modified CSM would together weigh about 15 tons. Of that, five tons were allotted to the CSM. All of their designs retained the Saturn IB rocket second stage, the S-IVB, for use as a counterweight. With its liquid hydrogen/liquid oxygen propellants spent, the S-IVB stage would weigh 10 tons.

Mason and Ferguson set the spin rate at a maximum of four rotations per minute. At that rate, and at a distance of 40 feet from the spin axis, the acceleration an astronaut would feel would vary by 15% between their feet and their head, with maximum acceleration being felt at their feet, farthest from the spin axis. Maximum acceleration would be limited to one Earth gravity; minimum acceleration would not fall below one lunar gravity (0.2 Earth gravities).

The 11 images that follow each include two views. The laboratory launch configuration is on the left and orbital configuration is on the right. In all but two of the images, the Z axis/spin axis points at the viewer in both views; for Laboratory Designs 8 and 9, the Z axis in the launch configuration view is turned 90° relative to the orbital configuration view.

Laboratory Design 1: the first Mason and Ferguson artificial-gravity lab design is the simplest, though it also has one of the greatest maximum wobble angles (about 43°). Crew couches in the CSM are at the minimum distance (40 feet) from the spin axis (Z), but the entire two-deck MM is too near the spin axis to avoid a variation in acceleration level between astronaut head and feet greater than 15%. Equipment weight is 12,496 pounds, structure weight is 7504 pounds, and pressurized volume is 2504 cubic feet. Thrusters located at the ends of the lab would expend 52.9 pounds of propellant to start it spinning at a rate of four rotations per minute.
Laboratory Design 2:  An alternate method of solar array deployment improves stability (wobble angle slightly more than 9°) by increasing lab width and mass along the Y axis. Structure weight is 8235 pounds and equipment weight is 11,765 pounds. Pressurized volume is 2505 cubic feet. Unfortunately, no part of the CSM or MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Thrusters expend 55.3 pounds of propellant to spin up the laboratory. 
Laboratory Design 3:  Equipment modules of unspecified function deploy along the Y axis; this helps to reduce maximum wobble angle to about 3.5°. Structure weight including the equipment modules is 12,492 pounds. Equipment weight — 7508 pounds — is the least of any of the designs. Pressurized volume is 2396 cubic feet. No part of the CSM or MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feetThrusters expend just 49.7 pounds of spin-up propellant. 
Laboratory Design 4: A tunnel between the CSM and the MM places the CSM crew couches 43.3 feet from the spin axis. Unfortunately, the maximum distance from the spin axis within the MM is just 18.3 feet. Placing the relatively massive CSM far from the spin axis and relatively narrow structure along the Y axis contribute to a wobble angle of nearly 44°. Structure weight is 8687 pounds and equipment weight is 11,313 pounds. Pressurized volume is 2396 cubic feet. Thrusters expend 50.7 pounds of propellant to spin up the laboratory. 
Laboratory Design 5: The tunnel linking the CSM and MM is extendable, increasing CSM crew couch distance from the spin axis to 52.9 feet. The wobble angle is identical to that of Design 4. Structure weight is 8290 pounds and equipment weight is 11,710 pounds. Pressurized volume is 2400 cubic feet. The MM entirely surrounds the spin axis; in theory, an astronaut at the spin axis would be weightless while the station spun around them. No part of the MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Thrusters expend 65.7 pounds of propellant to spin up the laboratory. 
Laboratory Design 6: Both the CSM and the MM telescope away from the spin axis. The 45° maximum wobble angle is the greatest of the 11 designs. Structure weight is 7505 pounds and equipment weight is 11,765 pounds. Pressurized volume is just 1633 cubic feet, the least of any of the designs. About half the MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Thrusters expend 68.3 pounds of propellant to spin up the lab. 
Laboratory Design 7 is similar to Design 6, but its modified solar array configuration increases its width and mass along the Y axis, reducing its maximum wobble angle to slightly less than 29°. Structure weight is 7869 pounds, equipment weight is 12,131 pounds, and pressurized volume is 1743 cubic feet. Thrusters expend 68.3 pounds of propellant to spin up the laboratory.
Laboratory Design 8 combines features of Designs 3 and 7 to achieve a wobble angle of slightly less than 2.5°. A new feature of this design is a docking porfor a visiting modified CSM at the spin axis (Z). In many artificial-gravity station designs, docking ports at the spin axis rotate spin "backwards" so that they appear to remain still, facilitating docking. Mason and Ferguson gave no indication that their design would include a counter-spun docking port, however. Structure weight is 12,169 pounds, equipment weight is only 7831 pounds, and pressurized volume — without a second CSM — is 2048 cubic feet. All of the CSM and nearly all of the MM are far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Thrusters expend 66.9 pounds of propellant to spin up this design. 
Laboratory Design 9 includes new structural elements: a "fork" and cables that permit the spent S-IVB stage to be pivoted 90° relative to its launch axis. This reduces the wobble angle to slightly less than 1° — the least of any of the 11 designs. Unfortunately, no part of the CSM or MM is far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Structure weight is 8306 pounds and equipment weight, 11,694 pounds. Pressurized volume is 3118 cubic feet. Thrusters expend 64 pounds of spin-up propellants.
Laboratory Design 10 employs a "rigid support" and cables to pivot the spent S-IVB stage 90° relative to its launch axis. Maximum wobble angle is 1°. The CSM crew couches and part of the MM are far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. Structure weight is 8120 pounds and equipment weight is 11,880 pounds. Pressurized volume is 2400 cubic feet. Thrusters would expend 71.8 pounds of propellants to spin up this design.
Laboratory Design 11 includes no CSM in its launch configuration view because structure and equipment weight is too great. The large MM is extendible. The CSM is displayed in the orbital configuration view as it would appear after it launched separately and docked with the MM in orbit. The pivoted S-IVB stage and the solar panel arrangement help to compensate for the large MM, yielding a wobble angle only slightly greater than Design 9. Structure weight is 14,047 pounds and equipment weight is 15,953 pounds. Pressurized volume is by far the greatest of the 11 designs (3828 cubic feet), as is the amount of spin-up propellant required (116.1 pounds). Spin-up would take place after the CSM arrived. All parts of CSM and MM are far enough from the spin axis to avoid a greater than 15% variation in acceleration level between astronaut head and feet. 
Source

Project Apollo Conceptual Rotating Space Vehicle Designs Using Apollo Components for Simulation of Artificial Gravity, NASA Project Apollo Working Paper No. 1073, NASA Manned Spacecraft Center, 8 May 1963.

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