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)

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 portion of the adjoining seasonal temporary cap. Many features of the southern hemisphere are 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 rim 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 1% 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 ice cap of Mars in southern hemisphere summer. In winter, the entire image would be cloaked in red dust and carbon dioxide frost and ice. 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 the Viking Orbiter with attached Viking Lander aeroshell capsule. Image credit: NASA.

Schematic of 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 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 spacecraft 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.

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

Bridging the 1970s: Lunar Viking (1970)

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

Mars Airplane (1978)

Making Propellants from Martian Air (1978)

Catching Some Comet Dust: Giotto II (1985)

Safeguarding the Earth from Martians: The Antaeus Report (1978-1981)

Mars Airplane (1978)

Wings over Mars: the JPL Mars airplane swoops past a martian mountain so that its camera, mounted inside a clear plastic bubble on its belly, can turn sideways to image layers on the mountain slopes. Image credit: Jeff Bateman.

In the 1970s, as U.S. piloted spaceflight retreated to low-Earth orbit, NASA planning for advanced robotic Mars exploration missions came into its own. New information on the martian environment from Mariner 9 and the twin Vikings fueled engineer imaginations.

Many concepts that became actual missions in the 1990s and 2000s first received detailed study in the 1970s. Planners also looked at concepts that have yet to yield NASA missions: Mars sample return, balloons and blimps, small lander networks, and gliders and powered fixed-wing aircraft.

Spacecraft and space mission design and development decisions are complex and influenced by many factors. Scientific efficacy is but one factor considered when planning for a new exploration mission begins, and it is not always the most important one. Nevertheless, scientists are almost always involved at the outset even when they do not originate the mission concept under consideration. Often this involvement is achieved through the establishment of a supportive science working group for the proposed mission.

The Ad Hoc Mars Airplane Science Working Group met at the Jet Propulsion Laboratory (JPL) in Pasadena, California, on 8-9 May 1978, to review mission objectives and propose a possible Mars airplane instrument payload weighing between 40 and 100 kilograms. In its report, the Group noted that a Mars Airplane designed for landings and takeoffs would be able to collect samples in places other types of vehicles might find hard to reach. The plane might also be used to deploy small payloads at scattered locations by airdrop or landing.

Mostly, however, the Ad Hoc Science Working Group limited its deliberations to use of the plane as an aerial survey platform. The Group based its planning on a Mars airplane design derived from NASA Dryden Flight Research Center's "MiniSniffer" pilotless plane, which was designed to sample Earth's stratosphere.

The 300-kilogram airplane would arrive at Mars folded in an lozenge-shaped Viking-type aeroshell. After aeroshell parachute deployment and heat shield separation, it would spread its hinged wings to their full 21-meter span and detach from the parachute and aeroshell in mid-air.

Conceptual Mars airplane design. Image credit: Jeff Bateman.

Normally, the plane would cruise one kilometer above the martian surface, though it would be capable of flying as high as 7.5 kilometers. The 4.5-meter-diameter propeller at the front of its 6.35-meter-long fuselage would pull it through the thin (less than 1% of Earth atmosphere density) martian atmosphere at a speed of between 216 and 324 kilometers per hour.

Mars airplane endurance would depend on the weight of its payload and the choice of power plant. A plane with a 13-kilogram, 15-horsepower hydrazine-fueled piston motor, 187 kilograms of hydrazine fuel, and a 100-kilogram payload could, the Group estimated, fly up to 3000 kilometers in 7.5 hours, while one with a 20-kilogram electric motor, 180 kilograms of advanced lightweight batteries, and a 40-kilogram payload could fly up to 10,000 kilometers in 31 hours.

After it depleted its fuel or batteries, the plane would crash on Mars. The Group noted that the plane's short operational lifetime would dictate that its position after atmosphere entry be determined rapidly so that it could be directed quickly to its survey targets.

The Ad Hoc Group assumed that the Mars airplane would carry an inertial guidance system, radar and atmospheric-pressure altimeters, and terrain-following sensors (laser or radar) for navigation, and that these would serve double-duty as science instruments. The Group's selected science payload was intended to characterize possible landing sites for a follow-on Mars sample return mission and also to perform "topical" studies. The latter would address specific questions about Mars: for example, "Is Valles Marineris a rift valley?"

Visual imaging would be "fundamental" to the Mars airplane mission, so would receive top priority in the instrument suite. The Group determined that the airplane would be well-suited to serve as a camera platform because it would offer image resolution intermediate between orbiter and lander cameras and would obtain valuable "oblique" (from the side) images of the surface.

A Mars airplane might fly down a sinuous martian outflow channel, for example, collecting high-resolution images of layers exposed in its walls. The Mars airplane camera might be mounted on a movable platform inside a transparent dome on the plane's belly.

Other high-priority investigations would include wind speed, air pressure, and temperature measurements at various altitudes, infrared and gamma-ray spectroscopy and multispectral imaging to determine surface composition, and measurements of local magnetic fields. For magnetic field studies, the plane would fly a grid pattern over a selected region. The magnetometer, which might be mounted on a boom or a wingtip to minimize interference from airplane electrical sources, could also be used to seek out iron-rich surface materials and buried iron-rich volcanic structures.

The 1978 Mars airplane conceptual design effort fell victim to post-Viking disenchantment with Mars. Viking, which cost more than $1 billion in 1975 dollars, had been intended to find life, but its three biology experiments did not produce an unequivocally positive result. The Mars community did not at first recognize that it would need to restore support for Mars exploration before it proposed new Mars missions; that is, that Viking had made it more difficult to sell Mars exploration, not easier.

In addition, Space Shuttle development experienced setbacks. It was difficult to justify development of a vehicle for flying in the thin atmosphere of Mars when NASA had difficulty building one to fly in the thin upper atmosphere (and thicker lower atmosphere) of Earth.

Mars missions would resume, but not until 1992, when NASA launched a sophisticated orbiter called Mars Observer. The spacecraft was meant to inaugurate a new era of Mars exploration by providing a new overview of the planet. The loss of Mars Observer as it neared its destination on 25 September 1993 was a major setback; for a time, it appeared that recriminations over the very public failure might halt NASA Mars exploration.

The Curiosity rover landed in Gale Crater on 6 August 2012 and, after a checkout period, began its slow climb up the geologically complex layered slopes of Aeolus Mons (seen here in a color-corrected montage of images captured on 9 September 2015). At this writing, six-wheeled Curiosity has traveled about 22 kilometers. A Mars airplane could provide a perspective on Aeolus Mons, Valles Marineris, and other large features of Mars intermediate between that of a rover and that of an orbiter. Image credit: NASA.

Sources

Final Report of the Ad Hoc Mars Airplane Science Working Group, JPL Publication 78-89, NASA Jet Propulsion Laboratory, 1 November 1978.

Mars Airplane Presentation Material Presented at NASA Headquarters, JPL 760-198, Part II, Jet Propulsion Laboratory, 9 March 1978.

More Information

The Russians are Roving! The Russians are Roving! A 1970 JPL Plan for a 1979 Mars Rover

After Venus: Pioneer Mars Orbiter with Penetrators (1974)

Purple Pigeon: Mars Multi-Rover Mission (1977)

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

Making Propellants from Martian Air (1978)

Flying Brickyard Postponed: A 1972-1973 Study of an Interim Ablative Space Shuttle Heat Shield

Space Shuttle Orbiter during atmosphere reentry as conceived by North American Rockwell (NAR) in July 1972. The following month, NASA would make NAR the Space Shuttle prime contractor. Image credit: NASA.

Launch, ascent to orbit, and Earth atmosphere reentry are the most risk-fraught phases of most piloted space missions to date. They are also the mission phases that most tax the ingenuity of engineers who design reusable spacecraft.

Aerodynamic heating creates challenges during reentry and, to a lesser degree, during ascent to orbit. Before the Space Shuttle, almost all piloted spacecraft designed to operate for some portion of their mission in an atmosphere withstood such heating by employing single-use ablative heat shields. (The only exception was the X-15A-2 rocket plane, which, for part of its career, included a replaceable ablative heat shield — please see "More Information" at the end of this post.) During reentry, ablative heat shields char and break away, carrying away heat.

The Space Shuttle, approved for development by President Richard Nixon on 5 January 1972, marked a dramatic departure in heat shield technology. Originally conceived as a fully reusable, economical Space Station resupply and crew rotation vehicle, Nixon's partially reusable Shuttle had as its only approved goal a dramatic reduction in the cost of launching things into space. A reusable heat shield was believed to be essential for achieving that objective.

Over the decades, engineers have considered many reusable heat shield concepts, typically in combination. High on the list was a layer of overlapping "shingles" made of exotic metal alloys. Other approaches included liquid or solid heat sinks, thick metal or composite adjoining plates, or even an "active" system with cooling fluid circulating through a network of tubes behind a metal-alloy hull.

Unfortunately, all of these concepts would be heavy. To compensate for a heavy heat shield, engineers could design a more powerful booster system or could cut back on payload capacity (or both). Both approaches would boost development and operations costs. The Nixon White House had made clear that the Shuttle development budget of $5.15 billion was carved in stone, leaving NASA with little choice but to find new approaches — including some that accepted a significant increase in eventual operations cost.

Partial cutaway of the NAR ("NR") Shuttle Orbiter showing optional jet engine module installed in the aft portion of the Payload Bay, main engines, fore and aft reaction control system thruster clusters, aluminum wing structure, twin robot arms, crew module, and ring-shaped, nose-mounted docking unit forward of the crew module. Image credit: NASA.

For the Shuttle Orbiter, NASA and contractor engineers chose a lightweight combination of fabrics and brittle silica ceramic tiles, which they dubbed Reusable Surface Insulation (RSI). The tiles could withstand temperatures of up to 2300° Fahrenheit. Reinforced Carbon-Carbon composite panels would protect the Orbiter's wing leading edges, nose, and other areas subject to the highest reentry temperatures (as high as 3000° Fahrenheit).

Though RSI was meant to block almost all heat, enough would get through that, combined with aerodynamic buffeting, the Orbiter's mostly aluminum skin would tend to warp and flex ("flutter"). This meant that large ceramic panels affixed to the skin would crack, leaving it vulnerable to reentry heating.

Shuttle engineers sought to avoid damage by gluing RSI ceramics to a flexible fabric "strain isolator" layer glued to the Orbiter's skin and by making individual ceramic elements small in size. By resorting to many small "tiles" in place of a relatively few large panels, engineers designed an RSI heat shield that was in effect "pre-cracked."

The tiles, each milled to conform to its place on the Orbiter's complexly curvaceous hull, would number in the tens of thousands. By late in the 1970s decade, when their number hovered around 31,000, the tiles earned the Orbiter the nickname "The Flying Brickyard."

Some engineers harbored doubts about RSI; enough that NASA Langley Research Center in Hampton, Virginia, paid the Denver Division of Martin Marietta Corporation (MMC) to examine an alternative. Between May 1972 and August 1973, MMC engineers sought to determine whether Space Shuttle Orbiters could employ an ablative heat shield.

The ablative shield was seen as a stand-in system meant to provide NASA with more time for RSI development should problems arise. In his October 1975 report on the ablative heat shield study, Rolf Seiferth, who managed the MMC study between 5 September 1972 and its conclusion on 31 August 1973, envisioned that the ablative shield might fill in for RSI for five years. Based on a November 1972 NASA-generated Space Shuttle traffic model, this meant that 151 flights between 1979 and the end of 1983 would rely on the stand-in ablative system.

Seiferth noted that, in past programs, ablative heat shield materials had been glued directly to the spacecraft hull. This was, he explained, a cost-saving, weight-saving approach; scraping away a used directly applied ablative shield would, however, add time to Orbiter refurbishment between flights and generate considerable debris, including invasive dust.

In addition to the directly applied heat shield, MMC examined three types of "mechanically attached" ablative panels. These had ablative material glued to panels made of aluminum, magnesium, graphite composite, or beryllium/aluminum "Lockalloy" sheet or honeycomb.

The panels would be joined to oversized holes in the Orbiter's skin using nut-and-bolt fasteners, enabling entire panels to be replaced as necessary. The oversized holes would allow for thermal expansion of the heat shield components.

The simplest mechanically attached ablative panel would see ablative material glued to a metal or composite sheet. Adhesive and sheet would together measure only about 0.06 inches thick. Attachment points for the sheet panel design would typically occur five inches apart over much of the Orbiter, though larger spacings (up to 20 inches) were also possible.

The two more complex mechanically attached ablative panels substituted metal or composite "honeycomb" for the metal or composite sheet. One had ablative material glued to the honeycomb, which was then bolted to oversized holes in the Orbiter's skin.

The other — to which MMC gave considerably less attention — added rib-like standoffs to the Orbiter's skin. The honeycomb was then mechanically attached to oversized holes in the standoffs, leaving a gap between the underside of the honeycomb and the Orbiter skin.

Honeycomb panel attachment points would typically occur 10 inches apart over much of the Orbiter. Larger (up to 20 inches) and smaller (down to five inches) spacings were possible.

Seiferth's team used computer models to determine required ablator thickness, which would vary depending on its location on the Orbiter. All models assumed a maximum reentry deceleration equal to 2.5 times Earth's surface gravity (that is, 2.5 G) and a maximum allowable Orbiter aluminum skin temperature of 350° Fahrenheit, variables which indicated a relatively benign reentry environment (as compared to an Apollo lunar-return reentry, for example).

MMC used for its calculations properties of several types of ablative material it had developed for other missile and space projects (notably, the Titan missile family and the Viking Mars lander). It found that, for most locations on the Orbiter, its least robust ablator would be sufficient.

The ablative layer for most locations could be surprisingly thin. For the simplest mechanically attached panel design, for example, the MMC computer models indicated that a point on the Orbiter's underside on the fuselage centerline 50 feet aft of its nose would need a layer of ablative material only 1.7 inches thick.

Assessing the cost of the ablative designs relative to RSI was difficult in part because Space Shuttle Program cost estimation was, for want of a better term, eccentric. Seiferth supplied no development or operations cost estimate for RSI in his report, though he did provide estimates for several of MMC's ablative designs.

A system with an ablator glued directly to the Orbiter's aluminum skin would, Seiferth estimated, cost a total of $164.8 million for 151 flights over five years. Of this, installation and removal would account for $27.9 million.

A mechanically attached system comprising an aluminum sheet, adhesive, and an ablator (that is, the simplest mechanically attached ablative system) with attachment points five inches apart would cost $168.3 million with an installation and removal cost of $21.9 million. The aluminum honeycomb system with no standoffs and attachment points five inches apart came in at $187.1 million with $25.7 million for installation and removal.

NASA provided MMC with an RSI weight estimate of 30,240 pounds, enabling an RSI/ablative system weight comparison. The MMC study determined that an ablator directly attached to the Orbiter's skin would weigh 27,199 pounds, while the sheet and honeycomb (no standoffs) mechanically attached systems would weigh 32,577 pounds and 32,158 pounds, respectively.

Seiferth noted that modifications to the Orbiter's aluminum skin design would need to be put in place soon if mechanically attached ablative panels were used. Delaying until after the Orbiter's skin was in place would make prohibitive the cost and difficulty of adopting the ablative Space Shuttle heat shield. By the time Seiferth's report saw print in October 1975 — more than two years after the MMC study concluded — a stand-in ablative heat shield, never high on NASA's list of Space Shuttle priorities, was in fact no longer an option.

Late in the 1970s decade, problems with the Space Shuttle Main Engine, RSI, computers, and other systems contributed to delays in STS-1, the Space Shuttle's orbital maiden flight. RSI problems in particular became very public in March 1979, when the Space Shuttle Orbiter Columbia was flown from California to NASA Kennedy Space Center (KSC), Florida, atop its 747 carrier aircraft. It was the first Orbiter's first visit to its home base. At the time, Columbia was scheduled to carry out STS-1 in November 1979.

Columbia rolls into the Orbiter Processing Facility at Kennedy Space Center, Florida, on 25 March 1979. Though the image displays only the area around the front of the fuselage, many RSI gaps are evident. Image credit: NASA.

For the cross-country flight, about 26,000 permanent RSI tiles were installed on Columbia, along with about 5000 foam "dummy" tiles. By the time the Orbiter/747 combination set down on the Shuttle Landing Facility strip at KSC on 25 March 1979, Columbia had lost more than 200 RSI tiles. Many were lost as more than 4800 of the dummy tiles tore loose, a condition which would not occur during space flight.

Some permanent RSI tiles had, however, fallen off Columbia for other reasons. Close examination revealed tile manufacturing flaws, installation errors, and an overall unexpected degree of fragility. Even as Columbia entered the processing flow for STS-1, NASA conceded that the flight might be delayed until 1980.

Much was made of the "zipper effect," a hypothetical catastrophic failure mode that would begin with the loss of a single tile during reentry. The Orbiter was believed likely to survive loss of a single tile unless it occurred in an especially critical area. Loss of a single tile anywhere would, however, weaken surrounding tiles, potentially leading to a cascading loss of thermal protection. In fact, few tiles fell off Orbiters during the series of 135 Shuttle missions that began with Columbia's first launch on 12 April 1981.

The RSI system did, however, prove prone to impact damage during processing, launch, landing, and transport. The most extreme example before January 2003 occurred during STS-27 (2-6 December 1988), a classified Department of Defense mission. Eighty-five seconds after liftoff, debris broke free from the right Solid Rocket Booster, battering the right wing of Orbiter Atlantis. More than 700 RSI tiles were damaged and one was lost. Because the mission was classified, the near-disaster was not widely known for nearly 20 years.

This closeup of the right wing of the Orbiter Discovery was taken from the International Space Station (ISS) during STS-114 (26 July-9 August 2005), the first post-Columbia "Return-to-Flight" Mission. After the Columbia accident, NASA modified the External Tank design to eliminate the possibility of debris separation; nevertheless, two pieces of icy foam insulation broke free during STS-114, with one striking Discovery. In addition to a tile repair kit, which the STS-114 crew tested during a scheduled spacewalk, Discovery carried a Shuttle Remote Manipulator ("robot arm") extension that enabled its crew to inspect its RSI surfaces; it also performed a slow flip near the ISS so that astronauts on the station could inspect and photograph it. Though no damage was found, NASA prudently grounded the Shuttle fleet for another year after STS-114 returned to Earth so that it could continue its efforts to solve the External Tank debris problem. Image credit: NASA.

The Space Shuttle Orbiter Columbia lifted off on 16 January 2003 at the beginning of mission STS-107, its 28th flight and one of the few remaining non-ISS missions NASA had scheduled for the Shuttle fleet. During ascent, a piece of water ice-impregnated insulating foam weighing almost two pounds broke free from the External Tank to which Columbia was mounted. It struck the Reinforced Carbon-Carbon leading edge of the Orbiter's left wing, punching a hole at least 10 inches wide.

The debris strike was captured on video and immediately became the subject of urgent debate within the Shuttle Program. Knowledge of the strike was not shared widely. The viewing angle meant that the strike area was not visible in launch video recorded from the ground and its location meant that the STS-107 crew could not see it. Managers decided that Columbia's wing leading edge was probably intact.

The hole admitted hot gas as Columbia reentered on 1 February 2003. Its internal structure compromised, NASA's oldest Orbiter broke up over east Texas and western Louisiana, killing its seven-person crew and grounding the Space Shuttle fleet for 30 months.

The following January, President George W. Bush declared that the Space Shuttle would be retired after it performed its last International Space Station (ISS) assembly mission. The final Shuttle flight, STS-135 (8-21 July 2011), saw Atlantis, veteran of the STS-27 near miss, deliver supplies to ISS ahead of an anticipated gap in U.S. piloted space flights of indefinite duration.

Sources

"Space Shuttle Orbiter and Subsystems," D. Whitman, Rockwell International Corporation; paper presented at the 11th Space Congress in Cocoa Beach, Florida, 17-19 April 1974.

Ablative Heat Shield Design for Space Shuttle, NASA CR-2579, R. Seiferth, Denver Division, Martin Marietta Corporation, October 1975.

"Thermal Tile Production Ready to Roll," R. O'Lone, Aviation Week & Space Technology, 8 November 1976, pp. 51, 53-54.

"First Orbiter Ready for Florida Transfer," B. Smith, Aviation Week & Space Technology, 5 March 1979, pp. 22-23.

"Thermal Tile Application Accelerated," C. Covault, Aviation Week & Space Technology, 21 May 1979, pp. 59, 61-63.

"Space Shuttle Orbiter Status April 1980," S. Jones, NASA Johnson Space Center; paper presented at the 17th Space Congress in Cocoa Beach, Florida, 30 April-2 May 1980.

STS-27R OV-104 Orbiter TPS Damage Review Team, Volume I, Summary Report, NASA TM-100355, February 1989.

More Information

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

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

What If a Space Shuttle Orbiter Had to Ditch? (1975)

What If a Space Shuttle Orbiter Struck a Bird? (1988)

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)

Bridging the 1970s: Lunar Viking (1970)

NASA's lunar soft-landers: in the background, the Apollo 12 Lunar Module Intrepid; in the foreground with Apollo 12 Commander Charles Conrad, Surveyor 3. Image credit: NASA.

In the 1960s, U.S. space assets included two spacecraft designed to soft-land on the Moon. These were automated three-legged Surveyor, of which seven were launched on Atlas-Centaur rockets between June 1966 and January 1968 (five Surveyors landed successfully), and the piloted four-legged Apollo Lunar Module (LM), which landed at six sites between July 1969 and December 1972.

Even as Surveyor 7 successfully soft-landed near the great ray crater Tycho, NASA, science advisory groups, Congress, and President Lyndon Baines Johnson considered plans for a project to soft-land spacecraft on Mars. Originally conceived in late 1967/early 1968 as "Titan Mars 1973," Project Viking, as it became known, received new-start funding in the Fiscal Year (FY) 1969 budget.

NASA's Langley Research Center (LaRC) managed Viking. LaRC, located in Hampton, Virginia, contracted with Martin Marietta in Denver, Colorado, to build two new-design Viking Landers. Meanwhile, the Jet Propulsion Laboratory (JPL) in Pasadena, California, began work on two Viking Orbiters based on its Mariner flyby spacecraft design first flown in 1962. The twin Viking spacecraft would each comprise a Lander and an Orbiter, and each Lander-Orbiter combination would leave Earth atop a Titan rocket with a Centaur upper stage.

NASA at first planned to launch the Vikings in July 1973, when an opportunity for a minimum-energy Earth-Mars transfer would occur. In January 1970, however, tight funding planned for FY 1971 forced a slip to the August-September 1975 minimum-energy Earth-Mars transfer opportunity.

For NASA's piloted space program, 1970 was eventful even though only a single mission took place. The mission, Apollo 13 (11-17 April 1970), was intended to build on the experience gained through the Apollo 11 (16-24 July 1969) and Apollo 12 (14-24 November 1969) landings. The Apollo 11 LM Eagle landed long, but the Apollo 12 LM Intrepid set down close by derelict Surveyor 3 on the Ocean of Storms, demonstrating that the LM could successfully reach a predetermined target.

Landing accuracy was important for planning geologic traverses, the first of which was to have taken place at Fra Mauro during Apollo 13. An explosion in the Service Module of the Apollo 13 Command and Service Module (CSM) Odyssey scrubbed the landing and put off the first lunar geologic traverse to Apollo 14 (31 January-9 February 1971), which also was directed to Fra Mauro.

The Apollo 13 accident and postponement of subsequent missions meant that much of the activity in NASA's piloted program in 1970 concerned planning and budgets. President Richard Nixon saw no cause for a large-scale Apollo-type goal in the 1970s; NASA Administrator Thomas Paine begged to differ. Nixon appointed the Space Task Group (STG) in February 1969 — less than a month after his inauguration — and made his Vice President, Spiro Agnew, its chair. Paine, a Washington neophyte, misjudged Agnew's importance in the Nixon White House, so believed that he had scored big when Agnew declared at the Apollo 11 launch that he believed NASA should put a man on Mars before the end of the 20th century.

Paine took Agnew's statement as an endorsement of the Integrated Program Plan (IPP), NASA's proposal for its future after Apollo. The IPP included a large Earth-orbital "Space Base," nuclear rockets, lunar orbital and surface bases, a piloted Mars landing mission, and Mars orbital and surface bases. At Paine's insistence, the STG's September 1969 report The Post-Apollo Space Program: Directions for the Future offered the White House only the IPP with three different timetables for carrying it out. Nixon's aides, more cognizant of their boss's thoughts on spaceflight, added an introduction outlining a future with no major goals and no target dates.

This NASA Marshall Space Flight Center illustration from 1970 displays Integrated Program Plan hardware elements planned to be operational in the 1990s. 

Paine largely ignored this clear message, instead focusing his efforts on making a permanent Earth-orbiting Space Station NASA's 1970s goal. In addition to a host of Earth-focused uses, the Station would permit astronauts to live and work in space for long periods. This would enable aerospace physicians to certify that humans could remain in space long enough to reach and return from Mars, a voyage that might last three years. A reusable piloted logistics resupply & crew rotation spacecraft — a Space Shuttle — would economically service the Station.

Paine expected that NASA would use a two-stage version of the Saturn V rocket to launch the core Station and other large IPP hardware elements. In January 1970, however, he found himself obliged to announce that Saturn V production would end with the fifteenth rocket in the series. Apollo missions through Apollo 19 would occur at six-month intervals, ending in 1974, and Apollo 20 would be canceled so that its Saturn V, the last of the original Apollo buy, could launch the Skylab Orbital Workshop. Skylab was the last remnant of President Johnson's post-Apollo piloted program, the Apollo Applications Program (AAP), which aimed to apply successful Apollo technology to new space goals; that is, to squeeze the U.S. investment in Apollo for all it was worth.

NASA advance planning developed a split personality in 1970. Some planners assumed that Saturn V rockets would be available indefinitely; others, that the Space Shuttle would launch all IPP hardware.

For example, even as Paine announced the end of Saturn V production, NASA piloted spaceflight planners studied a versatile reusable chemical-propellant Space Tug which could double as a Saturn V fourth stage. As early as 1980, a four-stage Saturn V would launch a Lunar Orbit Space Station (LOSS). The Saturn V S-IVB third stage would boost the LOSS/Space Tug toward the Moon and detach; the Space Tug would then correct the LOSS's course en route to the Moon and slow it so that the Moon's gravity could capture it into lunar orbit.

Subsequent Saturn V missions would build up a propellant farm and fleet of Space Tugs in lunar orbit. Astronauts in Space Tugs with crew cabins and landing legs would then descend from the LOSS to resume piloted lunar surface exploration and build a Lunar Surface Base (LSB).

Space Tug outfitted for piloted lunar landings. Image credit: NASA.

In June 1970, five planners with Bellcomm, the NASA Headquarters planning contractor, completed a multi-part memorandum in which they bemoaned the "prolonged gap in the lunar program. . .of at least six years" that NASA's Space Tug/LOSS/LSB plans would create. They argued that the gap would threaten the multidisciplinary community of lunar scientists Apollo and its robotic precursors had created. The gap also meant that Apollo exploration would make discoveries that could not be followed up until at least 1980. Construction of the LSB could not proceed immediately after the LOSS was established; piloted Space Tug missions to check out prospective LSB sites would need to take place first.

The Bellcomm team proposed a novel method of filling the gap after Apollo 19 and hastening construction of the LSB. They sought to repurpose spacecraft designs expected to become available in 1975: namely, the robotic Orbiter and Lander of the Viking Mars exploration program.

At the time they wrote, neither the Viking Orbiter nor Viking Lander designs were final. The Lander, for example, would eventually carry three biology experiments and two scanning cameras, but the Bellcomm team assumed only two biology experiments and one camera. They saw this as an advantage, for it meant that the Mars Viking design was not so far along that it could not to some degree take into account anticipated Lunar Viking needs.

Lunar Viking Lander. The design depicted includes a pair of scanning cameras.  Image credit: NASA/Russell Arasmith.

The most obvious modification to the Mars Viking design for lunar missions would be replacement of the Lander aeroshell, heat shield, and parachutes with a solid-propellant landing rocket. The Lunar Viking Orbiter would expend liquid propellants to slow itself and the Lunar Viking Lander so that the Moon's gravity could capture the combination into lunar orbit, then would perform maneuvers to adjust its orbit ahead of Lander release. The Lander would then detach and, at the proper time for a landing at its target site, ignite the solid-propellant rocket.

After its propellant was expended, the motor casing would fall away. The Lunar Viking Lander would then complete descent and soft-landing using liquid-propellant vernier rockets.

The Bellcomm team outlined six basic Lunar Viking missions; some included several variants. For example, the first Lunar Viking mission, the Orbital Survey Mission, would have three variants. None would include a Lander and all would use only instruments planned for the Mars Viking Orbiter. All three would complete their main objectives a month after capture into lunar orbit.

The Orbital Survey Mission variant #1 would see a Viking Lunar Orbiter map the entire Moon in visual wavelengths at eight-meter resolution from 460-kilometer-high lunar polar orbit. Variant #2 would map the entire lunar surface in stereo at 12-meter resolution. For variant #3, a Lunar Viking Orbiter would operate in 100-kilometer orbit. This, the Bellcomm planners explained, would enable it to image potential Lunar Viking Lander and Space Tug landing sites at two-meter resolution.

The Mars Viking Orbiter was meant to transmit data at a rate of just 1000 bits per second over a distance ranging from tens of millions to hundreds of millions of kilometers (that is, from Mars to Earth). The Lunar Viking Orbiter, on the other hand, would transmit from only about 380,000 kilometers (that is, from the Moon), so in theory could transmit about 75,000 bits per second. The Viking Orbiter data recorder could, Bellcomm estimated, store up to 100 images. The Lunar Viking Orbiter would use these capabilities to image the Moon while it was out of radio contact over the farside hemisphere and transmit the farside images to Earth while it passed over the Nearside hemisphere.

A Titan III-C rocket would be sufficient to place the Lunar Viking Orbiter into a 100-kilometer circular lunar polar orbit with plenty of propellant remaining on board for additional maneuvers. An Atlas-Centaur SLV-3C rocket would suffice if after lunar-orbit capture no other maneuvers were planned.

The second type of Orbiter-only Lunar Viking mission would use a Titan III-C-launched Orbiter outfitted with a scientific instrument suite tailored specifically for lunar investigations. The Bellcomm team modeled their specialized Lunar Viking Orbiter science payload on instruments expected to be mounted in the Service Module of the advanced Apollo 16, Apollo 17, Apollo 18, and Apollo 19 CSMs.

The Bellcomm team's third Lunar Viking mission would establish twin Farside Geophysical Observatories. A Titan III-D/Centaur rocket - the rocket intended in 1970 to launch the 1975 Mars Vikings - could, they calculated, place a stripped-down Lunar Viking Orbiter with two Lunar Viking Landers attached into a 600-kilometer circular equatorial orbit. The twin Landers would then detach and land at two different Farside sites, out of direct radio contact with Earth. The Orbiter would serve as a communications satellite for retransmitting radio signals from the twin Landers. Landing site selection would be based on Orbital Survey Mission images.

The Farside Geophysical Observatory payload on the twin Landers would comprise instruments similar to those in the Apollo Lunar Scientific Experiment Package (ALSEP) the Apollo astronauts first deployed during Apollo 12. This would extend the exclusively Nearside Apollo seismic monitoring network to the farside hemisphere.

Unfortunately, a Lunar Viking Orbiter in 600-kilometer equatorial orbit could receive signals from each Lunar Viking Lander only about 10% of the time. The Bellcomm planners noted that an Orbiter in a 5000-kilometer circular equatorial orbit could communicate with a Lander at Tsiolkovskii crater (23° south latitude) 26% of the time. Launching on the Titan III-D/Centaur would, they explained, enable the stripped-down Lunar Viking Orbiter to carry enough propellants to capture into 600-kilometer orbit and, after it released the Landers, maneuver to a 5000-kilometer communications orbit for the remainder of the mission.

Bellcomm's fourth Lunar Viking mission, the Farside Geochemical Mission, would see a Lunar Viking Orbiter/augmented Lunar Viking Lander combination leave Earth atop a Titan III-D/Centaur and capture into a 2000-kilometer circular equatorial orbit. The augmented Lunar Viking Lander would detach and ignite its chemical-propellant motors to place itself into a 2000-kilometer-by-100-kilometer elliptical orbit, then would ignite them again to reach a 100-kilometer circular equatorial orbit.

Finally, it would use its solid-propellant motor to deorbit and chemical-propellant verniers to soft-land at a geologically interesting Farside site. The Bellcomm team proposed that it transport to the surface a rover weighing up to 2000 pounds. Neither the augmented Lunar Viking Lander nor the rover was described. The Orbiter, again stripped down to serve mainly as a communications satellite, would remain in its initial 2000-kilometer orbit throughout the mission.

The Polar Mission, fifth on Bellcomm's list, would see the Lunar Viking Orbiter and Lander perform science together much as the Mars Viking Orbiter and Lander were meant to do. The Orbiter would again serve as a relay, but would also carry a suite of scientific instruments. The Lunar Viking Orbiter would capture into a 100-kilometer lunar polar orbit. As it passed over the Moon's poles, it would search permanently shadowed polar craters for ice deposits.

If ice were found, the Orbiter would release the Lander and maneuver to a higher orbit to improve communications. The Lander, meanwhile, would touch down in cold darkness and use an arm-mounted scoop or perhaps a drill to collect surface material for analysis in an on-board automated lab.

The sixth and most complex Lunar Viking mission, the Transient Event Mission, would aim to find and study Transient Lunar Phenomena (TLP). The Bellcomm team, which devoted an entire appendix of their report to TLP studies, noted that TLP had been recorded for decades at many sites on the Moon by telescopic observers. Appearing as bright spots, color changes, and hazes, TLP were generally interpreted as volcanic gas releases tied, perhaps, to the tides Earth raises in the solid crust of the Moon.

According to the Bellcomm planners, about half of all TLP recorded by 1970 had occurred in and around 40-kilometer-wide Aristarchus crater, located just west of Mare Imbrium in one of the most geologically diverse areas of the Moon. The Lunar Viking Orbiter would thus spend as much time as possible within sight of Aristarchus. This requirement would, along with the need for good image resolution, dictate Lunar Viking Orbiter altitude and maneuvers.

Aristarchus is the largest and brightest crater in this Apollo 15 image. Image credit: NASA.

In June 1970, the Mars Viking Orbiter was expected to operate during a six-month Earth-Mars cruise and then for at least three months in Mars orbit. This meant that — in theory — the Lunar Viking Orbiter could be expected to seek TLP for nine months in lunar orbit. In practice, the spacecraft would pass in and out of night several times each day as it orbited the Moon from very near the beginning of its mission, placing added stress on its solar arrays, batteries, and temperature-sensitive systems.

The Bellcomm team expected that the Lunar Viking Orbiter might not last for nine months, but that it would last long enough to detect a pattern in the occurrence of TLP events. Based on this pattern, the Lunar Viking Lander would be directed to a site where it would be likely to witness a TLP event up close.

If the Lunar Viking Orbiter could not spot enough TLP events to enable scientists to detect a pattern, the Lander would be dispatched to Aristarchus. There it would seek evidence of past TLP and stand by in the hope that it might witness a TLP event.

The Bellcomm planners lamented an expected six-year gap in U.S. lunar landings. One wonders how they would have greeted the news that NASA would soft-land no spacecraft on the Moon after Apollo 17 in December 1972 - that after almost 50 years, Apollo 17 remains the last U.S. lunar soft-lander. Three automated soft-landers followed Apollo 17: the Soviet Union's Luna 21, which delivered the eight-wheeled Lunokhod 2 rover (1973); Luna 24, which collected and launched to Earth a small sample of lunar surface material (1976); and China's Chang'e 3 lander (2015), which delivered the small Yutu rover.

20 August 1975: Viking 1 launch atop a Titan III-E/Centaur rocket. Image credit: NASA.

The Viking 1 and Viking 2 spacecraft exceeded all expectations. Viking 1 reached Mars orbit on 19 June 1976. The Viking 1 Lander separated from its Orbiter and soft-landed on 20 July 1976. Viking 2 reached Mars on 7 August 1976, and its Lander touched down on 3 September 1976. The Viking Landers performed multiple life-detection experiments (with equivocal results). Together, the four spacecraft of Viking 1 and Viking 2 transmitted to Earth more than 100,000 images.

The Viking 2 Orbiter suffered a propulsion system leak and was turned off on 25 July 1978; the Viking 2 Lander suffered battery failure and was switched off on 11 April 1980. The Viking 1 Orbiter depleted its attitude-control gas supply and was turned off on 17 August 1980. Though designed to operate on Mars for 90 martian days (Sols), the Viking 1 Lander transmitted from Mars until 13 November 1982 — a total of 2245 Sols. It might have lasted longer, but a faulty command caused it to break contact with Earth.

NASA and its contractors proposed many Viking-derived missions for the late 1970s and early 1980s. These included rover and dual-rover missions, sample-returners, and landers and rovers for the martian moons Phobos and Deimos. Their planning efforts in some ways resembled those of Apollo planners in AAP and its successor/remnant, the Skylab Program. The Earth-orbiting Skylab Orbital Workshop was staffed three times in 1973-1974. There was, however, no Viking Applications Program; despite Viking's success, its spacecraft designs saw no further application.

Sources

The Post-Apollo Space Program: Directions for the Future, Space Task Group Report to the President, September 1969.

America's Next Decades in Space: A Report for the Space Task Group, NASA, September 1969.

Internal Note: Integrated Space Program - 1970-1990, IN-PD-SA-69-4, T. Sharpe & G. von Tiesenhausen, Advanced Systems Analysis Office, Program Development, NASA Marshall Space Flight Center, 10 December 1969

"U. S. Space Pace Slowed Severely," W. Normyle, Aviation Week & Space Technology, 19 January 1970, p. 16.

"Presentation Outline [Space Tug]," NASA Manned Spacecraft Center, 20 January 1970.

"NASA Budget Hits 7-Year Low," W. Normyle, Aviation Week & Space Technology, 2 February 1970, pp. 16-18.

"Viking Spacecraft for Lunar Exploration - Case 340," R. Kostoff, M. Liwshitz, S. Shapiro, W. Sill, and A. Sinclair, Bellcomm, Inc., 30 June 1970.

On Mars: Exploration of the Red Planet, 1958-1978, NASA SP-4212, E. Ezell and L. Ezell, NASA, 1984, pp. 128-153, pp. 185-201, pp. 245-284.

More Information

"Assuming That Everything Goes Perfectly Well In The Apollo Program. . ." (1967)

The Russians are Roving! The Russians are Roving! A 1970 JPL Plan for a 1979 Mars Rover

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

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