19 August 2017

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

The Viking 2 rover on the frosty plain at Utopia. Image credit: Pat Rawlings/NASA
Even before Viking 1 landed on Mars (20 July 1976), NASA and its contractors studied post-Viking robotic Mars missions. Prominent among them was Mars Sample Return (MSR), considered by many to be the most scientifically significant robotic Mars mission.

The Viking missions reinforced this view of MSR, and also revealed the perils of making too many assumptions when planning costly and complex Mars exploration missions. The centerpiece of the $1-billion Viking mission, a briefcase-sized package of three biology experiments, yielded more questions than answers. Most scientists interpreted their data as evidence of previously unsuspected reactive soil chemistry, not biology. The truth, however, was that no one could be certain what the Viking biology experiment results meant.

With that unsatisfying experience in mind, A. G. W. Cameron, chair of the National Academy of Sciences Space Science Board, wrote in a 23 November 1976 letter to NASA Administrator James Fletcher that
[to] better define the nature and state of Martian materials for intelligent selection for sample return, it is essential that precursor investigations explore the diversity of Martian terrains that are apparent on both global and local scales. To this end, measurements at single points. . .should be carried out as well as intensive local investigations of areas 10-100 [kilometers] in extent.
Soon after Cameron wrote his letter, NASA Headquarters asked the Jet Propulsion Laboratory (JPL) to study a 1984 MSR precursor mission. The JPL study, results of which were due by July 1977, was meant to prepare NASA to request "new start" funds for the 1984 mission in Fiscal Year 1979. NASA also created the Mars Science Working Group (MSWG) to advise JPL on the mission's science requirements. The MSWG, chaired by Brown University's Thomas Mutch, included planetary scientists from several NASA centers, the U.S. Geological Survey (USGS) Astrogeology Branch, and Viking contractor TRW.

The MSWG's July 1977 report called the Mars 1984 mission the "next logical step" in "a continuing saga" of Mars exploration and a "required precursor" for an MSR mission, which it targeted for 1990. Mars 1984 would, it explained, provide new insights into the planet's internal structure and magnetic field, surface and sub-surface chemistry and mineralogy ("especially as related to the reactive surface chemistry observed by Viking"), atmosphere dynamics, water distribution and state, and geology of major landforms.

The Mars 1984 mission would also seek answers to "The Biology Question." The MSWG declared that
on-going exploration of Mars must address the issue of biology. Although there does not appear to be active biology at the two Viking landing sites, there may be other localities with special environments conducive to life. Life-supportive aspects of the Martian environment must be defined in greater detail. The characterization of former environments [and] a search for fossil life. . .should be conducted.
Mars 1984 would begin in December 1983-January 1984 with two Space Shuttle launches no less than seven days apart. The piloted, reusable Space Shuttle Orbiters would each place into low-Earth orbit a Mars 1984 spacecraft comprising one 3683-kilogram orbiter based on the Viking Orbiter design, three penetrators with a combined mass of 214 kilograms, and one 1210-kilogram lander/rover combination housed in an extended Viking bioshield/aeroshell. Together with an adapter linking it to a two-stage Intermediate Upper Stage (IUS), each Mars 1984 spacecraft would weigh a total of 5195 kilograms.

A Viking orbiter releases an aeroshell containing a Viking Mars lander. The Mars 1984 orbiter would have a similar design; the aeroshell, however, would stand taller to provide sufficient room for the lander/rover combination within it.

Viking aeroshell (left) and Mars 1984 aeroshell. Image credit: Martin Marietta
The Shuttle Orbiters would each deploy a spacecraft/IUS combination from its payload bay, then would maneuver away before IUS first-stage ignition. The MSWG calculated that the IUS would be capable of placing 5385 kilograms on course for Mars on 2 January 1984, near the middle of a launch opportunity spanning 28 days.

The twin Mars 1984 spacecraft would reach Mars from 14 to 26 days apart between 25 September and 18 October 1984, after voyages lasting a little more than nine months. Each would perform a final course-correction rocket burn using attitude control thrusters a few days before planned Mars Orbit Insertion (MOI). Their penetrators would separate two days before MOI and fire small solid-propellant rocket motors to steer toward their target impact sites on Mars. The motors would then separate from the penetrators.

During MOI, each spacecraft would fire a solid-propellant braking rocket motor, then the orbiter's liquid-propellant maneuvering engine would ignite to place it into a 500-by-112,000-kilometer "holding" orbit with a five-day period. Spacecraft #1's orbit would be near-polar, while spacecraft #2 would enter an orbit tilted from 30° to 50° relative to the martian equator. MOI completed, flight controllers would turn the orbiter's cameras toward Mars to assess weather conditions ahead of lander separation.

The Bendix Mars penetrator was designed to enter Mars's atmosphere directly from an interplanetary trajectory and embed itself in solid rock. A = radio antenna; B = meteorology package and magnetometer; C = isotope heater; D = aft body electronics; E = Aft body/fore body separation plane; F = cable linking aft body and fore body; G = accelerometer and neutron detector; H = fore body electronics; I = drill assembly; J = sampling drill bit; K = geochemical analysis package; L = seismometer; M = batteries; N = radioisotope thermal generator. Image credit: Bendix Corporation 
At about the time the twin spacecraft entered their respective holding orbits, the six penetrators would impact at widely scattered points. Each would split at impact into two parts linked by a cable. The aft body, which would include a weather station and an antenna for transmitting data to the orbiters, would protrude from the martian surface after impact. The fore body would include a drill for sampling beneath Mars's surface and a seismometer. According to the MSWG, penetrators were "the only economic means" of establishing a Mars-wide sensor network. Establishing a network of widely scattered seismometers was considered vital for charting the planet's interior structure.

After several months in holding orbit, spacecraft #2 would move to a 300-by-33,700-kilometer "magneto orbit," where it would explore Mars's magnetospheric bow wave and tail. It would then maneuver to a 500-by-33,500-kilometer "landing orbit" with a period of one martian day (24.6 hours). During a one-month landing site certification period, scientists and engineers would closely inspect orbiter images of the candidate landing site. Spacecraft #1, meanwhile, would proceed directly from holding orbit to landing orbit.
The Mars 1984 landing system for delivering the Mars 1984 rover to the surface would include five main parts. 1= top bioshield for protecting the sterilized lander and rover from contamination; 2 =  top aeroshell for protecting the lander from reentry heating; 3 = folded lander (rover not displayed); 4 = bottom aeroshell with attitude control/deorbit thrusters and propellant tanks; 5 = bottom bioshield/heat shield. Landing would occur as follows: the top bioshield would be left behind on the Mars 1984 orbiter as the rest of the lander moved away; motors on the bottom aeroshell would ignite to deorbit the lander; following reentry, the top aeroshell would deploy a single large parachute; the bottom aeroshell/heat shield would fall away; and the lander would fall free of the top aeroshell and ignite its landing motors for final descent. Image credit: Martin Marietta
The Mars 1984 landers would have one purpose: to deliver the Mars 1984 rovers to Mars's surface. Lander #2 would set down first at about 6° south latitude and lander #1 would land at about 44° north latitude at least 30 days later. JPL estimated that imaging data from the Viking orbiters would enable each Mars 1984 lander to set down safely within a "error ellipse" 40 kilometers wide by 65 kilometers long (for comparison, Viking's landing ellipse measured 100 kilometers wide by 300 kilometers long).

The Mars 1984 landers, based on a Martin Marietta design, would each include a "terminal site selection system." This would steer them away from boulders and other hazards as they descended the final kilometer to the martian surface. In other respects, their deorbit and landing systems would closely resemble those of the Vikings.

After lander separation, orbiter #1 would maneuver to a 500-kilometer near-polar circular orbit and orbiter #2 would move to a 1000-kilometer near-equatorial circular orbit. Orbiter #1's low near-polar orbit would permit global mapping at 10-meter resolution, while orbiter #2's more lofty near-equatorial orbit would enable it to map the equatorial region at 70-meter resolution. Low-flying Orbiter #1 would serve as the radio relay for the six penetrators, which would transmit relatively weak signals, while orbiter #2 would relay signals to and from the twin rovers.

The MSWG expected that most orbiter science operations would require minimal planning, since they would "be highly repetitive with most instruments acquiring data continuously and sending it to Earth in real time without tape recording." The exception would be imaging operations, since imaging data would be "acquired at a rate many times too great for real-time transmission." The MSWG suggested that the orbiters relay to Earth about 80 images of Mars per day.

Mars 1984 rover. A = antenna for signal relay through orbiter #2; B = antenna for direct transmission to and from Deep Space Network antennas on Earth; C = optics port cluster and strobe light (1 of 2); D = imaging/laser rangefinder mast (1 of 2); E = selenide radioisotope thermal generator (cover removed to display cooling vanes); F = rover chassis; G = manipulator arm with sampling drill (folded in travel position); H = sample-analysis inlet port; I = hazard detectors; J = loopwheel mobility system (1 of 4).

Mars 1984 rover and lander folded within their aeroshell and bioshield. A = folded landing leg (1 of 3); B = Viking-type landing footpad (1 of 3); C = lander body; D = Viking-type terminal descent engine (1 of 3); E = Viking-type parachute canister with deployment mortar; F = terminal site selection system sensors; G = folded rover ramp (1 of 2); H = folded loopwheel mobility system (2 of 4); I = stowed imaging/laser rangefinder mast (1 of 2); J = folded antenna for direct communication with Earth; K = rover chassis; L = radioisotope thermal generator; M = outer surface of aeroshell (tanks and thrusters not shown); N = outer surface of bioshield (heat shield not shown); O = attachment point linking bioshield to Mars 1984 orbiter. Image credit: Martin Marietta
The MSWG envisioned that the Mars 1984 rovers would be "substantial vehicles" capable of traveling up to 150 kilometers in two years at a rate of 300 meters per day. They based their rover concept on a Jet Propulsion Laboratory (JPL) design. Each would include four "loop-wheel" treads on articulated legs, a radioisotope thermal generator providing heat and electricity, laser range-finders for hazard avoidance, an "improved Viking-type manipulator" arm, twin cameras for stereo imaging, a microscope, a percussion drill for sampling rocks to a depth of 25 centimeters, and a sample processor for distributing martian materials to an on-board automated laboratory for analysis.

The MSWG acknowledged that a costly automated lab on an MSR precursor mission might be hard to justify, given that the MSR mission meant to follow it was intended to return samples to well-equipped labs on Earth for detailed analysis. The group argued, however, that clues to the nature of the reactive soil chemistry found by the Vikings might "reside in loosely bound complexes or interstitial gases" that "would be extraordinarily difficult to preserve in a returned sample." The scientists might also have worried that the planned MSR mission would be postponed or cancelled, leading them to attempt to exploit every opportunity to acquire new data.

The rovers would store particularly interesting samples for collection during the MSR mission and test the effects of Mars's reactive soil chemistry on MSR sample container materials. They would also each drop off three seismometer/weather stations as they moved over the surface to create a pair of 20-kilometer-wide regional sensor networks.

The rovers would employ three Mars surface operation modes. The first, Site Investigation Mode, would enable "intensive investigation of a scientifically interesting site." The rover would be fully controlled from Earth.

In Survey Traverse Mode, the second mode, the rover would operate nearly autonomously in a "halt-sense-think-travel-halt" cycle. Each survey/traverse cycle would last about 50 minutes and move the rover forward from 30 to 40 meters. Science operations would occur during the "halt" portion and while the rover was parked at night. Flight controllers would update rover commands once per day. The rover would cease autonomous operations and alert Earth when it encountered a hazard or a feature of scientific interest.

The third mode, Reconnaissance Traverse Mode, would occur when the terrain was sufficiently smooth (and scientifically dull) to allow the rover to move at its top speed of 93 meters per hour. The rover would make few science stops and would travel both by day and by night.

Valles Marineris with Mars 1984 landing ellipses marked in red and labeled. Image credit: NASA
To conclude its report, the MSWG drew on USGS studies based on Mariner 9 and Viking orbiter data to offer two candidate near-equatorial landing sites for lander #2. Capri Chasma, at the eastern end of Valles Marineris, included heavily cratered (thus ancient) highlands terrain, lava flows of different ages, lava channels, and possible water-related channels and deposits. Candor Chasma, a north-central branch of Valles Marineris, included at least two rock types in its four-kilometer-high canyon walls. The group expected that a Mars 1984 rover might find ancient crystalline rocks on the canyon floor.

New Mars missions stood little chance of acceptance in the late 1970s, when NASA's limited resources were largely devoted to Space Shuttle development and public enthusiasm for the Red Planet was (thanks the equivocal Viking biology results) at a nadir. Though MSR remained a high scientific priority (as it does today), the planetary science community opted to seek support for missions to other destinations: for example, the Jupiter Orbiter and Probe mission, later renamed Galileo, got its start in NASA's Fiscal Year 1978 budget.

NASA's next Mars spacecraft, the Mars Observer orbiter, was approved in 1985 for a 1990 launch; launch was subsequently postponed until September 1992, then the spacecraft failed during Mars orbit insertion in August 1993. NASA would return successfully to Mars for the first time since Viking in July 1997, when the 264-kilogram Mars Pathfinder spacecraft landed in Ares Valles bearing the 10.6-kilogram rover Sojourner.


Post-Viking Biological Investigations of Mars, Committee on Planetary Biology and Chemical Evolution, Space Science Board, National Academy of Sciences, 1977

Mars '84 Landing System Definition: Final Report, "Technical Report," Martin Marietta, April 1977

A Mars 1984 Mission, NASA TM-78419, "Report of the Mars Science Working Group," July 1977

More Information

Robot Rendezvous at Hadley Rille (1968)

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

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

12 August 2017

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

Cutaway of a reentering Apollo Command Module showing the position of its crew. Image credit: NASA
On 14 April 1964, a NASA Atlas-D rocket lifted off from Cape Kennedy, Florida, bearing the first Flight Investigation Reentry Environment (FIRE) payload. Project FIRE aimed to gather data on atmosphere reentry at lunar-return speed - about 36,000 feet per second (fps) - to enable Apollo engineers to develop the heat shield for the conical Apollo Command Module (CM).

Initiated in 1961 and managed by NASA's Langley Research Center (LaRC) under direction of the NASA Headquarters Office of Advanced Research and Technology, FIRE focused mainly on testing instrumented sub-scale model CM capsules in wind tunnels and thermal chambers at LaRC. Engineers realized, however, that there could be no substitute for data gathered in the actual spaceflight environment.

NASA rolls back the gantry structure surrounding the Atlas-D rocket bearing the first Project FIRE spacecraft, April 1964. Image credit: NASA
The Atlas-D rocket lobbed the Project FIRE payload, the 14-foot-long, 4150-pound Velocity Package (VP), onto an arcing course toward remote Ascension Island in the South Atlantic Ocean, a British possession that since 1957 had been home to U.S. missile tracking facilities. The VP cast off its two-part aerodynamic shroud and separated from the spent Atlas-D a little more than five minutes after liftoff. Attitude control motors mounted in its roughly cylindrical support shell then ignited to adjust its pitch so that it pointed its nose at Earth at a shallow angle.

About 21 minutes after separation from the Atlas-D and 800 kilometers above Earth, three rockets on the support shell ignited to spin the VP, giving it gyroscopic stability. Three seconds later, the VP cast off the support shell, revealing the engine bell of its solid-propellant Antares II-A5 rocket motor. Three seconds after support shell separation, the 24,000-pound-thrust motor ignited, driving the VP toward Earth's atmosphere.

The Antares motor burned out 33 seconds later, with the VP moving at nearly 37,000 fps. About 26 seconds later, the Apollo CM-shaped Reentry Rackage (RP) separated. Seven seconds after that, the 200-pound capsule fell past 400,000 feet, where the aerodynamic effects of reentry began to become obvious.

Image credit: NASA
Project FIRE Reentry Package. Image credit: NASA
The FIRE RP's heat shield heated rapidly as the falling capsule compressed and heated the atmosphere in its path. More than 300 sensors gathered data on the high-speed reentry environment. As the RP achieved a maximum speed of about 38,000 fps, the shockwave in front of the heat shield reached about 20,000° Fahrenheit (that is, about twice as hot as the Sun's surface).

Reentry heating formed a sheath of ionized gas around the FIRE RP, blocking radio signals. During the "blackout" period, which lasted for about 40 seconds, the RP stored data on magnetic tape. It transmitted the data after blackout ended.

Observers on Ascension Island - where the Sun had set - were able to track the FIRE RP visually as it automatically threw off two layers of heat shield material. They also observed the destructive reentry of the spent Antares II-A5 motor.

Thirty-two minutes after launch, the RP splashed into the Atlantic southeast of Ascension, about 5200 miles from Cape Kennedy. It was not designed for recovery.

NASA carried out the Project FIRE II test 13 months later, on 22 May 1965. The FIRE II RP was nicknamed the "flying thermometer" because it transmitted more than 100,000 temperature readings before ocean impact 5130 miles from Cape Kennedy. After FIRE II, engineers felt confident that they understood the atmosphere reentry effects the Apollo CM would experience as it returned from the moon.

The unmanned Apollo 4 (November 1967) and Apollo 6 (April 1968) Saturn V test missions carried out full-scale Apollo CM reentry tests. Astronauts first put the CM heat shield to the test at lunar-return speed during the Apollo 8 mission, which saw the second manned Apollo Command and Service Module (CSM) spacecraft orbit the moon 10 times on Christmas Eve 1968. Frank Borman, James Lovell, and William Anders reentered Earth's atmosphere in the Apollo 8 CM at nearly 36,000 fps on 27 December and splashed down safely in the Pacific southwest of Hawaii.

The FIRE flight tests were fresh in the minds of D. Cassidy, H. London, and R. Sehgal, engineers with Bellcomm, when they wrote a 14 April 1966 memorandum that proposed heat shield tests ahead of piloted Mars and Venus missions. Bellcomm was formed in 1962 to serve as the NASA Headquarters Apollo planning contractor, but almost immediately had extended its bailiwick to include planning beyond Apollo.

A piloted flyby spacecraft of the 1970s dispenses automated probes near Mars while a radar dish and a telescopic camera scrutinize the planet. Image credit: NASA
The three engineers wrote that Mars has a noticeably elliptical orbit around the Sun. Because of this, a piloted Mars flyby mission with a duration of 1.5 years would return to Earth at speeds ranging between 45,000 and 60,000 fps depending on where Mars was in its orbit when the flyby took place. A two-year Mars flyby mission would reenter Earth's atmosphere at between 45,000 and 52,000 fps. An opposition-class (short-stay) Mars "stopover" (orbiter or landing) mission would reenter at between 50,000 and 70,000 fps.

Venus, by contrast, has a nearly circular orbit around the Sun, so all flyby missions would return to Earth moving at about 45,000 fps. All Venus stopovers would reach Earth moving at between 45,000 and 50,000 fps. An opposition-class Mars stopover mission that flew past Venus before reaching Mars to speed up so that it could use a slow Earth-return path or flew past Venus during return from Mars to slow its approach to Earth would also reenter at between 45,000 and 50,000 fps.

Cassidy, London, and Sehgal wrote that, at speeds beyond 50,000 fps, reentry data gathered through testing for Apollo lunar missions no longer applied. Reentry heating would occur through different mechanisms and encompass a broader swath of the electromagnetic spectrum. This would increase turbulence and decrease the effectiveness of Apollo-type ablative heat shields (that is, heat shields designed to char and erode to dissipate reentry heat). In fact, at speeds beyond 50,000 fps, shield fragments detached by ablation could contribute to turbulence and heating.

The Bellcomm engineers acknowledged that braking propulsion might be used to slow a crew capsule to a better-understood Earth-atmosphere reentry velocity. They calculated, however, that slowing a piloted crew capsule derived from the Apollo CM from 70,000 fps to 50,000 fps would double the Earth-departure mass of the entire Mars stopover spacecraft. This would occur because extra propellants would be needed to launch the Earth-reentry braking propellants from Earth orbit to Mars and back again. Doubling the mass of the Mars spacecraft would in turn double the number of expensive heavy-lift rockets required to launch its components and propellants from Earth's surface to assembly orbit about the Earth.

They acknowledged that ground tests had provided some data on the interplanetary reentry velocity regime, but warned that the problem of aerodynamic surface heating involved "a complex interaction of vehicle size, shape[,] and heat protection characteristics." There would be, they added, “no substitute for testing specific configurations and materials in the actual environment of interest."

Cassidy, London, and Sehgal proposed that up to eight reentry capsules with attached solid-propellant motors be added to an Apollo Applications Program (AAP) Saturn V flight. AAP was NASA's planned post-Apollo program of Earth-orbital and lunar missions. The program aimed to use Apollo lunar mission vehicles in new ways. In addition to keeping the Apollo industrial team intact, AAP would see astronauts perform pioneering space biomedical and technology testing in Earth and lunar orbit, paving the way for piloted interplanetary voyages in the mid-to-late 1970s and the 1980s.

Image credit: NASA
Saturn V S-IVB third stage with cutaway and section showing spin tables and reentry capsules within the aft adapter that would link the stage to the Saturn V S-II second stage. Also shown is an Apollo Lunar Excursion Module (LEM)-derived lunar laboratory within the forward adapter that would link the top of the S-IVB to the bottom of the Apollo CSM. Image credit: Bellcomm/NASA
The Bellcomm trio proposed an interplanetary reentry test during a piloted lunar-orbital mission. The eight reentry capsules, each with a solid-propellant motor, might be housed in the adapter linking the bottom of the Saturn V S-IVB third stage with the top of the S-II second stage. Normally S-IVB separation would see the adapter left behind on the S-II, but for this mission it would remain attached to the S-IVB. Each reentry capsule-motor combination would be mounted on an individual spin table to spin it about its long axis for gyroscopic stability before release.

The AAP mission Cassidy, London, and Sehgal envisioned would include an Apollo CSM and a small lunar-orbital laboratory derived from the Apollo Lunar Module (LM) lander. The S-IVB's single J-2 engine would accelerate the S-IVB stage, the S-II/S-IVB adapter, the eight reentry capsules and their associated hardware, the LM Lab, and the CSM out of Earth parking orbit into a high elliptical Earth orbit.

After S-IVB shutdown, the crew in the CSM would detach their spacecraft from the stage, turn it end for end, and dock it with the LM Lab. They would extract the LM Lab from the front end of the S-IVB stage, then ignite the CSM's Service Propulsion System (SPS) main engine to place the CSM/LM Lab combination on course for the moon. A few days later they would fire the SPS again to enter orbit around the moon.

The S-IVB stage would retain about 30,000 pounds of liquid hydrogen/liquid oxygen propellants after the CSM and LM Lab went on their way. About 12 hours after departure from parking orbit, the S-IVB, with its cargo of reentry capsules and solid-propellant motors, would reach its maximum altitude above the Earth. The stage would aim at Earth, restart, and burn all of its remaining propellants, attaining a velocity of about 41,100 fps.

After J-2 engine shutdown, the spin tables would spin up the eight reentry capsules and their motors, then springs would push them out of the S-II/S-IVB adapter. Once clear of the S-IVB stage, the motors would ignite to further accelerate the reentry capsules.

Cassidy, London, and Sehgal calculated that Project FIRE's Antares II-A5 motor could increase a 10-pound reentry capsule's speed to 56,100 fps after release from the S-IVB stage. It could boost a 200-pound capsule to 48,500 fps. A TE-364 solid-propellant motor of the type used to brake unmanned Surveyor landers during descent to the lunar surface could accelerate a 10-pound capsule to nearly 60,000 fps. A 200-pound capsule with a TE-364 motor could attain 53,500 fps.


"NASA Schedules Project FIRE Launch," NASA News Release No. 64-69, April 14, 1964

Astronautics & Aeronautics, 1964: Chronology on Science, Technology, and Policy, NASA SP-4005, NASA Historical Staff, Office of Policy Planning, 1965, pp. 135, 350

"Reentry Heating Experiment on Saturn V AAP Flights or Unmanned Saturn IB Flights - Case 218," D. Cassidy, H. London, and R. Sehgal, Bellcomm, 14 April 1966

Astronautics & Aeronautics, 1965: Chronology on Science, Technology, and Policy, NASA SP-4006, NASA Historical Staff, Office of Policy Analysis, 1966, pp. 244

Project FIRE in Langley Researcher - https://crgis.ndc.nasa.gov/crgis/images/2/26/Project_Fire_Newsletters.pdf

More Information

Starfish and Apollo (1962) (Bellcomm)

After EMPIRE: Using Apollo Technology to Explore Mars and Venus (1965) (piloted flybys)

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

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

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

27 July 2017

Flyby's Last Gasp: North American Rockwell's S-IIB Interplanetary Booster (1968)

Stacking a Saturn V rocket: inside the Vertical Assembly Building at Kennedy Space Center, a giant crane gingerly lowers an S-II second stage onto an S-IC first stage. Image credit: NASA
NASA abandoned work toward piloted Mars and Venus flyby missions based on hardware developed for Apollo and its planned successor, the Apollo Applications Program, during the final months of the pivotal year 1967. Until August of that year, however, the concept was viewed by many as a plausible interim step between 1960s Apollo moon landings and 1980s piloted Mars landings.

Though NASA awarded no new piloted flyby study contracts, studies performed in 1965, 1966, and 1967 continued to report out at aerospace conferences and in NASA briefings during 1968 and 1969. In March 1968, for example, North American Rockwell (NAR) engineers W. Morita and J. Sandford summed up a study they completed in April 1967 for NASA's Marshall Space Flight Center (MSFC) in Huntsville, Alabama. Their study looked at how a modified NAR-built S-II rocket stage might be used to boost a piloted flyby spacecraft out of Earth orbit (that is, "inject" it onto an interplanetary trajectory). They presented results of their study at the Fifth Space Congress in Cocoa Beach, Florida.

Image credit: NASA
The 33-foot-diameter, 81.5-foot-long S-II, the second stage of the Apollo Saturn V rocket, weighed about 40 tons empty. A single propellant tank divided by a dome-shaped "common bulkhead" held a total of more than 400 tons of liquid oxygen (LOX) and liquid hydrogen (LH2) propellants. LH2 is of low density, so the LH2 section in the top/front part of the tank measured more than twice as long as the LOX section.

The propellants fed a cluster of five J-2 rocket engines, each producing 200,000 pounds of thrust. Together they consumed more than a ton of propellants per second during their 6.5 minutes (390 seconds) of operation, boosting the Saturn V's speed from 6000 miles per hour at separation from the Saturn V S-IC first stage to 17,400 miles per hour (just short of Earth-orbital velocity) at S-II shutdown.

NAR proposed to launch the S-II interplanetary boost stage, which it designated the S-IIB, into Earth orbit on a two-stage Saturn V. The S-IIB would include two or three improved J-2S engines in place of the S-II's five J-2s. After separation from the spent S-II, the J-2S engines would fire briefly to place the S-IIB into an elliptical Earth orbit. An auxiliary propulsion system made up of three solid-propellant motors would perform orbit circularization, and eight thruster modules based on the Apollo Command and Service Module (CSM) attitude control system would carry out orbit corrections and rendezvous and docking with the piloted flyby spacecraft.

Proposed North American Rockwell-built piloted flyby payloads are shown in red. Image credit: NAR/DSFPortree
The S-IIB would reach orbit with about 76 tons of LH2 fuel on board. NAR's analysis determined that, if only standard S-II thermal insulation were employed, boil-off caused by solar heating in orbit would reduce this to only 25 tons in less than five days. NAR proposed to reduce boil-off by installing a hydrogen gas-filled "vapor barrier" between the LH2 and LOX sections of the propellant tank and by applying "super-insulation" panels to the stage exterior. These modifications would reduce total LH2 boil-off over 10 days - the rated orbital lifetime of the S-IIB - to less than five tons.

The S-IIB would need to lift off with its LOX tank empty if the two-stage Saturn V was to place it in Earth orbit. Separately launched automated LOX tankers would then dock with it to fill the tank. The NAR engineers examined S-II-based tankers, tankers based on the Apollo Saturn S-IVB stage, and a wholly new tanker Lockheed Corporation designed in a separate study for MSFC.

LOX tankers considered in the North American Rockwell study. Green represents each design's LOX cargo volume. Image credit: NAR/DSFPortree
Morita and Sandford described two S-II-based tankers. The first, the S-IIB/TK, would measure about 25 feet shorter than the standard Saturn V S-II stage. It would separate from the S-II second stage of the two-stage Saturn V that launched it, fire its twin J-2S engines for 3.5 minutes to attain a 100-nautical-mile-by-263.5-nautical-mile orbit, then fire them again at apogee (the high point in its orbit about the Earth) to raise its perigee (the low point in its orbit about the Earth). The circularization burn would leave the S-IIB/TK into a 263.5-nautical-mile-high parking orbit.

The 92 tons of LOX remaining after the circularization burn would constitute the tanker's payload. Solar heating would cause the LOX to boil off over time, so after 163 days - the longest period the tanker would need to loiter in Earth orbit before transferring its payload to the S-IIB injection stage - 75 tons would remain.

NAR's second S-II tanker variant, the S-II/TK, would have a LOX tank four feet longer than that of the standard Saturn V S-II. It would serve double-duty as a Saturn V second stage and a tanker. After it separated from the S-IC first stage, its five J-2S engines would boost it into a 100-nautical-mile-by-263.5-nautical-mile orbit, Earth orbit, then two engines would fire a second time at apogee to circularize its orbit. The S-II/TK would retain about 105 tons of LOX after the circularization burn and about 82 tons after 163 days in orbit.

Sandford and Morita next examined tankers based on the Douglas Aircraft Company-built S-IVB stage. The 22-foot-diameter S-IVB served as the the second stage of the Saturn IB rocket and the third stage of the Saturn V moon rocket.

The first S-IVB tanker design would trim cost by retaining - but leaving empty - the S-IVB stage LH2 tank. The second would delete the LH2 tank, making for a tanker that was shorter and lighter, but more heavily modified and thus more costly. The first design would deliver 110.5 tons of LOX to 263.5-nautical-mile orbit, of which about 99 tons would remain after 163 days. The second S-IVB-based design would deliver 107.5 tons to a 263.5-nautical-mile circular parking orbit. Of this, 92.5 tons would remain after 163 days.

The third tanker Morita and Sandford investigated was Lockheed's Orbital Tanker. Because it would be purpose-built to serve as a tanker, it would be more efficient than the NAR S-II and Douglas S-IVB tankers, but also more costly. Efficiency in this case would be measured in terms of the expected amount of LOX boil-off.

After launch on a two-stage Saturn V, the Orbital Tanker would fire LH2/LOX or solid-propellant rocket motors to place itself into a 263.5-nautical-mile-high parking orbit. The Orbital Tanker would reach orbit bearing 114.9 tons of LOX in an insulated spherical tank. Of this, 110.9 tons would remain after 163 days.

Sandford and Morita looked at Mars and Venus flybys, but emphasized a Mars flyby that would leave Earth orbit in late September 1975. Their proposed Mars flyby launch schedule took into account the narrow range of Earth-orbit departure dates, the planned 10-day lifetime in Earth orbit of the S-IIB injection stage, and the existence of only two Launch Complex 39 Saturn V launch pads at NASA's Kennedy Space Center in Florida.

Assuming an Earth-orbit departure date of 20 September 1975, the piloted Mars flyby mission would begin with three LOX tanker launches in April-May 1975. They would lift off between 153 and 130 days before the scheduled launch to Earth orbit of the S-IIB injection stage. A Saturn V bearing a fourth, backup tanker would be held in reserve.

Following the launch of the third LOX tanker in May 1975, KSC ground teams would refurbish the twin Launch Complex 39 pads for launch of the backup tanker (if necessary), the piloted flyby spacecraft, and the S-IIB injection stage. NAR estimated that KSC workers would need no more than one eight-hour shift per day to ready the pads in time for the piloted flyby spacecraft and S-IIB stage launches in September 1975. More shifts would be added if the backup tanker became necessary; that is, if one of the first three tankers failed to reach orbit or malfunctioned in orbit while awaiting arrival of the spacecraft and S-IIB stage.

On 15 September 1975, the S-IIB injection stage would lift off, followed within 24 hours by the piloted flyby spacecraft. Spacecraft and stage would rendezvous and dock within 12 hours, then the combination would set out in pursuit of the waiting tankers.

The piloted flyby spacecraft/S-IIB combination would dock with the three LOX tankers about 12 hours apart. Each would in turn link up with the aft end of the S-IIB, transfer its LOX cargo, and detach.

The piloted flyby astronauts and mission controllers on Earth would then perform a detailed systems check of the piloted flyby spacecraft/S-IIB stage combination. If all checked out as normal, they would be certified ready to depart Earth orbit on 20 September, just as the launch window opened for a minimum-energy Earth-Mars free-return transfer.

The quantity of propellants required to depart Earth orbit on a Mars flyby trajectory would increase steadily from the moment the launch window opened. At the same time, boil-off would cause the quantity of propellants in the S-IIB stage to steadily decrease. Morita and Sandford calculated that the S-IIB stage would retain sufficient LH2 to boost the Mars flyby spacecraft out of Earth orbit toward Mars for five days after the launch window opened; that is, until 25 September 1975.


"The S-II Injection Stage for the Mars/Venus Flyby Mission," W. H. Morita and J. W. Sandford, Proceedings, Fifth Space Congress: The Challenge of the 1970s, pp. 10.1-1 – 10.1-22; paper presented in Cocoa Beach, Florida, 11-14 March 1968

More Information

After EMPIRE: Using Apollo Technology to Explore Mars and Venus (1965)

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

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

Two for the Price of One: 1980s Piloted Missions with Stopovers at Mars and Venus (1969)

23 July 2017

Plush Bug, Economy Bug, Shoestring Bug (1961)

President John F. Kennedy, flanked by Vice President Lyndon Baines Johnson (right) and NASA officials, addresses employees of the Manned Spacecraft Center (MSC) in Houston, Texas, on 12 September 1962. The Apollo lunar lander mockup in this image is a North American Aviation design. Image credit: John F. Kennedy Library/NASA
Following President John F. Kennedy's 25 May 1961 "moon speech" before a joint session of Congress, NASA re-directed Project Apollo. When first conceived, Apollo was to have been the next U.S. piloted spacecraft program after Mercury. In a plan drafted in 1959-1960, the Apollo spacecraft was described as an Earth-orbital vehicle designed for independent flight or ferry flights to an Earth-orbiting space station.

NASA planners hoped that eventually - by about 1970 - Apollo might lead to a circumlunar or lunar-orbital flight. Following JFK's speech, however, the program's goal became to land a man on the moon by 1970 and return him safely to the Earth. Almost immediately, many asked the obvious question: how would Apollo accomplish this epic feat?

As many as six lunar mission modes received consideration in 1961-1962, though two - Earth-Orbit Rendezvous (EOR) and Direct Ascent - emerged as early favorites. Both modes included several variants.

In the EOR mode, one or more propulsion stages and a piloted moonship (and sometimes tankers for filling the propulsion stages with propellants) were brought together in Earth orbit. The propulsion stage or stages were then fired to place the moonship and its crew on course for the moon.

In Direct Ascent, a single large rocket boosted the piloted moonship from Earth's surface directly to the moon. After the piloted spacecraft was placed on course for the moon, the EOR and Direct Ascent mission modes would be essentially identical.

The process by which the Apollo lunar mission mode decision was taken was complex and involved many players at NASA Headquarters and the NASA field centers, some of whom backed different modes at different times. Throughout the entire untidy 14-month-long process, however, engineer John Houbolt of the NASA Langley Research Center (LaRC) in Hampton, Virginia, staunchly advocated Lunar-Orbit Rendezvous (LOR). Houbolt did not originate the LOR mode: it dates back at least to 1948, when H. E. Ross described it in London before a meeting of the British Interplanetary Society.

John Houbolt at the chalkboard. He points toward an "LEV" (Lunar Excursion Vehicle); this would become the Apollo Lunar Module. Image credit: NASA
Houbolt gave briefings on LOR to several of the groups involved in the Apollo mode decision, including the Lundin Committee, the Ad Hoc Task Group for Study of Manned Lunar Landing by Rendezvous Techniques (the Heaton Committee), and the LaRC-based (but otherwise independent) Space Task Group. The Large Launch Vehicle Planning Group (the Golovin Committee) requested a detailed written report on LOR after Houbolt briefed it in August 1961.

Houbolt and his colleagues at LaRC explained in their 31 October 1961 report to the Golovin Committee that LOR would differ from EOR and Direct Ascent in the nature of its spacecraft. As noted above, in EOR and Direct Ascent a single Apollo spacecraft would accomplish all phases of the lunar landing mission. It would bear its occupants from the Earth to the moon, land them on the moon, and then transport them back to Earth. LOR, on the other hand, would see lunar mission phases divided between two distinct piloted vehicles. In the LOR scenario LaRC described, these spacecraft were the Apollo and the single-stage "Bug" lunar lander.

The three-man Apollo spacecraft, which in the LOR mode would come no nearer to the moon than lunar orbit, would include the mission's Earth-atmosphere reentry vehicle and a pair of propulsion modules for performing major maneuvers. The Bug would detach from the Apollo in lunar orbit, descend toward the moon's surface, land one or two astronauts gently on the moon, and then return to the Apollo mothership in lunar orbit. The crew would cast off the spent Bug, then the Apollo would return to Earth.

LaRC examined three Bug designs, which it dubbed "Shoestring," "Economy," and "Plush." The first, with a dry (no propellants or other expendables) mass of only 1270 pounds, would land one man on the moon for only a brief period and return no more than 50 pounds of lunar samples to the orbiting Apollo for transport to Earth. Of the three designs, the Shoestring Bug would come closest to fulfilling the strict letter of President Kennedy's mandate - that "a man" land on the moon.

The second design, the Economy Bug, would support two men on the moon for 24 hours. The lander's dry mass would total 2234 pounds; it would transport up to 100 pounds of rock samples from the moon's surface to the orbiting Apollo.

The Plush Bug, the scientists' favorite, would support two men on the moon for one week, providing them with adequate time to perform field geology at the lunar landing site. Plush Bug dry mass would total 3957 pounds; it could lift 150 pounds of samples to the orbiting Apollo.

According to Houbolt's team, an LOR landing would be safer than either a Direct Ascent or EOR landing because the Bug would be designed only for that function - in other modes the landing function would be compromised by the need to take into account other functions, such as Earth atmosphere reentry. Houbolt also proposed a spacecraft configuration that would further enhance astronaut safety: an Apollo with two Shoestring Bugs. If the first Shoestring Bug became trapped on the lunar surface, then the second would be used to mount a rescue.

John C. Houbolt's concept drawing of a Lunar Orbit Rendezvous spacecraft with twin Shoestring Bugs. The stack depicted would reach Earth parking orbit atop a Saturn C-3 rocket. Weights (in pounds) are for "wet" hardware (that is, loaded with propellants and expendables). Image credit: NASA 
As with the other proposed Apollo modes, many preparatory and precursor missions would lead up to LaRC's LOR moon landing. LaRC's proposed five-year "Master Flight Plan" would begin with 11 Ranger automated lunar rough-landing missions before October 1963. These would help engineers and scientists characterize the lunar surface so that they could design the Bug, lunar surface space suits, and surface exploration equipment.

Meanwhile, NASA would fly four Mercury manned Earth-orbital missions, each completing 18 Earth orbits. The missions, flown between February and August 1963, would enable doctors to gather basic data on human performance in space.

NASA would launch 15 Surveyor automated soft-landing missions to the moon between August 1963 and March 1966. Meanwhile, back on Earth, the space agency would drop Apollo reentry vehicles from aircraft 20 times between September 1963 and June 1964 to test glide characteristics, parachutes, and land landing systems (the LaRC engineers assumed a landing on U.S. soil).

Astronauts would practice rendezvous and docking using maneuverable Mercury Mark II spacecraft launched into Earth orbit atop modified Titan missiles. The two-seater Mercury Mark II, which was renamed Gemini in January 1962, would dock with separately launched unmanned Agena upper stage target vehicles during five missions spanning October 1963 to June 1964.

Six manned Mercury Mark II missions between August 1964 and June 1965 would see astronauts practice docking with Bug landers in Earth orbit. For each mission, the Bug and the Mercury Mark II would be launched together on a Saturn C-1 rocket. Saturn C-1 was envisioned as the first NASA launch vehicle designed specifically for piloted spaceflight.

The Mercury Mark II/Bug Saturn C-1 missions would overlap with eight Saturn C-1-launched Apollo suborbital and Earth-orbital flight tests spanning September 1964 through August 1965. A pair of Saturn C-1-launched manned Apollo/Bug test missions in Earth orbit would follow in September-October 1965, laying the groundwork for four piloted Apollo high-elliptical Earth-orbit and circumlunar/lunar-orbit missions between November 1965 and February 1966.

LaRC suggested that the high-elliptical and circumlunar/lunar-orbit missions, each of which would leave Earth on a Saturn C-3 or C-4 rocket, might be converted into manned lunar landing attempts if necessary: for example, if the Soviet Union were believed to be on the verge of launching its first piloted lunar landing attempt. Assuming, however, that they were not turned into landing missions, then the first four manned lunar landing attempts of the Apollo Program would occur between March and June 1966.

LaRC's LOR missions would begin with launch on a Saturn C-3 or C-4 from Cape Canaveral, Florida. The LaRC team noted that a direct flight to the moon from a fixed site on the Earth's surface could begin only during a short period each month if it sought to land at a specific lunar landing site. To circumvent this limitation, LaRC's Apollo mothership/Bug lander stack would enter low-Earth parking orbit before setting out for the moon. This would in effect give the mission a mobile launch site, providing planners with "complete freedom" in selecting lunar landing mission start time.

The first and largest of the two Apollo propulsion modules would burn all of its propellants to push the Apollo mothership with its small propulsion module and a Bug lander out of Earth orbit, then would separate. LaRC opted for an Earth-moon transfer lasting from 2.5 to three days.

LaRC noted that the mission could follow "a free return circumlunar trajectory" that would enable the Apollo/Bug stack to swing around the moon and return directly to Earth without additional propulsion. This would come in handy if the Apollo/Bug stack suffered a propulsion malfunction after Earth-orbit departure.

Assuming, however, that all occurred as planned, the astronauts in the Apollo/Bug stack would turn the small propulsion module toward its direction of motion as it looped behind the moon. Over the center of the lunar Farside hemisphere, they would ignite the module so that the moon's gravity could capture the Apollo/Bug stack into lunar orbit.

LaRC recommended a 50-mile-high circular orbit over the moon's equator, "especially if the exploration time on the lunar surface [was] to be of the order of a week." A spacecraft in such an orbit would pass over all points on the moon's equator every two hours. This meant that every two hours the Bug would have an opportunity to descend to a specific equatorial target landing site or to lift off from an equatorial site and perform a rendezvous with the orbiting Apollo.

If, on the other hand, the Apollo entered an orbit inclined relative to the lunar equator so that the Bug could descend to a non-equatorial landing site, the moon's slow rotation would gradually take the site out of the Apollo's orbital plane (that is, the mothership would no longer pass over the Bug landing site). The Bug or the Apollo (or both) would then need to expend propellants to perform a plane-change maneuver to match orbits before rendezvous and docking could take place. LaRC noted, however, that the plane change necessary after a one-day stay at a non-equatorial site would be "insignificant."

LaRC proposed that the moon landing occur close to local lunar midnight under a full Earth, "thus avoiding the bright glare and black shadows of the sunlit side." Prior to Bug separation, the crew would examine the moon from orbit to make their final landing site selection. One or two astronauts would then enter the Bug, undock, and fire its engines briefly to move away from the Apollo. This would prevent the Apollo from being enveloped in the Bug's engine plume when the more powerful "lunar letdown" maneuver began. The Bug's engines - LaRC recommended two for redundancy and improved maneuverability - would be capable of being throttled and gimbaled (that is, pivoted for steering).

Halfway around the moon from the selected landing site - that is, out of view of Earth, over the Farside - the Bug pilot would fire the engines to slow the lander by 60 feet per second. This would nudge its orbit so that it would intersect the surface at the landing site. The Bug would then coast for an hour, steadily losing altitude. Bug and Apollo would remain in visual and radio contact throughout the descent.

About 100 miles from the landing site, the Bug pilot would fire the twin engines to reduce speed, then would commence landing maneuvers. The Bug would gradually tip so that it would reach the landing site with its engines and landing leg footpads pointed down. The pilot would then have one minute of hover time to choose a safe spot for final letdown.

All of LaRC's Bug designs would employ a single pair of engines for descent and ascent. If landing proved impossible, the pilot could throttle up the engines and abort back to orbit. Assuming that all went as planned, the Bug would gently settle on the surface at the target landing site as its pilot throttled back to zero.

Following a period of surface activity, the astronaut or astronauts would prepare the Bug for liftoff. Just before liftoff, the orbiting Apollo mothership would climb into view above the Bug's horizon. The Bug pilot would spot it visually and with radar, then would ignite the Bug's engines. The lander would climb 10 miles high at 0.5 gravities of acceleration, then would coast along an arcing course for up to 33 minutes. Meanwhile, the Apollo spacecraft would orbit over the landing site and pull ahead of the Bug.

A gyroscope-equipped "inertial attitude reference" would provide guidance data to the Bug pilot; if electronic aids failed, however, he could complete rendezvous and docking using visual cues. The Bug pilot would start homing on the Apollo's blinking light beacon about 250 feet out. Docking would take place over the lunar night hemisphere to avoid sun glare and improve beacon visibility. After docking, the astronaut or astronauts would transfer to the Apollo and cast off the Bug.

The Apollo mothership's small propulsion module would ignite for a second time to push it out of lunar orbit, then would be cast off. LaRC reported that studies of optimum Earth-return trajectories for accomplishing land landings in the U.S. were in progress.

NASA formally adopted LOR in July 1962. The Apollo spacecraft became known as the Command and Service Module (CSM) and the Bug was designated the Lunar Excursion Module (LEM - later changed to Lunar Module, or LM). On 7 November 1962, NASA awarded the LEM contract to Grumman Aircraft Engineering Corporation in Bethpage, Long Island, New York. The Grumman design featured separate descent and ascent stages.

Grumman's winning November 1962 LEM concept featured five landing legs, two docking ports, and large curved glass windows. The design would evolve rapidly as the company and NASA confronted the challenges of landing a man on the moon. Image credit: NASA
Early LM designs took as their design inspiration small helicopters, so featured curved surfaces and large windows; as the LM design evolved, however, it became faceted and asymmetrical, with small triangular windows. These changes reduced the LM's mass; large, curved, multi-pane glass windows were, it was found, heavier square centimeter for square centimeter than the LM's metal skin.

No LM was as light as the heaviest LaRC Bug - the Apollo 11 LM had a dry mass of 9271 pounds, or about 2.5 times the dry mass of the Plush Bug. It is unlikely that Houbolt and his colleagues knowingly low-balled their mass estimates; rather, most Apollo systems ended up heavier than at first estimated because no one had built piloted lunar spacecraft before.

20 July 1969: The Apollo 11 LM Eagle pirouetted in lunar orbit so that Command Module Pilot Michael Collins, alone on board the Command Module Columbia, could inspect and photograph it. At Eagle's controls were Apollo 11 Commander Neil Armstrong and Lunar Module Pilot Edwin Aldrin. Less than two hours after Collins took this photograph, Eagle touched down on the moon's Sea of Tranquility, an ancient dusty plain pocked by impact craters. Image credit: NASA
In preparation for the first manned moon landing, the NASA carried out robotic Ranger, Surveyor, and Lunar Orbiter missions. They mainly focused on equatorial and near-equatorial areas of the lunar Nearside hemisphere. Three Rangers imaged small areas close-up as the fell toward destructive impact, five Surveyors soft-landed and imaged, analyzed, and dug into the lunar surface, and five Lunar Orbiters imaged large regions and small candidate LM landing sites. Meanwhile, 10 two-man Gemini missions in Earth orbit gave astronauts rendezvous and spacewalk practice.

The LM reached space for the first time atop a Saturn IB (an uprated Saturn C-1 variant) during the unmanned Apollo 5 mission in January 1968. Apollo 7, also launched on a Saturn IB, was a piloted CSM test in Earth orbit.

Apollo missions 8 through 17 each launched on a three-stage Saturn V rocket; originally designated Saturn C-5, the Saturn V was more powerful than the Saturn C-3 and Saturn C-4 rockets described in LaRC's report to the Golovin Committee, which were never built. Apollo 8 was a CSM-only piloted lunar orbital flight. Apollo 9 was a piloted test of the CSM and LM in Earth orbit. Apollo 10, which included a CSM and an LM, was a lunar-orbital dress rehearsal for the first lunar landing attempt scheduled for Apollo 11. Astronauts used the LOR mode to land successfully on the moon six times between July 1969 and December 1972 (Apollos 11, 12, 14, 15, 16, and 17).

In April 1970, the LM served as a lifeboat for Apollo 13 astronauts James Lovell, Fred Haise, and Jack Swigert after an oxygen tank explosion crippled the CSM Odyssey en route to the moon. Odyssey's propulsion, electricity generation, and life support systems were all compromised. Fortunately, they had an undamaged spacecraft at their disposal.

The Apollo 13 crew used LM Aquarius's single descent engine and navigation aids to change their course to a free-return path and speed their docked CSM and LM spacecraft back to Earth. They relied on the LM's life support system to provide oxygen and scrub exhaled carbon dioxide from their cabin air. Had an EOR or Direct-Ascent Apollo spacecraft suffered a similar mishap, its crew would almost certainly have perished through asphyxiation, collision with the moon (if moon-bound), or uncontrolled Earth-atmosphere reentry (if Earth-bound).


"Manned Lunar Landing Via Rendezvous," NASA Langley Research Center, presentation materials, 19 April 1961

Manned Lunar Landing Through Use of Lunar-Orbit Rendezvous, NASA Langley Research Center, 31 October 1961

The Apollo Spacecraft: A Chronology, NASA SP-4009, The NASA Historical Series, I. Ertel and M. Morse, Vol. I, pp. 81-202

Enchanted Rendezvous: John C. Houbolt and the Genesis of the Lunar-Orbit Rendezvous Concept, Monographs in Aerospace History #4, J. Hansen, NASA History Office, December 1995

More Information

Starfish and Apollo (1962)

The Spacewalks That Never Were: Gemini Extravehicular Planning Group (1965)

If an Apollo Lunar Module Crashed on the Moon, Could NASA Investigate the Cause? (1967)

"Still Under Active Consideration": Five Proposed Earth-Orbital Apollo Missions for the 1970s (1971)

09 July 2017

SEI Swan Song: International Lunar Resources Exploration Concept (1993)

In the top image, the Soviet Union's two-stage Energia heavy-lift rocket and Buran reusable shuttle orbiter ride a transporter to a launch pad at Baikonur Cosmodrome in Kazakhstan. A sturdy armature designed to hoist the combination upright on the pad obscures Energia's lower half. In the bottom image, 59-meter-tall Energia stands on a launch pad bearing Polyus, an experimental military payload developed in response to the U.S. Strategic Defense Initiative. After the Soviet Union crumbled and Russia became a potential international supplier of rockets and spacecraft, hopeful NASA advance planners tentatively tapped Energia to launch hardware for piloted moon and Mars missions. Image credit: NPO Energia
By the close of 1992, the handwriting had been on the wall for the Space Exploration Initiative (SEI) for more than two years. President George H. W. Bush had launched his moon and Mars exploration initiative on the 20th anniversary of the Apollo 11 lunar landing (20 July 1989), but it had almost immediately run headlong into a minefield of fiscal and political difficulties. The change of Presidential Administration in January 1993 was the final nail in SEI's coffin. Nevertheless, exploration planners across NASA continued to work toward SEI goals into early 1994.

In the same period, the Soviet Union was falling apart. Even as Bush called on NASA to return astronauts to the moon and launch them onward to Mars, Soviet domination in eastern Europe collapsed, then the Soviet Union itself began to disintegrate. A bungled coup d'etat in August 1991 undercut the authority of Soviet President Mikhail Gorbachev and led to the official demise of the Soviet Union on 26 December 1991. The largest state on Earth divided into more than a dozen countries, with the Russian Federation under President Boris Yeltsin emerging as the most significant.

The end of the U.S.-Soviet Cold War created dangers and opportunities. Some feared that, impelled by economic chaos in the former Soviet Union, scientists and engineers would sell their skills and knowledge abroad, leading to unprecedented global nuclear proliferation.

Others noted that high-level Soviet space officials had begun to peddle their space hardware at important aerospace meetings in the late 1980s. They saw an opportunity to, among other things, save the U.S./European/Japanese/Canadian Freedom Space Station from cancellation. Yeltsin and Bush agreed to wide-ranging space cooperation in June 1992, partly in the hope that NASA money might help to forestall an exodus of Russian aerospace talent.

In February 1993, Kent Joosten, an engineer in the Exploration Program Office (ExPO) at NASA's Johnson Space Center (JSC) in Houston, Texas, proposed a plan for lunar exploration which, he hoped, would take into account the emerging realities of post-Cold War space exploration. His International Lunar Resources Exploration Concept (ILREC) would, he wrote, reduce "development and recurring costs of human exploration beyond low-Earth orbit" and "enable lunar surface exploration capabilities significantly exceeding those of Apollo." It would do these things by exploiting the abundant oxygen in the lunar regolith (that is, surface material) as oxidizer for burning liquid hydrogen fuel brought from Earth, shipping most cargo to the moon separate from crews, employing Earth-based and moon-based teleoperations, and cooperating with the Russian Federation.

Joosten's concept was a variant of the Lunar Surface Rendezvous (LSR) mission mode. The Jet Propulsion Laboratory (JPL) in Pasadena, California, put forward LSR in 1961 as a candidate mode for achieving President John F. Kennedy's goal of a man on the moon by the end of the 1970s. In 1962, after NASA selected Lunar Orbit Rendezvous (LOR) as its Apollo lunar mission mode, the LSR scheme faded into obscurity. Joosten's concept was not inspired by the early 1960s scenario; instead, his work drew upon contemporary In-Situ Resource Utilization (ISRU) and Mars surface rendezvous techniques proposed for use in NASA's Mars Design Reference Mission 1.0 and Martin Marietta's Mars Direct scenario.

The Apollo LOR mode was designed to permit the U.S. to reach the moon quickly and relatively cheaply, not to support a sustained lunar presence. It split lunar mission functions between two piloted spacecraft, each of which comprised two modules. Modules were discarded after they fulfilled their functions.

Joosten's ILREC piloted moonship would be roughly intermediate in size between the Apollo Lunar Module (LM) (left) and the Apollo Command and Service Module (CSM) (right). This NASA artwork from 1966 is a partial cutaway showing two blue-clad astronauts moving from the CSM to the LM in preparation for undocking and landing on the moon. A third astronaut, who will remain in lunar orbit, awaits LM undocking strapped into his CSM couch.
At the start of an Apollo lunar mission, a Saturn V rocket launched a Command and Service Module (CSM) mothership and a Lunar Module (LM) moon lander. The mighty rocket's S-IVB third stage boosted the CSM and LM into a parking orbit about the Earth; then, about 90 minutes later, reignited to push itself, the CSM, and the LM out of Earth orbit toward the moon. This maneuver, called Trans-Lunar Injection (TLI), marked the real start of the lunar voyage.

After TLI, the CSM separated from the spent S-IVB, turned end-for-end, docked with the LM, and extracted it from the S-IVB. The S-IVB then vented propellants to change its course so that it would not interfere with CSM/LM navigation. Beginning with Apollo 13, the S-IVB was intentionally crashed on the moon to trigger seismometers left behind by previous Apollo expeditions.

As they neared the moon, the Apollo crew fired the CSM engine to slow down so that the moon's gravity could capture the joined Apollo spacecraft into lunar orbit. The LM then separated from the CSM bearing two of the astronauts and descended to the lunar surface using the engine in its Descent Stage.

After a maximum of three days on the moon, the Apollo lunar crew lifted off in the LM Ascent Stage using the Descent Stage as a launch pad. The astronaut in the CSM performed a rendezvous and docking with the Ascent Stage in lunar orbit to recover the moonwalkers - hence the name Lunar Orbit Rendezvous - then the crew discarded the LM Ascent Stage and fired the CSM engine to depart lunar orbit for Earth. Nearing Earth, they cast off the CSM's drum-shaped Service Module and reentered Earth's atmosphere in its conical Command Module (CM).

According to Joosten, a spacecraft that flew from Earth to the lunar surface, arrived on the moon with empty oxidizer tanks, and reloaded them for the trip home with liquid oxygen mined and refined from lunar regolith, could have about half the TLI mass of an equivalent LOR spacecraft. The Apollo 11 CSM, LM, and spent S-IVB stage had a combined mass at TLI of about 63 metric tons; the ILREC spacecraft and its spent TLI stage would have a mass of about 34 metric tons. This substantial mass reduction would permit use of a launch vehicle smaller than the Apollo Saturn V, potentially slashing lunar mission cost.

Lunar regolith is on average about 45% oxygen by weight. According to Joosten, literally dozens of lunar oxygen (LUNOX) extraction methods are known. He listed 14 as examples, including one, Hydrogen Ilmenite Reduction, for which the U.S. Patent Office had issued a patent to the U.S.-Japanese Carbotek/Shimizu consortium.

Joosten assumed that an automated LUNOX extraction process involving "solid-state high-temperature electrolysis" could produce 24 metric tons of LUNOX in cryogenic liquid form per year. He estimated that the process would need between 40 and 80 kilowatts of continuous electricity, and suggested that a nuclear reactor would be the best power-supply option. Such a reactor would have ample reserve power for charging electrically powered teleoperated mining vehicles and could supply crew electricity needs when astronauts were present.

Joosten acknowledged that ILREC emphasized technologies "in somewhat different areas than most exploration scenarios." Among these were teleoperated surface vehicles and surface mining and processing. On the other hand, the technological areas it emphasized had a "high degree of terrestrial relevance," a fact which, he argued, might prove to be a selling point for the new piloted lunar program.

Automated exploration missions would precede the new piloted lunar program. These might take the form of Lunar Scout orbiters and Artemis Common Lunar Landers, both JSC-proposed projects. The automated missions would have some "science linkages," Joosten explained, but would serve mainly to locate landing sites with abundant oxygen-rich regolith, perform ISRU experiments under real lunar conditions using real lunar materials, and map candidate landing sites to enable mission planners to certify them as safe for landings and rover traverses.

The NASA JSC engineer envisioned a three-phase piloted lunar program, though he provided details only for Phases 1 and 2. In Phase 1, three cargo landers would deliver equipment to the target landing site ahead of the first piloted mission. Flight 1 of Phase 1 would deliver the nuclear reactor on a teleoperated cart and the automated liquid oxygen production facility (the latter would remain attached to its lander); flight 2 would deliver teleoperated diggers, regolith haulers, oxygen tankers, and carts for auxiliary fuel-cell power and consumables resupply; and flight 3 would deliver a pressurized moon bus exploration rover and science equipment for the astronauts who would reach the moon on flight 4.

Following launch on an Energia rocket, translunar injection, and an Earth-moon voyage lasting up to about a week, a U.S.-built cargo lander bearing a self-deploying LUNOX regolith processing payload descends toward the lunar surface on a direct-descent trajectory. The lander is arranged horizontally, not vertically, to reduce the risk of tipping and, as important, to provide the astronauts who will follow it to the moon with easy access to its cargo. Image credit: NASA
After touchdown, the LUNOX regolith processing payload pivots into vertical operational position and deploys ramps so that teleoperated regolith hauler rovers (two are shown on the left side of the image) can reach its screen-covered input hopper. Meanwhile, a teleoperated tanker rover (right) collects and stores LUNOX in preparation for the arrival of a piloted ILREC spacecraft. Image credit: NASA
An Energia-launched cargo lander slowly lowers a U.S.-built pressurized moon bus lunar rover to the surface ahead of the arrival of the first two-person ILREC crew. Image credit: NASA
The one-way automated cargo landers, each rectangular in shape and capable of delivering 11 metric tons of payload to the moon's surface, would be assembled and packed in the U.S. and shipped to Russia in C-5 Galaxy or Antonov-124/225 transport planes, then launched on Energia rockets from Baikonur Cosmodrome, a Russian enclave in independent Kazakhstan. The Soviet Union's Energia heavy-lift rocket and Buran reusable shuttle were developed beginning in 1976 in response to the planned U.S. Space Shuttle. Energia replaced the Soviet answer to the U.S. Saturn V rocket, the N-1, which was cancelled in 1974 after four failed test flights. 

In contrast to the N-1, Energia flew successfully both times it was launched. Energia payloads were required to perform a short burn after they separated from the rocket so that they could achieve a stable orbit about the Earth. Polyus, launched 15 May 1987, did not orient itself properly ahead of the burn and did not reach orbit, while the unpiloted Buran completed a single orbit as planned and landed on a Baikonur runway on 15 November 1988. 

Based on data Russia provided to NASA, launch teams at Baikonur could prepare two Energia rockets for launch simultaneously. Three Energia launch pads were available - two originally built for the Soviet N-1 moon rocket and an all-new pad. Energia could launch a 5.5-meter-diameter canister containing a U.S.-built cargo lander attached to a Russian "Block 14C40" upper stage. Following an Earth-orbit insertion burn, the upper stage would perform a TLI burn, boosting the cargo lander toward the moon.

Shuttle-derived heavy-lift boosters would launch Joosten's piloted landers from the twin Kennedy Space Center (KSC) Complex 39 pads. The pads, monolithic Vehicle Assembly Building, and other KSC facilities, most of which were originally constructed in the 1960s for the Apollo moon program, were modified in the 1970s to serve the Space Shuttle. They would require new modifications to support the ILREC program; Joosten assured his readers, however, that no wholly new facilities would need to be constructed at the Florida spaceport.

Joosten considered both Shuttle-C and in-line Shuttle-derived launchers. The Shuttle-C design had a cargo module with attached Space Shuttle Main Engines (SSMEs) mounted on the side of a Shuttle External Tank (ET) in place of the delta-winged Shuttle Orbiter. The in-line design, a conceptual ancestor of the Space Launch System presently (2017) under development, would place the cargo module on top of a modified ET and three SSMEs underneath. The tank would have attached to its sides twin Advanced Solid Rocket Motors more powerful than their Space Shuttle counterparts. Joosten appears to have favored the Shuttle-C design.

The image above is slightly confusing: it displays a piloted ILREC lander and, below that, a conical TLI stage with three engines, but does not make clear that, except for the white, black, and gray conical crew capsule at the top, both lander and stage would be hidden from view under a streamlined white launch shroud. Missing from this illustration is the solid-propellant launch-escape system tower mounted on the crew capsule's nose. Image credit: NASA
A piloted ILREC lander descends toward a landing near the regolith processing lander and the teleoperated tanker rover. The aft compartment, located between the two rear landing gear, holds up to two tons of cargo. Image credit: NASA
Shortly after touchdown, the teleoperated tanker rover moves into position beside the ILREC crew lander and extends an umbilical so that it can refill the lander's empty liquid oxygen tanks with LUNOX for the trip home to Earth. Note the position of the crew hatch and two of the lander's four engines. Image credit: NASA
The Shuttle-derived heavy-lift rocket would launch the piloted lander, bearing an international crew and about two tons of cargo, into Earth orbit. About 4.5 hours after liftoff, following a systems checkout period, the TLI stage would place the piloted lander on a direct trajectory to the moon. The stage would then be cast off.

Joosten's crew lander design outwardly resembled the fictional "Eagle" transport spacecraft from the 1970s Gerry Anderson TV series Space: 1999. The crew compartment, a conical capsule modeled on the Apollo Command Module (but lacking a nose-mounted docking unit), would be mounted on the front of a horizontally oriented three-legged lander. The three landing legs would fold against the lander's belly beneath a streamlined shroud during ascent through Earth's lower atmosphere.

On the moon, the crew hatch would face downward, providing ready access to the surface via a ladder on the lander's single forward leg; on the launch pad, the hatch would permit horizontal access to the capsule interior much as did the Apollo CM hatch. The crew compartment windows would be inset into the hull and oriented to enable the pilot to view the landing site during descent. The crew spacecraft would land on and launch from the moon using the same set of four belly-mounted throttleable rocket engines.

During descent to the lunar surface, the engines would burn Earth oxygen and hydrogen. Soon after lunar touchdown, the lander would be reloaded with liquid oxygen from the automated lunar oxygen plant.

During return to Earth, Joosten's spacecraft would burn Earth hydrogen and lunar oxygen. The entire crew lander would lift off from the moon; only descent stages that delivered automated payloads would remain on the moon to clutter up the site. After a brief period in lunar parking orbit, the ILREC lander would ignite its four engines again to place itself on course for Earth.

Nearing Earth, the crew capsule would separate from the lander section and orient itself for reentry by turning its Apollo-style bowl-shaped heat shield toward the atmosphere. The lander section, meanwhile, would steer toward a reentry point well away from populated areas. The crew capsule would deploy a steerable parasail-type parachute. Joosten recommended that NASA recover the capsule on land - perhaps at Kennedy Space Center - to avoid the greater cost of an Apollo-style CM splashdown and water recovery. Most of the lander section would burn up during reentry.

The first piloted ILREC lander, with a U.S.-Russian crew of two on board, would spend two weeks on the moon. The crew would inspect the automated mining and oxygen production systems and explore using the moon bus rover. In Phase 1, the moon bus would be capable of traveling away from the crew lander landing site for two or three days at a time.

Several Phase 1 piloted missions to the site would be possible; alternately, NASA and Russia could skip immediately to Phase 2 - establishment of a temporary lunar outpost - after only a single Phase 1 piloted flight. In ILREC Phase 2, three more cargo flights would deliver to the same site a second moon bus rover, a rover support module with an attached airlock derived from Space Station hardware designs, consumables in a cart-mounted pressurizable Space Station-derived module, and science equipment.

An Energia-launched cargo lander would deliver the U.S.-built airlock/rover support node to the outpost site and lower it to the lunar surface. Astronauts in the pressurized moon bus rovers would drive it to a flat area using teleoperations techniques, then would use robot arms on their rovers to lower stilt-like supports. These would level and raise the airlock/node. After the airlock/node's wheels became raised off the ground, they would be removed, clearing the way for the twin rovers to "dock" with the node's two round side ports (one port is visible below the observation cupola just right of center). Image credit: NASA
Phase 2 ILREC temporary lunar outpost. Two pressurized rovers are docked tail-first to the support node, as is a pressurized consumables cart (at the end of the node opposite the airlock). Hanging regolith-filled bags on the node provide added protection from ionizing radiation. Wheels removed from the airlock/node are stacked to the left of the surface access gangway; they serve as spares for the pressurized moon bus rovers. A buried electrical cable (visible as a curved line in the lunar dirt running from center to lower right) leads toward a nuclear reactor (out of view). Image credit: NASA
Phase 2 outpost with components identified. The lower image is turned 90 degrees relative to the top image. Image credit: NASA
A piloted flight would then deliver a four-person crew for a six-week lunar surface stay. The crew would divide up into pairs, with each pair living in and operating a moon bus rover. The support module/airlock would include docking ports so that the two moon buses and the consumables module cart could link to it, forming a small outpost.

The moon buses would tow auxiliary power carts in Phase 2 to enable longer traverses across the lunar surface. The moon bus/cart combinations might travel in pairs along parallel routes or one moon bus might remain at the outpost while the other moon bus and its power cart ventured far afield. In the event that a moon bus rover failed beyond walking distance from the outpost and could not be repaired, the other moon bus could rescue its crew.

ILREC Phase 3 was poorly defined: it might see larger lunar crews venturing further afield, or NASA might change direction and use technology developed for the lunar program to put humans on Mars (perhaps still in partnership with Russia). Joosten identified the piloted moon lander crew capsule, Shuttle-derived heavy-lift rocket, pressurized moon bus rovers, and Energia as candidate Mars mission hardware. Both Energia and the Shuttle-derived rocket might be upgraded for piloted Mars missions; they might even be merged to create a single international heavy-lift rocket more powerful than either Energia or the Shuttle derivative.

Joosten envisioned that in Phases 1 and 2 Russia would pay for Energia and the Block 14C40 TLI stage, while NASA would pay for the Shuttle-derived rocket and its TLI stage, the crew and cargo landers, moon bus rovers and teleoperated carts, and lunar oxygen production systems. In exchange for Russia's participation, its cosmonauts would walk on the moon in the early years of the 21st century. If U.S.-Russia space cooperation were for any reason curtailed, NASA could continue the moon program by using Shuttle-derived launchers to launch moon-bound cargo - provided, of course, that U.S. policy makers determined that an all-U.S. moon program was worth the added cost.


Mir Hardware Heritage, NASA Reference Publication 1357, NASA Johnson Space Center Reference Series No. 3, David S. F. Portree, March 1995, pp. 168-170

"International Lunar Resources Exploration Concept," Kent Joosten, Low Cost Lunar Access Conference Proceedings, 1993, pp. 25-61; paper presented at the AIAA Low Cost Lunar Access conference, Arlington, Virginia, 7 May 1993

International Lunar Resources Exploration Concept, Presentation Materials, Kent Joosten, Exploration Programs Office, NASA Johnson Space Center, February 1993

Press Kit: Apollo 11 Lunar Landing Mission, NASA, 6 July 1969

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