09 September 2017

NASA Glenn Research Center's 2001 Plan to Land Humans on Mars Three Years Ago

August 2014. Image credit: NASA
In October 2001, at the 52nd International Astronautical Congress in the European aerospace center of Toulouse, France, nuclear propulsion engineers at NASA's Glenn Research Center (GRC) in Cleveland, Ohio, led by Stanley K. Borowski, Advanced Concepts Manager in GRC's Space Transportation Project Office, described a variant of NASA's 1998 Mars Design Reference Mission (DRM) based on Bimodal Nuclear-Thermal Rocket (BNTR) propulsion. The BNTR DRM concept, first described publicly in July 1998, evolved from nuclear-thermal rocket mission designs Borowski and his colleagues had developed during President George H. W. Bush's abortive Space Exploration Initiative (SEI), which got its start with a July 1989 presidential speech commemorating the 20th anniversary of Apollo 11, the first piloted moon landing mission.

This post contains more than its share of significant acronyms. As an aid to the reader, these are grouped alphabetically and defined at the bottom of the post, just ahead of the list of sources.

NASA's first Mars DRM, designated DRM 1.0 in 1997, was developed by a NASA-wide team during the 1992-1993 period. It was based on Martin Marietta's 1990 Mars Direct mission plan. SEI's demise temporarily halted NASA Mars DRM work in 1994. The civilian space agency resumed its Mars DRM studies after the announcement in August 1996 of the discovery of possible microfossils in martian meteorite ALH 84001. This enabled NASA planners to release their baseline chemical-propulsion DRM 3.0 in 1998. There was no official DRM 2.0, though a "scrubbed" (that is, mass-reduced) version of DRM 1.0 bears that designation in at least one NASA document.

Shortly thereafter, NASA's Johnson Space Center (JSC) in Houston, Texas, which led the DRM study effort, was diverted from DRM work by the in-house COMBO lander study (more on this below). Left largely to its own devices, NASA GRC developed a pair of DRM 3.0 variants: a solar-electric propulsion (SEP) DRM 3.0 and the BNTR DRM 3.0 discussed here.

In BNTR DRM 3.0, two unpiloted spacecraft would leave Earth for Mars during the 2011 low-energy Mars-Earth transfer opportunity, and a third, bearing the crew, would depart for Mars during the corresponding opportunity in 2014. Components for the three spacecraft would reach Earth orbit on six Shuttle-Derived Heavy-Lift Vehicles (SDHLVs), each capable of launching 80 tons into 220-mile-high assembly orbit, and in the payload bay of a winged, reusable Space Shuttle Orbiter, which would also deliver the Mars crew.

The SDHLV, often designated "Magnum," was a NASA Marshall Space Flight Center conceptual design. The Magnum booster would burn liquid hydrogen (LH2)/liquid oxygen (LOX) chemical propellants in its core stages and solid propellant in its side-mounted boosters. Magnum drew upon existing Space Shuttle hardware: its core stages were derived from the Space Shuttle External Tank and its twin solid-propellant rocket boosters were based on the Shuttle's twin Solid-Rocket Boosters.

The mighty Magnum was the conceptual ancestor of the equally conceptual Ares V and the Space Launch System, now under development. Image credit: NASA
SDHLV 1 would launch BNTR stage 1 with 47 tons of LH2 propellant on board. Each BNTR DRM mission would need three 28-meter-long, 7.4-meter-diameter BNTR stages. The BNTR stages would each include three 15,000-pound-thrust BNTR engines developed as part of a joint U.S./Russian research project in 1992-1993.

SDHLV 2 would boost an unpiloted 62.2-ton cargo lander into assembly orbit. The cargo lander would include a bullet-shaped Mars aerobrake and entry heat shield (this would double as the cargo lander's Earth launch shroud), parachutes for landing, a descent stage, a 25.8-ton Mars surface payload including an in-situ resource utilization (ISRU) propellant factory, four tons of "seed" LH2 to begin the process of manufacturing propellants on Mars, and a partly fueled Mars Ascent Vehicle (MAV) made up of a conical Earth Crew Return Vehicle (ECRV) capsule and an ascent stage. The cargo and habitat lander engines would burn liquid methane fuel and LOX.

SDHLV launch 3, identical to SDHLV launch 1, would place into assembly orbit BNTR stage 2 containing 46 tons of LH2 propellant. SDHLV launch 4 would place the unpiloted 60.5-ton habitat lander into assembly orbit. The habitat lander would include a Mars aerobrake & entry shield/launch shroud identical to that of the cargo lander, parachutes, a descent stage, and a 32.7-ton payload including the crew's Mars surface living quarters.

The BNTR stage forward section would include chemical thrusters. These would provide maneuvering capability so that the stages could dock with the habitat and cargo landers in assembly orbit. During flight to Mars, the thrusters would provide each stage/lander combination with attitude control.

2011: the unmanned BNTR 1 stage/cargo lander and BNTR 2 stage/habitat lander spacecraft orbit the Earth prior to departure for Mars. Image credit: NASA
The BNTR 1/cargo lander combination would have a mass of 133.7 tons, while the BNTR 2/habitat lander combination would have a mass of 131 tons. Both combinations would measure 57.5 meters long. As the 2011 launch window for Mars opened, the BNTR stages would fire their engines to depart assembly orbit for Mars.

Each BNTR engine would include a nuclear reactor. When moderator elements were removed from its nuclear fuel elements, the reactor would heat up. To cool the reactor so that it would not melt, turbopumps would drive LH2 propellant through it. The reactor would transfer heat to the propellant, which would become an expanding very hot gas and vent through an LH2-cooled nozzle. This would propel the spacecraft through space.

Following completion of Earth-orbit departure, the BNTR engine reactors would switch to electricity-generation mode. In this mode, they would operate at a lower temperature than in propulsion mode, but would still be capable of heating a working fluid that would drive three turbine generators. Together the generators would make 50 kilowatts of electricity. Fifteen kilowatts would power a refrigeration system in the BNTR stage that would prevent the LH2 it contained from boiling and escaping.

Much like the LH2 propellant in BNTR propulsion mode, the working fluid would cool the reactor; unlike the LH2, however, it would not be vented into space. After leaving the turbine generators, it would pass through a labyrinth of tubes in radiators mounted on the BNTR stage to discard leftover heat, then would cycle through the reactors again. The cycle would repeat continuously throughout the journey to Mars.

2012: Cargo lander/Mars Ascent Vehicle Landing. Image credit: NASA
As Mars loomed large ahead, the turbine generators would charge the lander batteries. The BNTR stages would then separate and fire their engines to miss Mars and enter a safe disposal orbit around the Sun. The landers, meanwhile, would aerobrake in Mars's upper atmosphere. The habitat lander would capture into Mars orbit and extend twin solar arrays to generate electricity. The cargo lander would capture into orbit, then fire six engines to deorbit and enter the atmosphere a second time. After casting off its heat shield, it would deploy three parachutes. The engines would fire again, then landing legs would deploy just before touchdown. The GRC engineers opted for a horizontal landing configuration; this would, they explained, prevent tipping and provide the astronauts with easy access to the lander's cargo.

As illustrated in the cargo lander image above and the MAV launch image below, the four MAV engines would serve double-duty as cargo lander engines. In addition to saving mass by eliminating redundant engines, this would test-fire the engines before the crew used them as MAV ascent engines.

2012: Automated propellant manufacture for MAV ascent begins. Image credit: NASA
The cargo lander would touch down on Mars with virtually empty tanks. After touchdown, a teleoperated cart bearing a nuclear power source would lower to the ground and trundle away trailing a power cable. Controllers on Earth would attempt to position it so that the radiation it emitted would not harm the astronauts (for example, behind a sand dune or boulder pile). The reactor's first job would be to power the lander's ISRU propellant plant, which over several months would react the seed hydrogen brought from Earth with martian atmospheric carbon dioxide in the presence of a catalyst to produce 39.5 tons of liquid methane fuel and LOX oxidizer for the MAV ascent engines.

SDHLV launch 5, identical to SDHLV launches 1 and 3, would mark the start of launches for the 2014 Earth-Mars transfer opportunity. It would place BNTR stage 3 into assembly orbit with about 48 tons of LH2 on board. Because it would propel a piloted spacecraft, its BNTR engines would require a new design feature: each would include a 3.24-ton shield to protect the crew from the radiation it produced while in operation. The shields each would create a conical radiation "shadow"; the radiation shadows would overlap to create a safe zone in which the crew would remain while they were inside or close to their spacecraft.

2013: the BNTR 3 stage and the first Crew Transfer Vehicle components dock automatically in Earth orbit. Image credit: NASA
Thirty days after SDHLV launch 5, SDHLV launch 6 would place into assembly orbit a 5.1-ton spare Earth Crew Return Vehicle (ECRV) attached to the front of an 11.6-ton truss. A 17-meter-long tank with 43 tons of LH2 and a two-meter-long drum-shaped logistics module containing 6.9 tons of contingency supplies would nest along the truss's length. BNTR stage 3 and the truss assembly would rendezvous and dock, then propellant lines would automatically link the truss tank to BNTR stage 3.

A Shuttle Orbiter carrying the Mars crew and a 20.5-ton deflated Transhab module would rendezvous with the BNTR stage 3/truss combination one week before the crew's planned departure for Mars. Following rendezvous, the spare ECRV would undock from the truss and fly automatically to a docking port in the Space Shuttle payload bay. Astronauts would then use the Orbiter's robot arm to hoist the Transhab from the payload bay and dock it to the front of the truss in the spare ECRV's place.

2014: Crew and a deflated Transhab arrive on board a Space Shuttle Orbiter to complete Crew Transfer Vehicle assembly. Image credit: NASA
The Mars astronauts would enter the spare ECRV and pilot it to a docking at a port on the Transhab's front, then enter the cylindrical Transhab's solid core and inflate its fabric-walled outer volume. The inflated Transhab would measure 9.4 meters in diameter. Unstowing floor panels and furnishings from the core and installing them in the inflated volume would complete assembly. Transhab, truss, and BNTR stage 3 would make up the 64.2-meter-long, 166.4-ton Crew Transfer Vehicle (CTV).

The CTV's truss-mounted tank and BNTR stage 3 would hold 90.8 tons of LH2 at the start of CTV Earth-orbit departure on 21 January 2014. The truss tank would provide 70% of the propellant needed for departure. In the most demanding departure scenario, the BNTR engines would fire twice for 22.7 minutes each time to push the CTV out of Earth orbit toward Mars.

2014: Crew Transfer Vehicle departs Earth orbit. Image credit: NASA
Transhab cutaway (weightless design). Floor and ceiling would be reversed in the NASA Glenn artificial-gravity design. "Down" would thus be toward the top of this image, where the airlock and Earth Crew Return Vehicle capsule would be located. Image credit: NASA
Following Earth-orbit departure, the crew would jettison the empty truss tank and use small chemical-propellant thrusters to start the CTV rotating end over end at a rate of 3.7 rotations per minute. This would create acceleration equal to one Mars gravity (38% of Earth gravity) in the Transhab module. Artificial gravity was a late addition to BNTR DRM 3.0; it made its first appearance in a June 1999 paper, not in the original July 1998 paper describing BNTR DRM 3.0.

In artificial-gravity mode, "down" would be toward the spare ECRV on the CTV's nose; this would make the Transhab's forward half its lower deck. Halfway to Mars, about 105 days out from Earth, the astronauts would stop rotation and perform a course-correction burn using the attitude-control thrusters. They would then resume rotation for the remainder of the trans-Mars trip.

The CTV would arrive in Mars orbit on 19 August 2014. The crew would halt rotation, then three BNTR engines would fire for 12.3 minutes to slow the spacecraft for Mars orbit capture. In its loosely bound elliptical Mars orbit, the spacecraft would circle the planet once per 24.6-hour martian day.

2014: Crew Transfer Vehicle arrival in Mars orbit. Image credit: NASA
The crew would pilot the CTV to rendezvous with the habitat lander waiting in Mars orbit, taking care to place it in the CTV's radiation shadow. If the cargo lander on the surface or the habitat lander in Mars orbit malfunctioned while awaiting the crew's arrival, then the crew would remain in the CTV in Mars orbit until Mars and Earth aligned for the flight home (a wait time of 502 days). They would survive by drawing upon contingency supplies in the drum-shaped logistics module attached to the truss.

If the orbiting habitat lander and landed cargo lander checked out as healthy, however, then the crew would fly the spare ECRV to a docking port on the habitat lander's side. After discarding the spare ECRV and the habitat solar arrays, they would fire the habitat lander's engines, enter the martian atmosphere, and land near the cargo lander.

The habitat lander's horizontal configuration would provide the astronauts with ready access to the martian surface. After the historic first footsteps on Mars, the astronauts would inflate a Transhab-type habitat attached to the side of the habitat lander, run a cable from the habitat lander to the nuclear power source cart, unload at least one unpressurized crew rover, and commence a program of Mars surface exploration that would, if all went as planned, last for nearly 17 months.

In case of hardware failure or other emergency, the crew could retreat to the MAV and return early to the orbiting CTV. They would, however, have to wait in Mars orbit until Mars and Earth aligned to permit a minimum-energy Mars-Earth transfer (that is, until the originally planned end of their stay at Mars).

2014-2015: The first Mars campsite. In the foreground is the habitat lander with inflated Transhab surface habitat; in the background, the nuclear power source cart and the cargo lander with Mars Ascent Vehicle. Image credit: NASA
2014-2015: Exploring Mars with a crew rover and two teleoperated robot rovers, one small and one large. Image credit: NASA
2014-2015: Drilling for water, geologic history, and, just possibly, life. Image credit: NASA
2015: Mars Ascent Vehicle liftoff. Image credit: NASA
Near the end of the surface mission, the unmanned CTV would briefly fire its nuclear engines to trim its orbit for the crew's return. The MAV bearing the crew and about 90 kilograms of Mars samples would then lift off. Taking care to remain within the the radiation shadows of the CTV's BNTR engines, it would dock at the front of the Transhab, then the astronauts would transfer to the CTV. They would cast off the spent MAV ascent stage, but would retain the MAV ECRV for Earth atmosphere reentry.

The CTV would leave Mars orbit on 3 January 2016. Prior to Mars orbit departure, the astronauts would abandon the contingency supply module on the truss to reduce their spacecraft's mass so that the propellant remaining in BNTR stage 3 would be sufficient to launch them home to Earth. They would then fire the BNTR engines for 2.9 minutes to change the CTV's orbital plane, then again for 5.2 minutes to escape Mars and place themselves on course for Earth.

Soon after completion of the second burn, the crew would fire attitude-control thrusters to spin the CTV end-over-end to create acceleration equal to one Mars gravity in the Transhab. About halfway home they would stop rotation, perform a course correction, then resume rotation. Flight home to Earth would last 190 days.

2016: Return to Earth. Image credit: NASA
Near Earth, the crew would stop CTV rotation for the final time, enter the MAV ECRV with their Mars samples, and undock from the CTV, again taking care to remain in the BNTR engine radiation shadows as they moved away. The abandoned CTV would fly past Earth and enter solar orbit. The MAV ECRV, meanwhile, would reenter Earth's atmosphere on 11 July 2016.

The authors compared their Mars plan with the baseline chemical-propulsion DRM 3.0 and with the NASA GRC SEP DRM 3.0. They found that their plan would need eight vehicle elements, of which four would have designs unique to BNTR DRM 3.0. The baseline DRM 3.0, by contrast, would need 14 vehicle elements, 10 of which would be unique, and SEP DRM 3.0 would need 13.5 vehicle elements, 9.5 of which would be unique. BNTR DRM 3.0 would require that 431 tons of hardware and propellants be placed into Earth orbit; the baseline DRM 3.0 would need 657 tons and SEP DRM 3.0, 478 tons. Borowski and his colleagues argued that fewer vehicle designs and reduced mass would mean reduced cost and mission complexity.

The BNTR DRM 3.0 variant became the basis for DRM 4.0, which was developed during NASA-wide studies in 2001-2002 (though NASA documents occasionally back-date DRM 4.0 to 1998, when BNTR DRM 3.0 was first proposed). DRM 4.0 differed from BNTR DRM 3.0 mainly in that it adopted a "Dual Lander" design concept developed as part of JSC's 1998-1999 COMBO lander study. COMBO was the brainchild of William Schneider, NASA JSC Engineering Directorate boss.

Dual Lander concept. The lander in the foreground is the habitat; the background lander is the Mars Descent/Ascent Vehicle. Image credit: NASA
The Dual Lander concept grew from COMBO's main design guideline, which was to develop a low-mass "Apollo-style" piloted Mars landing mission. A major change from past Mars DRMs was no reliance on ISRU. As in BNTR DRM 3.0, two cargo missions would leave Earth one minimum-energy Earth-Mars transfer opportunity ahead of the crew; in DRM 4.0, however, these would take the form of a Mars lander that would also include an ascent vehicle for returning the crew to the CTV in Mars orbit and a cargo lander with an inflatable donut-shaped habitat. The former could by itself support a short-stay (~30-day) Mars surface mission; the latter would enable a Mars surface stay of more than 400 days.

In 2008, a decade after BNTR DRM 3.0 first was made public, NASA released a version of DRM 4.0 modified to use planned Constellation Program hardware (for example, the Ares V heavy-lift rocket in place of the Magnum and the Orion Multi-Purpose Crew Vehicle in place of the ECRVs). The space agency dubbed the new DRM Design Reference Architecture (DRA) 5.0.

The DRA 5.0 Mars plan acknowledged that, largely as a result of the 1 February 2003 Columbia accident, the Space Shuttle would be retired after the remaining Orbiters - Endeavour, Discovery, and Atlantis - completed their part of the task of building the International Space Station. The last Space Shuttle mission, STS-135, took place in July 2011.

DRA 5.0 also saw the return of ISRU. A Descent/Ascent Vehicle (DAV) and a Surface Habitat (SHAB) would capture into Mars orbit in the first minimum-energy Earth-Mars transfer opportunity. The DAV would descend, land, and begin making propellants for its ascent stage. The SHAB would loiter in orbit awaiting arrival of a crew on board a Mars Transfer Vehicle (MTV) launched from Earth during the second Earth-Mars transfer opportunity of the mission. The crew would transfer to the SHAB in an Orion/service module and land on Mars near the DAV. After a stay on Mars lasting more than 400 days, they would lift off in the DAV ascent stage, dock with the waiting MTV, and return to Earth.

Though DRA 5.0 exerts influence on current NASA planning, the precise form a piloted Mars mission will eventually take remains unclear at this writing. NASA increasingly has shifted its attention toward finding low-cost stepping stones that could lead to a piloted Mars landing in 2033. A crew-tended - that is, not permanently staffed - Deep Space Gateway space station in cislunar space, for example, could be established by 2026 through a series of Orion missions launched using the Space Launch System (SLS) heavy-lift rocket (SLS replaced Ares V in 2010). Other possible interim steps toward Mars include an SLS-launched robotic Mars sample-return mission in the mid-2020s and a piloted mission to Mars orbit in 2030 using a Deep Space Transport based partly on Deep Space Gateway hardware.


BNTR = Bimodal Nuclear Thermal Rocket
CTV = Crew Transfer Vehicle
DAV = Descent/Ascent Vehicle
DRA = Design Reference Architecture
DRM = Design Reference Mission
ECRV = Earth Crew Return Vehicle
GRC = Glenn Research Center
ISRU = In-Situ Resource Utilization
JSC = Johnson Space Center
LH2 = liquid hydrogen
LOX = liquid oxygen
MAV = Mars Ascent Vehicle
MTV = Mars Transfer Vehicle
SDHLV = Shuttle-Derived Heavy-Lift Vehicle
SEI = Space Exploration Initiative
SEP = Solar-Electric Propulsion
SHAB = Surface Habitat
SLS = Space Launch System


"Bimodal Nuclear Thermal Rocket (NTR) Propulsion for Power-Rich, Artificial Gravity Human Exploration Missions to Mars," IAA-01-IAA.13.3.05, Stanley K. Borowski, Leonard A. Dudzinski, and Melissa L. McGuire; paper presented at the 52nd International Astronautical Congress in Toulouse, France, 1-5 October 2001

"Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using 'Bimodal' NTR and LANTR Propulsion," AIAA-98-3883, Stanley K. Borowski, Leonard A. Dudzinski, and Melissa L. McGuire; paper presented at the 34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit in Cleveland, Ohio, 13-15 July 1998

"Artificial Gravity Vehicle Design Option for NASA's Human Mars Mission Using 'Bimodal' NTR Propulsion," AIAA-99-2545, Stanley K. Borowski, Leonard A. Dudzinski, and Melissa L. McGuire; paper presented at the 35th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit in Los Angeles, California, 20-24 June 1999

NASA Exploration Team (NEXT) Design Reference Missions Summary, NASA, 12 July 2002 [draft]

"Enabling Human Deep Space Exploration with the Deep Space Gateway," Tim Cichan, Bill Pratt, and Kerry Timmons, Lockheed Martin; presentation to the Future In-Space Operations telecon, 30 August 2017

More Information

A Forgotten Pioneer of Mars Resource Utilization (1962-1963)

Two For The Price of One: 1980s Piloted Missions With Stopovers at Mars and Venus (1969)

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

Humans on Mars in 1995! (1980-1981)

Bridging the Gap Between Space Station and Mars: The IMUSE Strategy (1985)

The Collins Task Force Says Aim for Mars (1987)

Sally Ride's Mission to Mars (1987)

Footsteps to Mars (1993)

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, finally, the lander would fall free of the top aeroshell and ignite its landing motors for terminal 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 transmit 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 loop-wheel 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
Following lander touchdown, the rovers would each unfold their various appendages and stand up on their articulated legs. The landers, meanwhile, would each extend a pair of ramps. Controllers on Earth would then command the rovers to crawl forward and down the ramps on their loop-wheel treads.

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

"The Case for Life on Mars," A. Chaikin, Air & Space Smithsonian, February/March 1991, pp. 63-71

More Information

Robot Rendezvous at Hadley Rille (1968) (AAP & drivable/robotic lunar rover)

The Russians are Roving! The Russian are Roving! A 1970 JPL Plan for the 1979 Mars Rover (Soviet robotic exploration plans & JPL's response)

Safeguarding the Earth from Martians: The Antaeus Report (1978-1981) (Mars Sample Return & planetary protection & early Shuttle optimism)

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)