Rabu, 10 Mei 2017
JPL's MSR spacecraft, which used the Mars Orbital Rendezvous mission mode, consisted of an orbiter for transporting a lander to Mars and returning the Mars samples to Earth, a rover for sample collection, a Mars Ascent Vehicle (MAV) for boosting the collected samples to Mars orbit for retrieval by the orbiter, and a lander for delivering the rover and MAV to Mars's surface.
In May 1998, JPL rover engineer Brian Wilcox proposed a possible solution: replace the MSR mission design's massive (about 500 kilograms) liquid-propellant MAV with a low-mass solid-propellant MicroMAV. The following month, JPL engineers Duncan MacPherson, Doug Bernard, and Bill Layman began a preliminary study to attempt to validate Wilcox's concept. As part of their effort, they held a "mini-workshop" at which they consulted with space industry propulsion engineers. By September, MacPherson was ready to present his group's findings to the second meeting of the NASA-appointed Mars Architecture Team (MAT).
Wilcox had envisioned an alternative MSR scenario in which a large rover would carry and launch his 20-kilogram MicroMAV. MacPherson, Bernard, and Layman proposed a MAV that burned solid propellants but had a more realistic estimated mass of 110 kilograms. This would, they found, require a return to a more traditional MSR scenario in which the MAV would lift off from a stationary lander. A rover would collect samples and deliver them to the MSR lander, which would load them into a container in the MAV's third stage.
Wilcox assumed that, during first-stage flight, airflow over four canted fins on his MicroMAV's first stage could spin it to provide gyroscopic stability. MacPherson, Bernard, and Layman judged, however, that martian air was not dense enough for canted fins to be effective. Prior to first stage ignition, thus, a spin table on the MSR lander would spin their MAV about its long axis 300 times per minute to provide gyroscopic stability. The first stage motor would then ignite and hurl the MAV skyward at from six to 10 gravities of acceleration.
Industry experts attending the mini-workshop had told MacPherson, Bernard, and Layman that metal-based solid propellant yields molten slag when it burns. In a rapidly spinning motor, the centrifugal force causes the slag to adhere to the nozzle, producing unpredictable mass imbalances. These could destablilize the ascending rocket, causing it to tumble out of control. A high spin rate could also cause uneven solid propellant burning. MacPherson told the MAT that metal-free solid propellant would eliminate both problems, though at the price of reduced motor performance.
After first stage burnout, a small despin motor would slow the MAV's rate of spin to 20 revolutions per minute. The MAV would then coast to an altitude of 90 kilometers. Wilcox assumed no active attitude control during the coast, but MacPherson, Bernard, and Layman invoked cold-gas attitude control thrusters to compensate for winds and to orient the MAV accurately for the second stage burn.
An inertial measurement unit and a sun sensor would provide data to the thruster guidance system and to a timer that would govern subsequent MAV operations. The spent first stage would detach one second after timer activation, then the second stage motor would ignite one second after that. Second stage acceleration would peak at 35 times the pull of Earth's gravity just before burnout. The second stage would boost the MAV's apoapsis (orbit high point) to 300 kilometers above Mars, then would separate two minutes after timer start.
Wilcox gave little attention to the MicroMAV's role in preventing biological contamination of Mars (forward contamination) or Earth (back contamination). MacPherson noted that the second-stage motor's trajectory after separation would take it back into Mars's atmosphere, thus eliminating it as a possible source of back contamination.
As in the Wilcox design, the MacPherson/Bernard/Layman third stage motor nozzle would point forward during first stage and second stage flight, ensuring that it would point aft when the gyro-stabilized MAV attained apoapsis halfway through its first orbit. The timer would ignite the third stage motor 50 minutes after timer start; if all had functioned as planned up to that point, this would coincide with apoapsis. The brief burn would raise the MAV's periapsis (orbit low point) out of the atmosphere to an altitude of at least 300 kilometers.
As its last act, the timer would fire a motor that would halt the MAV's spin so that the orbiter could more easily capture it. The waiting orbiter would then maneuver to retrieve the MAV third stage and the precious Mars samples it carried.
MacPherson, Bernard, and Layman found that minor guidance errors, motor performance variations, and the vagaries of Mars's atmosphere could affect the MAV's final orbital parameters and thus the magnitude of the maneuvers the orbiter would need to perform to rendezvous with it. Wilcox, always optimistic about his MicroMAV's capabilities, had calculated that compensating for orbital uncertainties would require that the orbiter carry only enough propellants to enable velocity changes totaling about 100 meters per second. MacPherson's team, by contrast, estimated a possible MAV periapsis range of 300-to-500 kilometers, an apoapsis range of 600-to-800 kilometers, and an orbital inclination range spanning one degree. In the worst case scenario, this would mean that the MSR orbiter might need to make velocity changes totaling about 260 meters per second.
The MacPherson group's results might have thrown cold water on the MicroMAV concept. A 110-kilogram MAV was, however, an improvement over one with a mass of 500 kilograms. Even before they finished their work, JPL adopted the small solid-propellant MAV as part of its baseline MSR mission design.
One of O'Neil's first initiatives in his new role was a pair of workshops aimed at generating fresh ideas for JPL's Mars Sample Return (MSR) mission, which had become mired in fiscal and engineering problems. In its April 1998 iteration, a single launch vehicle boosted the MSR spacecraft to Mars in late 2004. By the start of the first MSR workshop in July, a redesign effort begun in June had yielded a baseline MSR plan that split the orbiter and lander between two rockets launched in August and September 2005. The lander would carry the Mars Ascent Vehicle (MAV), which would launch the Mars sample to orbit for retrieval by the orbiter and return to Earth. The 512-kilogram liquid-propellant MAV was overweight, contributing to mass limitations which meant that only a small, short-range sample-collecting rover could be included in the mission.
In his presentation to the first MSR workshop, Brian Wilcox, a JPL rover engineer and former model rocketry enthusiast, described a possible alternative to the baseline mission's liquid-propellant MAV based on the U.S. Navy's 1958 PILOT microsatellite launcher design. His "MicroMAV" was a 20-kilogram solid-propellant rocket with no moving parts. Wilcox noted that, unlike liquid propellants, solid propellants would not freeze during the frigid martian night.
Wilcox envisioned an MSR similar to one he proposed in 1989, in which a rover with six wheels and a top-mounted solar array would carry the three-stage MicroMAV with it while exploring Mars. The MicroMAV would ride slung horizontally along one of the rover's sides. The rover would collect an unspecified quantity of rocks and dirt and load them into the sample canister in the MicroMAV's third stage, then would pivot the MicroMAV onto the top of the solar array and point its nose skyward. The MicroMAV would then ignite its first stage motor.
The first stage, which would loft the MicroMAV above most of Mars's atmosphere, would have a total mass at ignition of 9.75 kilograms, of which 7.8 kilograms would comprise solid propellant. It would include four fins and a horizon sensor. The fins would be canted slightly so that the thin martian air rushing past them during ascent would spin the MicroMAV about its long axis for gyroscopic stabilization.
After first stage burnout, the MicroMAV would coast upward, still spinning about its long axis. As it neared the top of its trajectory, its nose would begin to tip downward toward the horizon. The horizon sensor would alternately "see" the sky above and the ground below.
When the sensor tallied a pre-set number of rotations, it would trigger second stage ignition. This would also discard the first stage. The second stage, which would supply most of the MicroMAV's orbital velocity, would have a mass of 9.4 kilograms with 7.8 kilograms of propellant. After second stage burnout and separation, the MicroMAV third stage would be in Mars orbit; its periapsis (orbit low point) would, however, remain within Mars's atmosphere. Second stage burnout would thus trigger a timer designed to ignite the third stage motor.
The tiny 0.85-kilogram third stage would include 0.05 kilograms of propellant and the Mars sample. During first and second stage flight, its motor would point forward. Because it would be spinning like a gyroscope, it would remain pointed in the same direction following second stage separation. This would mean that, half a Mars orbit later, the motor would point away from its direction of motion. At that same moment, the MicroMAV would attain apoapsis (orbit high point) and the timer would reach zero. The third stage engine would then ignite to raise the MicroMAV's periapsis to a safe altitude.
Third stage ignition would also ignite a "pyrotechnic layer" that would heat its exterior "white-hot for an instant." This would destroy any martian microbes that might have hitched a ride on the third stage and would also solder shut the sample canister to prevent the escape of any contaminants inside.
The grapefruit-sized MicroMAV sample canister would be entirely passive, with neither a radio beacon nor a flashing light to aid the orbiter in locating it. The orbiter would begin looking for the MicroMAV from a position about 100 kilometers above its orbit. For 18% of its orbit, the canister would be sunlit but set against Mars's nightside as seen from the orbiter. At such times, the orbiter would point its wide-angle imager toward the MicroMAV's predicted position and image the area several times to enable controllers on Earth to determine the MicroMAV sample canister's orbit. Wilcox estimate that controllers using orbiter images would need no more than 31 hours to locate the MicroMAV.
The MicroMAV concept excited much interest among JPL engineers. Though further study revealed the MicroMAV MSR scenario to be unworkable in the form Wilcox described - for example, JPL quickly abandoned rover launch in favor of a more conventional launch from a fixed lander (image above) - the concept of a simplified solid-propellant MAV profoundly influenced subsequent JPL MSR planning.
In a paper presented at the Fifth International Astronautical Federation Congress in 1954, Stuhlinger pitched interplanetary travel using low-thrust ion (electric) propulsion. The spacecraft design he proposed comprised three major parts: the crew/payload compartment at the ship's center; a 146.4-ton multi-unit solar-electric power system; and a multi-chamber low-thrust ion drive system.
Stuhlinger provided no details about the layout of his ship's crew/payload compartment, other than that it would carry up to 50 tons of crew and cargo. He did, however, offer abundant details on his solar-electric power and ion drive systems.
The former would include two 350-meter-wide "wings," each comprising 19 independent electricity-generating "sub-units." A dish-shaped mirror 50 meters wide would form the largest component of each 4400-kilogram sub-unit. Stuhlinger wrote that his spacecraft would gain speed very slowly, accelerating at a rate equal only to about 1/1000th of Earth's surface gravity. At such a low rate of acceleration, a fork dropped in the ship's messroom would need more than five minutes to strike the floor. The low acceleration would mean that the mirrors would have no need of robust construction; they might comprise "thin aluminum foil with a very light supporting frame."
Each 450-kilogram, 2000-square-meter mirror would concentrate sunlight onto a boiler, causing a working fluid within it to turn to steam. The steam would drive a turbine, which would in turn drive a generator capable of producing 200 kilowatts of electricity. The steam, meanwhile, would enter a disk-shaped radiator cooler and condense back into fluid. Boiler, turbine/generator, and cooler would revolve together as a unit, completing one revolution every 10 seconds. This would generate acceleration that would cause the working fluid to flow to the cooler's outer rim, from which it would be pumped back to the boiler.
The multi-unit solar-electric power system would have built-in redundancy, Stuhlinger noted. Even if a large "meteor" hit the ship, he wrote, "the total loss of one or two sub-units would mean only a minor reduction of the capacity of the power plant."
Stuhlinger rejected an "atomic pile" as a heat source; in addition to having a mass of "hundreds of tons," a reactor would emit harmful radiation that would demand heavy shielding and make in-flight repair difficult. He added, however, that "an atomic pile will be a very promising power source for an electrically propelled space ship as soon as the mass problem, the shielding problem, and the maintenance problem have been solved satisfactorily."
The third major part of Stuhlinger's ship, the ion drive, would consist of many clustered thrust chambers. Within each, electricity from the solar-electric power system would ionize cesium or rubidium vapor using heated platinum grids and paired positive and negative electrodes. The cesium or rubidium ions would then depart the chamber through an opening at a large fraction of the speed of light to push the ship through space.
Stuhlinger wrote that cesium would be a more efficient propellant than rubidium. A cesium-fueled ship would need only 1833 thrust chambers to produce as much thrust as a rubidium-fueled ship with 2200 chambers. He noted, however, that cesium is "a rare element which might not be available in quantities as required for space ships."
Despite its large number of thrust chambers, Stuhlinger's ion drive would generate at most nine kilograms of thrust. This would, however, be applied continuously for long periods. Assuming no interference from planetary or solar gravity, Stuhlinger's ship could in a year travel 183 million kilometers in a straight line and reach a velocity of 12 kilometers per second.
Stuhlinger calculated that his ship would need just 18.6 tons of rubidium to accelerate continuously for one year. Even with its elaborate solar-electric power and ion drive systems, his ship's mass would total just 280 tons. To reach the same 12-kilometer-per-second velocity, a chemical-propulsion Mars spaceship would need a mass of about 820 tons, most of which would comprise propellants. For his calculations, Stuhlinger assumed that the chemical ship's rocket motors would burn nitric acid oxidizer and hydrazine fuel. He also assumed that both the ion and chemical Mars ships would be assembled in Earth orbit from components launched atop chemical-propulsion cargo rockets; his ship's lesser mass meant that it would need about a third as many cargo launches for assembly as would its chemical counterpart.
An ion drive spaceship would, of course, not travel between Earth and Mars in a straight line; it would instead gradually spiral out of Earth orbit into solar orbit, follow a curved course around the Sun to Mars, capture into a distant Mars orbit, spiral gradually down to a low Mars parking orbit, spiral out of Mars orbit, follow a curved course around the Sun back to Earth, capture into distant Earth orbit, and gradually spiral down to low Earth parking orbit. Halfway to Mars and again halfway to Earth the ship would turn end for end to face its thrust chambers forward and begin a slow deceleration. Stuhlinger determined, nonetheless, that his low-thrust solar-electric ion drive spaceship could travel from Earth orbit to Mars orbit and back in just two or three years; that is, in approximately the same period of time that a high-thrust chemical spaceship would need.
Stuhlinger did not call his spaceship the Cosmic Butterfly; that name originated with Frank Tinsley (1899-1965), an artist, cartoonist, and author famed for his futuristic technical illustrations. Tinsley used the term "gigantic butterfly" in reference to Stuhlinger's design in a 1956 article in Modern Mechanix magazine. The illustration at the top of this blog post, which Tinsley painted in 1959 for an American Bosch Arma Corporation advertisement titled "Cosmic Butterfly," depicts a ship little different from Stuhlinger's 1954 design.
Senin, 08 Mei 2017
At the same time, the U.S. had developed new priorities - for example, prosecuting the war in Indochina. This was reflected in changes in NASA's budget and workforce. The space agency's budget peaked in 1967 at nearly $6 billion, or about 0.9% of Gross National Product (GNP) - the largest fraction of GNP NASA ever attained. By the time of the triumphant first Apollo moon landing on July 20, 1969, NASA's budget had been pared to $4 billion. In 1965, the NASA and NASA contractor workforce totaled 420,000 people across the United States; by the end of Fiscal Year 1969, this had slumped to 220,000, triggering an “aerospace depression.” States like California and Florida, where space contractors were concentrated, bore the brunt of the cuts. With this background in mind, NASA’s report to the STG began with the following prescient words:
At the moment of its greatest triumph, the space program of the United States faces a crucial situation. Decisions made this year will affect the course of space activity for decades to come. . .NASA argued that the Nixon Administration had before it a unique opportunity for greatness. Nixon could, if he so chose, become the President known for launching America to the planets.
This Administration has a unique opportunity to determine the long-term future of the Nation's space progress. We recommend that the United States adopt as a continuing goal the exploration of the solar system. . .To focus our developments and integrate our programs, we recommend that the United States prepare for manned planetary expeditions in the 1980s.In an effort to stem its critics, who increasingly asserted that NASA's programs and goals were irrelevant to America's many pressing problems, the NASA report devoted considerable attention to Earth-centered benefits of spaceflight - for example, the potential public health benefits of medical experiments performed aboard Earth-orbiting space stations - and claimed that "the national civilian space effort has contributed $35 billion in goods and services to the U.S. economy." At the time, a large Earth-orbiting space station was NASA’s top priority as an immediate post-Apollo goal. Recognizing that the Soviet space threat no longer carried the weight that it had a decade earlier, and knowing the Nixon Administration’s own geopolitical preferences, NASA proposed spaceflight as a vehicle for international cooperation, not competition.
The report then asserted that NASA should receive sufficient resources in the 1970s to build on the capabilities it developed in the 1960s, a period during which
the American space program progressed from the 31-pound Explorer 1 in earth orbit to Apollo spacecraft weighing 50 tons sent out to the moon; [and] from manned flights of a few thousand miles and 15-minute duration to the 500,000 mile round-trip 8-day [Apollo 11] mission which landed men on the moon and returned them safely to earth.Continued manned lunar exploration after Apollo would, the report explained, "expand man's domain to include the moon." Large space stations and a space transportation system comprising reusable vehicles - a winged shuttle for delivering crews and supplies to the Earth-orbiting station, a nuclear-propulsion cislunar shuttle for transportation between Earth orbit and a lunar-orbiting space station, and a chemical-propulsion space tug that would do double-duty as a moon lander - would support the post-Apollo lunar program. This “integrated program” would lay the groundwork for the first manned Mars landing in the 1980s.
This vision, often identified with rocketry pioneer Wernher von Braun, director of NASA’s Marshall Space Flight Center in Huntsville, Alabama, is probably better ascribed to George Mueller, who became NASA’s Associate Administrator for Manned Space Flight in November 1963, NASA’s new Administrator, Thomas O. Paine, an STG member and chief author of NASA's report to the STG, and U.S. Vice-President and STG chair Spiro Agnew. More politically savvy and technically conservative STG participants – for example, U.S. Air Force Secretary Robert Seamans, a former NASA Deputy Administrator - did their best to rein in the breathless enthusiasm of Washington neophytes Paine and Agnew. Seamans, Office of Management and Budget chief Robert Mayo, and others understood that neither President Nixon nor the Congress would support a new Apollo-scale space program, let alone one several times larger and more costly.
NASA’s report proposed four possible "program rates" based on available funding. The “maximum rate” (that is, fully funded) program would begin in 1975, immediately following the last Apollo moon missions and the Skylab Program, with the launch on a two-stage Saturn V of the Earth-orbiting station and the maiden flight of the winged Earth-to-orbit shuttle. The following year, NASA would use Saturn Vs to launch a space station to lunar orbit and would debut the space tug/lunar lander. The year 1978 would see introduction of the nuclear cislunar shuttle and a lunar surface base established using space tug landers. By 1980, a 50-man Space Base would orbit the Earth. The next year, NASA would launch the first in a series of three-year Mars expeditions. The Space Base, meanwhile, would expand by 1985 - just one decade after program start - to support a crew of 100.
NASA’s Program I was only a little less ambitious than the maximum rate. It would start a year later, in 1976, with the Earth-orbiting space station and Earth-to-orbit space shuttle. The lunar-orbiting station and space tug/lunar lander would be postponed two years to 1978. The nuclear cislunar shuttle would, however, also debut in 1978, the same year as in the maximum rate program. The year 1980 would see both the lunar surface base and the 50-man Space Base brought into service. The first Mars expedition would be bumped to 1983, but the 100-man Space Base would be in place by 1985, as in the maximum rate plan.
Program II, the pacing option Agnew favored, would get off to a delayed start, with the Earth-orbiting space station and Earth-to-orbit shuttle both coming on line in 1977. The lunar-orbiting station, space tug/lunar lander, and nuclear cislunar shuttle would begin operations simultaneously in 1981, with the lunar surface base following two years later. The following year (1984), 50 men would orbit Earth in a Space Base. Men would walk on Mars for the first time in 1986, and the Space Base population would reach 100 in 1989.
NASA’s Program III, tacked on almost as an afterthought, was hardly an attempt at conservatism: it was identical to Program II, except that it set no date for the first Mars expedition. The Nixon Administration paid lip service to elements of Program III, declaring that the decisions it was making about NASA's future would make possible a manned Mars voyage before the end of the 20th century. At the same time, however, Nixon continued to cut NASA's budget until it touched bottom at about $3 billion in Fiscal Year 1971. This spelled the end for the technologies and hardware NASA had identified as necessary for humans on Mars, including nuclear propulsion, the Saturn V, and the large Earth-orbiting station. Mueller saw the handwriting on the wall and left his NASA post in December 1969; Paine resigned close to the first anniversary of NASA's report to the STG.
For his part, Mueller's departure did not signify that he had abandoned support for an integrated NASA program including space stations, a moonbase, and Mars expeditions. Almost as a parting shot, he published "An Integrated Space Program for the Next Generation," the cover article in the January 1970 issue of Astronautics & Aeronautics.
Significantly, the cover illustration for Mueller's article depicted a two-stage fully-reusable Space Shuttle (top image above). The Shuttle had already begun its rise from a mere utilitarian logistics spacecraft in Mueller and Paine's plan to the multipurpose centerpiece of Nixon's post-Apollo space program. On January 5 of the election year 1972, Nixon and new NASA Administrator James Fletcher unveiled a partially reusable Space Shuttle design in California, a state crucial for Nixon's reelection bid. They announced that the Shuttle would be assembled there, creating tens of thousands of aerospace jobs.
For more than a year before President John F. Kennedy's May 25, 1961 call for a man on the moon, Apollo had been seen primarily as an Earth-orbital spacecraft capable of both independent manned missions and crew ferry flights to Earth-orbiting space stations. A decade after Kennedy's call, NASA was preparing for Skylab A, its first Earth-orbiting space station, which would receive at least three three-man crews on board Apollo Command and Service Module (CSM) spacecraft (images above). The agency also studied independent CSM missions in Earth orbit and CSM missions to Earth-orbiting stations other than Skylab A.
The CSM, which measured a little more than 11 meters long, comprised the conical Command Module (CM) and the drum-shaped Service Module (SM). The CM's nose carried a probe docking unit, and at the aft end of the SM was mounted the Service Propulsion System main engine. The CM also included the pressurized crew compartment, flight controls, a bowl-shaped heat shield for Earth atmosphere reentry, and parachutes, while the SM included hydrogen-oxygen fuel cells for making electricity and water, propellant tanks, four attitude-control thruster quads, and room for a Scientific Instrument Module (SIM) Bay.
On August 27, 1971, Philip Culbertson, director of the Advanced Manned Missions Program at NASA Headquarters in Washington, DC, dispatched a letter to Rene Berglund, Manager of the Space Station Project Office at NASA's Manned Spacecraft Center (MSC) in Houston, Texas, in which he outlined five Earth-orbital CSM missions that were "still under active consideration" at NASA Headquarters. Culbertson explained that his letter was meant to "emphasize the importance" of statements he had made in a telephone conversation with Berglund on August 19.
Culbertson referred to an unspecified new contract MSC had awarded to North American, prime contractor for the CSM. He told Berglund that, in "the early stages of your contract. . .you should concentrate on defining the CSM modifications required to support each of the [five] missions and possibly more important defining the effort at North American which would hold open as many as possible of the [five] options until the end of the [Fiscal Year] 1973 budget cycle." U.S. Federal Fiscal Year 1973 would end on October 1, 1973.
The first and simplest of the five missions was an "independent CSM mission for earth observations." The mission would probably use a CSM with a SIM Bay fitted out with remote-sensing instruments and cameras. At the end of the mission, an astronaut would spacewalk to the SIM Bay to retrieve film for return to Earth in the CM.
The second mission on Culbertson's list was an Apollo space station flight unlike any envisioned in the year before Kennedy diverted Apollo to the moon. It would have seen a CSM dock in Earth orbit with a Soviet Salyut space station.
Salyut 1, the world's first space station, had reached Earth orbit on April 19, 1971. The 15.8-meter-long station remained aloft as Culbertson wrote his letter, but had not been manned since the Soyuz 11 crew of Georgi Dobrovolski, Viktor Patsayev, and Vladislav Volkov had undocked on June 29, 1971, after nearly 24 days in space (a new world record). The three cosmonauts had suffocated during reentry when their capsule lost pressure, so the Soviet Union had halted manned missions while the Soyuz spacecraft underwent a significant redesign.
The third Earth-orbital CSM mission on Culbertson's list combined the first two missions into a single mission. The CSM crew would turn SIM Bay instruments toward Earth before or after a visit to a Salyut.
Culbertson's fourth CSM mission would see the Skylab A backup CSM (CSM-119) with a crew of three dock first with a Salyut for a brief time, then with Skylab A. CSM-119's crew would remain on board 26-meter-long Skylab A for an unspecified period. NASA planned that, during the three missions to Skylab A in the basic Skylab Program, CSM-119 would stand by as a rescue vehicle capable of carrying five astronauts (Commander, Pilot, and the three rescued Skylab A crewmen). It would thus need to be refitted for the Salyut-Skylab A mission. Culbertson added that the Salyut-Skylab A mission would begin 18 months after Skylab A reached orbit.
The fifth and final Earth-orbital CSM mission was really two (or, possibly, three) CSM missions. A pair of "90 day" CSMs would dock with the Skylab B station while a rescue vehicle modified to carry five astronauts stood by. Beginning in 1969 (that is, at the same time it started Skylab A funding), NASA had funded assembly of Skylab B as a backup in case Skylab A failed. Culbertson gave no date for the Skylab B launch, which would have required one of the two Apollo Saturn V rockets made surplus by the September 1970 cancellation of the Apollo 15 and 19 missions (the Apollo 20 mission had been cancelled in January 1970 to make its Saturn V available to launch Skylab A).
Of the five missions Culbertson declared to be on the table in August 1971, not one flew. Skylab A, re-designated Skylab I (but more commonly called Skylab), reached orbit on May 14, 1973. It suffered damage during ascent, but NASA and its contractors pulled it back from the brink. In August 1973, with Skylab I functioning well in Earth-orbit, NASA began to mothball its backup. Several plans for putting Skylab B to use were floated in the 1973-1976 timeframe, but Space Shuttle development had funding priority, so NASA's second space station wound up in the National Air and Space Museum.
The three CSM missions to Skylab spanned May 25-June 22, 1973, July 28-September 25, 1973, and November 16, 1973-February 8, 1974, respectively. Leaks in attitude control thrusters on the second CSM to dock with Skylab caused NASA to ready CSM-119 for flight; the leaks stopped of their own accord, however, so the rescue CSM remained earthbound.
In early April 1972, shortly before finalizing its agreement with NASA to conduct a joint Apollo-Salyut mission, the Soviet Union declared the concept to be impractical and offered instead a docking with a Soyuz. At the superpower summit in Moscow on May 24, 1972, U.S. President Richard Nixon and Soviet Premier Alexei Kosygin signed the agreement creating the Apollo-Soyuz Test Project (ASTP).
Apollo CSM-111 was the ASTP prime spacecraft, while CSM-119 was refitted to serve as its backup. In the event, the backup was not needed. CSM-111, designated simply Apollo, docked with Soyuz 19 on July 17, 1975. The last CSM to fly undocked on July 19 and returned from Earth orbit on July 24, 1975.
Jumat, 03 Maret 2017
By late 1998, JPL had settled on an MSR mission design based on the Mars Orbit Rendezvous (MOR) mode. This was not surprising, since the Pasadena laboratory had staunchly advocated MOR since the early 1970s. At that time, JPL was responsible for building the Viking Orbiter, and MSR missions in the late 1970s/early 1980s were expected to be based on Viking hardware. The MOR mode would require use of an Orbiter; MOR's chief rival, Direct-Ascent, would not because it would launch samples directly from the surface of Mars back to Earth. No Orbiter meant no role for JPL in Viking-based Direct-Ascent MSR; JPL thus supported MOR. This institutional preference became thoroughly ingrained by the early 1980s.
In MOR MSR's basic form, samples would reach Mars orbit in an ascent vehicle. An orbiter would perform rendezvous and collect the samples, then would depart Mars orbit for Earth. Splitting the Mars ascent and Earth return functions between ascent and orbiter vehicles would enable a smaller, lighter Mars lander than in Direct-Ascent, and thus would trim overall mission mass. The reduced mass of MOR MSR would mean that the mission could leave Earth on a smaller, cheaper launch vehicle or could include more science payload - for example, a rover for sample collection beyond the immediate landing site.
One can argue, however, that MOR increases mission complexity and thus risk. JPL's 1998-1999 MOR MSR plan aimed to reduce risk by collecting samples from two different martian sites using landers launched from Earth during two Earth-Mars transfer opportunities (specifically, in 2003 and 2005). After completing its 90-day sample collection mission, each lander would launch to Mars orbit a Mars Ascent Vehicle (MAV) bearing an Orbiting Sample (OS) canister. To help keep its MSR mission under a strict cost cap, NASA invited the French space agency, Centre National d'Etudes Spatiales (CNES), to provide the MSR orbiter.
At the August 1999 AAS/AIAA Astrodynamics Specialist Conference in Girdwood, Alaska, a team of engineers from JPL and another from JPL contractor Charles Stark Draper Laboratory (CSDL) presented papers in which they examined how the CNES orbiter might perform rendezvous with the 2003 and 2005 OSs. They proposed a complex three-phase MOR strategy for each OS consisting of preliminary, intermediate, and terminal rendezvous phases.
In 2003, OS preliminary rendezvous would begin with MAV liftoff. The 2003 MSR lander would be rated to function on Mars for 90 days, so its MAV would need to launch from Mars within 90 days of touchdown. The 2003 OS would thus reach Mars orbit no later than April 2004. To save money and ensure adequate development time, the JPL MSR mission would employ a simplified solid-propellant MAV with a spin-stabilized first stage and a second stage with only a simple guidance system.
In their paper, the JPL engineers noted that even small OS orbit dispersions could place significant rendezvous propulsion demands on the CNES orbiter. An OS dispersion of only 1° in inclination, for example, would require that the orbiter alter its velocity by an additional 60 meters per second to match orbits, which would require an additional 48 kilograms of propellants.
For their MOR calculations, they assumed that a MAV capable of reliably placing the OS into a circular orbit 600 kilometers above Mars (plus or minus 100 kilometers) and inclined 45° to the planet's equator (plus or minus 1°) could be developed. They assume that the OS would take the form of a 14-to-16-centimeter sphere covered with solar cells which would power a radio beacon. The OS power system would include no batteries, so the beacon would operate only when the cells were in sunlight.
Between July 24 and August 26, 2006, the CNES orbiter would arrive in 250-by-1400-kilometer Mars orbit inclined 45° to Mars's equator. Once there, it would activate its Radio Direction Finder (RDF) to begin a four-week hunt for the 2003 OS. The RDF, which would collect OS data for relay to controllers on Earth, would have a range of 3000 kilometers. The JPL engineers suggested that other spacecraft in Mars orbit (Europe's Mars Express, the U.S. Mars Surveyor 2001 orbiter, or a specialized U.S. navigation & communications orbiter proposed for launch in 2003) might augment data from the CNES orbiter's RDF.
On September 24, 2006, controllers on Earth would begin the intermediate rendezvous phase by commanding the CNES orbiter to perform the Nodal Phasing Initiation (NPI) maneuver, the first in a series of maneuvers over 19 weeks designed to nearly match orbits with the 2003 OS. Radio signal roundtrip travel time would gradually increase from 23 to 43 minutes over the 19 weeks as Mars and Earth moved apart in their Sun-centered orbits.
At the start of this phase, both OS and orbiter would be in orbits inclined about 45° to Mars's equator; however, their orbits would have different ascending and descending nodes (that is, they would cross the equator at different places) and thus different orbital planes. In the planned 2003 OS orbit, the nodes would shift along the equator at the rate of 6.09° per day. This shifting, called regression of the nodes, would occur because of irregularities in the martian gravity field. The NPI would adjust the CNES orbiter's orbit so that its nodes would shift at a slightly different rate, enabling it to gradually match nodes with the 2003 OS.
Between October 8 and November 5, 2006, Mars would be behind the Sun as viewed from Earth and largely out of radio contact. No maneuvers would occur during this solar conjunction period, though nodal phasing would continue.
The Nodal Phasing Termination maneuver on January 7, 2007, would see the 2003 OS and CNES orbiter in nearly the same orbital plane. At the end of the intermediate rendezvous phase (February 4, 2007), the orbiter would be two kilometers below and 400 kilometers behind the OS. In its slightly lower (thus slightly faster) orbit, the orbiter would close with the OS at a rate of 200 kilometers per day (about 8.3 kilometers per hour).
In their paper, the CSDL engineers proposed a "double coelliptic" rendezvous strategy for the week-long terminal rendezvous phase. The CNES orbiter would fire its rocket motor about two days before planned OS capture to place itself in an orbit only 0.2 kilometers lower than that of the OS. This would slows the closing rate to about 20 kilometers per day (about 0.8 kilometers per hour).
The orbiter would acquire the OS with its twin Light Detection and Ranging (LIDAR) lasers as it closed to within five kilometers. At a distance of 0.4 kilometers, the orbiter would perform several maneuvers to intersect the OS's orbit 80 meters ahead of the OS. As it crossed the OS's path, it would fire its motor again to precisely match orbits.
The orbiter would then keep station with the OS for four hours. During this period, controllers on the ground would check the orbiter's systems. If everything checked out as normal, they would then give it the go-ahead to perform OS capture. If all went as planned, the CNES orbiter would automatically capture the 2003 OS on February 11, 2007.
The 2005 OS preliminary rendezvous would overlap the 2003 OS intermediate rendezvous. For purposes of their study, the JPL engineers assumed that the 2005 MAV would deliver its OS to Mars orbit on October 8, 2006, the last possible day before the start of solar conjunction. The 2005 OS would be targeted to an orbit matching as closely as possible that planned for the CNES orbiter at the time it captured the 2003 OS.
Intermediate rendezvous in 2005 would begin immediately after 2003 OS capture (that is, at the end of the 2003 OS Terminal Rendezvous phase) on February 11, 2007. Nodal phasing would end after 13 weeks, on May 13, 2007, and the 2005 OS Intermediate Rendezvous phase would end on June 10, 2007.
The 2005 OS terminal rendezvous would resemble its 2003 counterpart. The CNES orbiter would capture the 2005 OS on June 17, 2007, then would begin a series of maneuvers over four weeks to place itself into the proper orbital plane for departure for Earth on July 21, 2007.
The JPL engineers calculated that each 10-meter-per-second velocity change during intermediate rendezvous would require about 8 additional kilograms of orbiter propellant and subsystem mass at launch from Earth, and that the orbiter would need to make velocity changes totaling 478 meters per second during intermediate rendezvous if it were to have a 99% probability of successfully capturing both the 2003 and 2005 OSs. This would imply a rendezvous propellant mass of 382.4 kilograms. They noted that the MSR Project required only a 99% probability of retrieving one OS, and that this level of reliability could be achieved with an orbiter capable of velocity changes totaling 349 meters per second (which implied a propellant mass of 279.2 kilograms).
The CSDL engineers added that a 99% probability of successfully retrieving one OS meant a 60% probability of retrieving both. They calculated that terminal rendezvous using the propellant-saving double coelliptic rendezvous strategy would require velocity changes totaling only a little more than one meter per second up to the 80-meter stationkeeping point, and no more than 4.6 meters per second from the 80-meter point up to capture.
Paraterraforming would offer other important advantages over traditional terraforming besides speed. A worldhouse could be constructed using technologies known since the 1960s, Lewis estimated. Terraforming, on the other hand, would demand technological breakthroughs. Initial investors in the paraterraforming project could live to see at least a small part of the uninhabitable world's surface roofed over and made to resemble Earth. The long timescale inherent in most traditional terraforming proposals, on the other hand, would mean that initial investors could expect to receive no gratification in return for their investment. Finally, paraterraforming's modular approach would allow "staged pay-as-you-go funding," so could proceed in fits and starts. Terraforming would require sustained high funding levels.
In his paper, Taylor emphasized Mars paraterraforming. He pointed out that Mars's gravity is only one-third as strong as Earth gravity, so terraformers would need to pile on an atmosphere 75% as massive as Earth's to give it an earthlike surface pressure. With only 28% of Earth's surface area, Mars might not contain enough gas in its crust and polar ice caps to create such an atmosphere, and importing gas from elsewhere in the Solar System might prove to be impractical. Taylor estimated that providing an earthlike surface pressure inside a two-kilometer-tall Martian Worldhouse (MWH) would require less than one-tenth as much gas as a terraformed Mars atmosphere.
Taylor proposed that MWH construction begin in a seismically stable area with little subsurface ice. The MWH roof would comprise inner and outer layers held in place by cables. Atmospheric pressure within the MWH would push its roof upwards, so supports within it would serve primarily to hold it down.
Taylor envisioned three types of MWH support towers. Inhabited Mars Support Towers (IMASTs) would resemble the 3.25-kilometer-high "vertical super-city" designed by the British architect W. W. Frischmann in 1965. Each would measure 110 meters across its foundation and consist of six load-bearing masts clustered around a hollow core outfitted to house 500,000 settlers. Mars Support Towers (MASTs), uninhabited IMASTs, could be converted into full-fledged IMASTs as martian population grew. IMASTs and MASTs would be spaced equidistantly six kilometers apart. Compression-Tension Towers (CTTs), uninhabited 30-meter-diameter tubes with tension cables running through their cores, would be spaced equidistantly two kilometers apart.
The first IMAST would provide a manufacturing and construction facility for MWH components. Six MASTs and 30 CTTs would be erected around it, yielding a habitable MWH "cell" 30 kilometers wide. Taylor envisioned adding MWH cells until eventually about 84% of Mars was roofed over.
The unroofed 16% of the planet would include Valles Marineris, a tectonic rift system with crustal layers. Taylor wrote that the abundance of mineral deposits found in Earth's rift zones suggested that Valles Marineris might provide materials for manufacturing MWH structures. Other unroofed areas would include the poles, which would provide ices and gases and serve as "dumping zones," and volcanoes taller than seven kilometers. The calderas of such volcanoes rise above most of the thin martian atmosphere; Taylor maintained that this would make them ideal locations for solar power generators.
He wrote that settlers might choose to flood with water areas of low elevation within the MWH to create lakes and seas. The towers standing in them would, he noted, need to be specially braced to stand against currents and waves.
Taylor then threw cold water on his paraterraforming concept and on the concept of planetary settlement in general, arguing that
The ASO was envisioned as a spent-tank space station; that is, it would start out as a rocket stage filled with liquid propellants and would be converted into a pressurized habitat for astronauts after it expended its propellants by placing itself into low-Earth orbit. The spent-tank station concept may have originated with Wernher von Braun in the 1940s. In the late 1950s, several space engineers, including Krafft Ehricke of General Dynamics and Kurt Strauss and Caldwell Johnson of NASA's Manned Spacecraft Center, developed spent-tank station designs. Beginning in late 1964, von Braun urged that the concept be made part of NASA's proposed Apollo-based post-Apollo space program. By 1966, the Saturn S-IVB stage-based "wet workshop" (images at top) had become a key element of the Apollo Applications Program (AAP).
Nissim proposed that the ASO be built into the second stage of a 107-foot-tall, 17-foot-diameter chemical-propellant rocket. He envisioned launching the ASO from near-equatorial Christmas Island, located in the Indian Ocean northwest of Australia. The rocket's first stage, with three engines generating 150,000 pounds of thrust each, would expend 154,266 pounds of liquid hydrogen fuel and liquid oxygen oxidizer during 145 seconds of operation, boosting the second stage to a velocity of 9800 miles per hour.
The second stage would separate from the spent first stage, coast for eight seconds, then ignite its single 150,000-pound-thrust engine to boost itself to a velocity of 16,300 miles per hour. Following engine shutdown, the second stage would coast to an apogee (highest point above the Earth) of 300 nautical miles. At apogee, the engine would ignite a second time to boost the second stage to an orbital velocity of 17,000 miles per hour and circularize its orbit, which would be inclined 40° relative to Earth's equator. The second stage would burn a total of 86,788 pounds of liquid hydrogen and liquid oxygen to achieve its operational orbit.
A = second-stage rocket engine; B = tanks holding gaseous oxygen and nitrogen for liquid hydrogen tank purging and pressurization; C = streamlined launch shroud segment with solar cells on concave inner surface (one of four); D = Schmidt telescope; E = star tracker for accurate telescope pointing; F = Cassegrain telescope; G = loop antenna for radio astronomy; H = emergency reentry vehicle; I = airlock hatch for spacewalks; J = emergency reentry vehicle launch escape/deorbit rocket motor (in airlock); K = relaxation area restraint positions (one of two); L = hatch from central column to interior of liquid hydrogen tank; M = central column; N = common bulkhead separating liquid hydrogen and liquid oxygen tanks; O = food lockers; P = life support equipment; Q = sleep area; R = radiation-shielded compartment; S = space suit storage.During launch and ascent to orbit, the initial four-man crew would ride in a conical emergency reentry vehicle with a dome-shaped nose, three fins, and a single solid-propellant motor. The emergency reentry vehicle would be mounted atop a six-foot-diameter cylindrical central column embedded in and protruding from the top of the second-stage hydrogen tank.
In the event of launch vehicle trouble during launch and ascent, the solid-propellant motor would ignite, blasting the emergency reentry vehicle to safety. The spent motor would then separate and the vehicle would descend to Earth nose first. During ascent, the astronauts would face forward in the direction of the vehicle's nose; during descent, their couches would pivot so that they would face in the direction of its tail. Shortly before landing, the emergency reentry vehicle would deploy a parachute to slow its descent.
Assuming, however, that they arrived safely in orbit, the astronauts would immediately begin to prepare the second stage for occupancy. First, they would turn it to maximize the amount of sunlight striking it and open valves in the second-stage engine. Solar heating would speed escape of any residual hydrogen through the engine nozzle into space.
Next, a space-suited astronaut would open a hatch in the emergency reentry vehicle leading into the airlock at the top of the central column. After sealing the hatch behind him, he would open a hatch into the radiation shelter, a section of the central column embedded within the hydrogen tank. There he would open a valve that would release into the hydrogen tank nitrogen gas stored in spherical tanks at the bottom of the second stage. The nitrogen would escape through the engine nozzle, purging the tank of any remaining hydrogen. The engine valves would then be closed.
A = second-stage rocket engine; B = tanks holding gaseous oxygen and nitrogen for liquid hydrogen tank purging and pressurization (five clusters); C = lower liquid oxygen tank bulkhead; D = liquid oxygen tank; E = common bulkhead separating liquid oxygen and liquid hydrogen tanks; F = bottom of central column; G = central column; H = sleep area; I = space suit storage; J = life support equipment access panel; K = lavatory; L = crew personal lockers; M = ventilation duct.The astronaut would next open a hatch leading from the central column into the hydrogen tank and move to the tank's bottom end. There he would permanently seal the hydrogen outlet port leading to the engine by welding a cover over it or by injecting a plastic sealant. He would then return to the central column, seal the hatch behind him, and release nitrogen into the hydrogen tank to check for leaks. While his shipmates monitored the tank's internal pressure, he would return to the emergency reentry vehicle.
Assuming that pressure in the tank remained steady, a space-suited astronaut would enter the central column to release oxygen into the hydrogen tank. According to Nissim, the pressure in the tank would equal the atmospheric pressure on Earth at 10,000 feet of altitude. The atmosphere in the tank would, however, contain as much oxygen as occurs at Earth's sea level. Located in the same area as the nitrogen tanks, the spherical oxygen tanks would contain enough gas to supply the ASO crew for 45 days.
The three astronauts waiting in the emergency reentry vehicle would then enter the hydrogen tank and doff their space suits. They would cut away metal covers welded over pre-installed equipment and openings (for example, air ducts), then would remove equipment and furnishings stowed in the central column and install them in the tank.
The crew would also point the emergency reentry vehicle's nose at the Sun and open four petal-like streamlined launch shroud segments located between the top of the second stage and the bottom of the emergency reentry vehicle. Besides revealing a "storage area" containing folded astronomical instruments, this would expose to the Sun electricity-generating solar cells covering the concave inner surfaces of the shroud segments. Attitude control thrusters and gyroscopes would keep the station properly oriented as it revolved around the Earth. (Nissim, by the way, proposed fueling the attitude control thrusters with crew urine.)
Pointing the emergency rentry vehicle at the Sun would also help to regulate temperature on board the ASO. The open shroud segments, telescopes, and emergency reentry vehicle would partially shade the spent stage part of the station. Alternating blue and white stripes of equal area would cover its hull. The blue stripes would absorb sunlight while the white stripes would reflect it. Most heating in the converted hydrogen tank would come from on-board equipment and the astronauts' bodies. Nissim estimated that the interior of the spent stage would maintain a temperature of 72° Fahrenheit.
The emergency reentry vehicle would be powered down, so would lack a significant internal heat source. It would, however, be in direct sunlight whenever the ASO was over the Earth's day side, so would be colored white with thin blue stripes so that it would reflect most of the sunlight striking it.
With ASO electricity, life support, and thermal control up and running, an astronaut would don a space suit, enter the central column airlock, pump the air it contained into the converted hydrogen tank, and open a hatch leading to the station's exterior. Linked to the airlock by a thin cable, he would deploy astronomical instruments from the storage area between the top of the stage and the bottom of the emergency reentry vehicle. By operating above Earth's obscuring atmosphere, Nissim explained, the ASO's instruments would for the first time in history permit astronomical observations of the entire electromagnetic spectrum from gamma rays to very long radio waves.
After deploying and checking the instruments, the spacewalker would return to and repressurize the airlock, then rejoin his colleagues in the tank. After doffing his space suit, he would settle into a routine that would see two crew members on duty, one asleep, and one off duty at all times.
In their explanatory text for the mockup, the Ideal Home Show organizers wrote that the initial crew would return to Earth in the emergency reentry vehicle, which they called the "reentry Vehicle (nosecone)." This would mean that only the initial crew could reside in the station before it was abandoned. According to its designer, however, the ASO would operate "forever," with new four-man crews and supplies arriving by ferry spacecraft every 30 days.
Nissim did not specify how astronauts would transfer between ferry and ASO. His design lacked docking ports, so he might have meant for astronauts to spacewalk between the two vehicles. Ferry spacecraft would remain at the ASO only long enough to rotate crews and drop off supplies. The emergency reentry vehicle, on the other hand, would remain part of the ASO throughout its career, enabling crews to evacuate immediately in the event of catastrophic meteoroid puncture, fire, or massive life support failure.
In February 1993, Kent Joosten, an engineer in the Exploration Program Office (ExPO) at NASA’s Johnson Space Center in Houston, Texas, proposed a plan for lunar exploration which, he hoped, would take into account the emerging realities of space exploration in the 1990s. His International Lunar Resources Exploration Concept 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 lunar oxygen as oxidizer for burning liquid hydrogen fuel brought from Earth, shipping most cargo to the moon separate from the crew, and relying on cooperation with Russia.
According to Joosten, a lunar lander making a direct flight from Earth's surface to the lunar surface that would arrive on the moon with empty oxidizer tanks and then reload with liquid oxygen mined and refined from lunar regolith (that is, surface material) would have half the mass of a lander that performed an Apollo-style Lunar-Orbit Rendezvous mission (itself a mass-saver) and brought to the moon oxidizer for the return trip from Earth. This would in turn permit a smaller launch vehicle, slashing costs.
One-way automated cargo landers, each rectangular in shape and capable of delivering 11 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 Russian Energia rockets from Baikonur Cosmodrome in Kazakstan. Joosten noted that launch teams at Baikonur could service two Energia rockets at the same time and that three Energia launch pads existed. An Energia would place a cargo lander into Earth orbit attached to a Russian "Block 14C40" upper stage that would then boost the lander toward the moon.
Shuttle-derived heavy-lift boosters would launch the piloted landers from Kennedy Space Center's twin Complex 39 Shuttle pads. The pads, monolithic Vehicle Assembly Building, and other KSC facilities would require modifications to support the new piloted lunar effort, but no wholly new facilities would need to be constructed, Joosten explained. The piloted lander, carrying an international crew and about two metric tons of cargo, would be placed into Earth orbit, then a new-design Trans-Lunar Injection Stage would put it on a direct trajectory to land near the pre-established automated oxygen production facilities.
Joosten's crew lander design outwardly resembled the "Eagle" transport in the 1970s TV series Space: 1999 (image at top of post). 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 horizontal, three-legged lander. At launch, the capsule would sit on top of the stack surmounted by a solid-propellant launch escape system. The three legs would fold against the lander's belly beneath a streamlined shroud during ascent from Earth. 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 Command Module 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 lander would land on and launch from the moon using four belly-mounted throttleable rocket motors. Soon after lunar touchdown, the lander would be reloaded with liquid oxygen from the automated oxygen plant. The entire lander would lift off the lunar surface for flight to Earth; no expendable descent stages would be left behind to clutter up the moon. Nearing Earth, the crew capsule would separate 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 away from populated areas. The crew capsule would use a steerable parasail-type parachute. Joosten recommended recovering the crew capsule on land - perhaps at Kennedy Space Center - to avoid the greater cost of Apollo-style water recovery.
Joosten envisioned a three-phase lunar program, but provided details for only Phases 1 and 2. In Phase 1, three cargo landers would deliver to the target landing site a nuclear power system, an automated liquid oxygen production facility, robotic diggers, loaders, power supply, and propellant transport "carts," and a pressurized exploration rover and science equipment. The first piloted lander carrying two astronauts would then arrive for a two-week stay, during which they would check out the automated systems and explore using the pressurized rover. Several Phase 1 piloted missions to the same site would be possible.
In Phase 2, three cargo flights would deliver a second pressurized rover, an airlock module, consumables on a cart, and science equipment. A fourth cargo flight would deliver a four-person crew for a six-week stay. The crew would divide up two to a pressurized rover. The airlock module would include docking units so that the two rovers and the consumables cart could link to it, forming a small outpost. Phase 3 might see larger crew sizes and longer stay times; alternately, NASA might change direction and use technology developed for the lunar program (for example, the crew capsule and Shuttle-derived heavy-lifter) to send humans to Mars.