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Water and Oxygen Resources:
A Lunar Discovery Mission
Back to the Moon: WORLD Mission
A Discovery Mission to Demonstrate
the Usefulness of Lunar Resources
for Future Space Explorers
A Space Systems Design Proposal for
ASE396 - Space Systems Design
University of Texas at Austin
December 7, 1994
This document is a response to the Discovery Mission's Announcement of Opportunity issued by the National Aeronautics and Space Administration (NASA) Headquarters on July 28, 1994. Contained herein is a proposal for a Discovery Class mission tentatively called Water and Oxygen Resources: A Lunar Discovery Mission (WORLD-M).
Twenty-five years ago humans first visited Earth's only natural satellite, the Moon. During that and subsequent piloted missions only a small fraction of the Moon's surface was explored and only limited samples from those explored areas were returned to Earth. Much remains to be learned about and from the Moon. But, from what scientists have learned from Apollo and other unpiloted missions, it is known that the Moon harbors many resources that could be used by future space explorers. A primary example is lunar regolith. Tests on Earth using this type of soil retrieved from the Moon reveal that significant amounts of oxygen can be extracted from the material by several different processes. Such in situ - produced oxygen could be used for Earth-return fuel oxidizer or for life support. Being able to produce needed resources on the Moon would reduce the amount of mass that would have to be launched from Earth.
While mass reduction is a core concern for any space endeavor, it will be the deciding factor for feasibility of human exploration of Mars. In that light Earth's Moon can be thought of as a "close by" proving ground for the technologies that will be needed to go to Mars and perhaps beyond. Besides oxygen production, the Moon can also be the place to test concepts for habitat construction, radiation shielding and extraction of other indigenous materials.
The accessibility of the Moon, it's known resources, it's still unknown features, and its natural allure to humans who have observed it and wondered about it for millennia are the thoughts that lead this team to propose a lunar mission. The primary objective of WORLD-M is to deliver a small, lunar oxygen demonstration device to the surface of the Moon and operate it to prove that the technology is feasible. The idea is that if the concept can be proven in an inexpensive way on a Discovery mission it would pave the scientific and political way for a full scale production device.
The secondary objective of the mission is to do remote sensing of the lunar polar regions from a very low lunar orbit, expressly looking for water. The presence of frozen water at the lunar poles has been suggested by various scientific theorists and has been given more credence by the recent flight of the Clementine mission. The remote sensing approach was chosen after design considerations eliminated a primary objective option of sending penetrators into the polar regions to directly sense for water. Validation of the existence of water on the Moon would be of interest both to the scientific community and to human space exploration planners and researchers.
Tertiary science instruments have been added to both the orbiter and lander components of the spacecraft as mass margins allowed to increase the scientific return of the mission. The details of these experiments will be explained in the proposal. There is, however, another goal of this mission that is perhaps more nebulous but, it could be argued, may be more important than the science and technology knowledge that is gathered. That goal is to generate public enthusiasm for space research and exploration. It is believed that the search for water and the production of oxygen on the Moon, the two resources of life, for a relatively small monetary outlay, would capture the attention, imagination and admiration of the general public. If such a goal is realized, even to a small extent, the dividends of WORLD-M will be reaped by the space program for years to come.
The WORLD-M mission is to be launched aboard the NASA-furnished Delta II launch vehicle into circular, low Earth orbit. From there the PAM-D upper stage of the Delta II will establish an elliptic orbit with perigee at the former circular altitude. The remainder of the main propulsion for the mission will be performed by a Rocketdyne liquid propellant engine - the cost of which will be included in the $150 million allotted for a Discovery mission. This engine will provide the additional thrust necessary to put WORLD-M on a translunar trajectory, the plane change that will allow polar, lunar orbit, the lunar orbit insertion burn and the lander's descent to the Moon's surface.
The lander, with its oxygen production plant and other scientific instruments, will be landed on the Moon's near side near the equator. The work of the oxygen plant will be accomplished in a day, but the remainder of the instruments will continue to operate, whenever the landing site is in sunlight, for several months. Data collected will be sent directly to the Earth from the surface through the Deep Space Network (DSN). The water sensing equipment and other devices that remain on the spacecraft in lunar polar orbit will operate for approximately one year, as power and other resources allow. Data will be telemetered to the Earth through the DSN except for times during each orbit when the spacecraft will be occulted by the Moon. This data will be recorded on board for later transmission.
O2 Production Experiment
Oxygen production on the Moon is a major objective for the future of lunar exploration for the production of propellant and life support. The experiment consists of testing in situ the process of production of oxygen from lunar soil by hydrogen reduction of ilmenite, which has already been tested in labs. Here, water will be produced instead of directly extracting oxygen, since the latter may be obtained from H2O by electrolysis. Small samples (< 10 g) of regolith will be collected with a robotic arm, crushed, weighed and treated with hydrogen in a furnace. This furnace is heated by a 1.5-meter diameter solar dish up to 1000 K. The outlet of the furnace contains both H2O and H2, which are separated from each other in a cooler. The detection of water produced will be measured with a spectrometer. The experiment will take place twice during the lunar day. The whole device, including the robotic arm, has a mass of about 30 kg and requires 500 W of electrical power for various components such as the cooler, pump, and sensor.
The Search for Water
The most commonly accepted theory of how water arrives on the Earth is by cometary impact. Under current theories, during the formation of the Earth it would have been too hot for Earth to have retained any of its own water. Therefore, the water would have had to come to the Earth after it had cooled. It is believed that comets are the means by which water was delivered to the Earth. If the Earth was being impacted by comets then the Moon must have been also. Therefore, it is possible that the Moon could have trapped water if there was a mechanism to trap the water. Steven Brandt [Brandt] discussed a possible capture mechanism during cometary impact. The conditions created in the first 100 meters of the crater floor are conductive to the formation of hydrated rock. The water may be concentrated in rocks as hydrated silicates which can possibly be remotely sensed.
This mission will use a Gamma-Ray Spectrometer to target cometary craters to search for hydrate rock. The craters in the polar regions are more promising since the Sun will have had less of a chance to drive off the water. The spectrometer will also be able to sense any ice fields if they exist.
Several instruments will gather scientific data from lunar orbit and from the surface. The primary instrument on the orbiter is a gamma-ray spectrometer. This is a passive instrument that returns data on the composition of the lunar surface, including possible water content. The orbiter will also carry a magnetometer to chart the Moon's magnetic field, and a high-resolution ultraviolet-visible (UV/Vis) camera. The lander will carry a "cosmic dust" sensor to detect the velocity and charge of incident particles. The lander will also have a low-resolution UV/Vis camera. This camera will provide pictures of the landing site and the probe itself.
The spacecraft is composed of two distinct parts: the lander, with most of the propulsion subsystem and the orbiter. The total mass of the structure at launch is estimated at 100.5 kg for the lander and 49 kg for the orbiter (including solar arrays structure, antennas , payload boom and appendages and landing legs).
The central part of the structure consists of an aluminum cylinder of 1.5 mm thickness 1.5 and 36 cm radius which enables the spacecraft to sustain the acceleration during the launch. It is also built so that the first lateral and longitudinal frequencies of the spacecraft are sufficiently high compared to the ones of the launcher. The base of the cylinder is terminated by a conical section that adapts to the spelda of the launcher.
Next, two platforms for both the lander and the orbiter provide a support for the equipment and the payload. These platforms are dimensioned to avoid dynamic coupling with the excitations due to the launcher. They are made of honeycomb with a thickness of 1 cm, and the face sheet is made of composite material (carbon fiber: GY 70) with a thickness of 0.28 mm. The upper platform has a circular that rotates to position the furnace in the axis of the mirror. This is possible due to a motor that is placed inside the top of the cylinder included in the lander.
The side walls of the lander are made of the same honeycomb plates. The purpose is to provide a good protection for the oxygen plant elements versus lunar dust and micro-meteoroids (those with a momentum less than 1.2 10-3 can be stopped). However, the side walls for the orbiter are only aluminum sheets of 0.05 mm.
The last items of interest for the primary structure are the fuel and oxidizer tanks. Sized to accommodate propulsion system needs, three spherical tanks (r=21.86 cm) are required for the N2O4 and three others (r=21.65 cm) for the MMH. By estimating the internal pressure, it was possible to estimate (with a 40% safety factor) the thickness of the tanks, which is around 0.5 mm. The tanks are fixed to both the conical section and the main cylinder. The lower part of the cylinder will be cut and ejected after the braking phase, along with the tanks and the main engine that are fixed on it. The tanks are made of titanium.
Finally some evaluations, using buckling and dynamic criteria, were performed to obtain the sizing of the secondary structures. Data were derived for the landing legs (thanks to Surveyor IV data on the shock at impact), the antennas (honeycomb sandwich with carbon fiber faces), the sample collector arm (made of beryllium) and the gamma-ray spectrometer boom (CFRP composite, 2.5 m long, a radius of 3.48 cm and a thickness of 5mm).
The thermal control systems of the orbiter and lander are designed to dissipate excess heat and to maintain the instruments and systems within their required temperature limits. Each vehicle has sources of heat such as the electronics and other systems, and also will be heated by solar radiation. A passive thermal control system is used to transfer the heat from these sources and radiate it into space using thermal radiators. This system is also used to maintain the required temperature limits of the on-board instruments. For the lander, the thermal radiator was found to require an area of 2 m2, while for the orbiter it may be necessary to use an active thermal control system because of the temperature limits of the gamma ray spectrometer.
WORLD-M will be launched by a Delta II 7925. The third stage will provide a part of the thrust for insertion into transfer orbit. The main propulsion system is a XLR-132 from Rocketdyne, with a thrust of about 17,000 N. The engine is part of the lander itself. It will be used for final insertion into LTO, midcourse corrections and insertion into LLO for the lander/orbiter complex. Then the two craft will separate from each other, and the engine will deorbit the lander and assume part of the breaking. A few hundred meters above the surface, the main engines and the tanks are jettisoned, breaking switches to small vernier engines, which drive the lander to a stop 10 m above the surface. The lander then falls and reaches the soil with a velocity of 5.7 m/s, which can be supported by the three legs. Attitude control of both the lander and orbiter is accomplished by small hydrazine thrusters.
Guidance, Navigation and Control
In general the guidance, navigation, and control (GNC) subsystem is responsible for sensing and determining attitude and position, and controlling the spacecraft such that all mission objectives are achieved. It is assumed that the launching vehicle and the upperstage provide their own separate GNC system. The orbiter will have an inertial measurement unit (IMU) to measure velocity and provide an accurate attitude reference, two fine sun sensors, and a horizon sensor. Three-axis stabilization control technique is employed through the use of reaction wheels with small thrusters for momentum dumping when needed. The lander will take full advantage of its propulsion system for control. An IMU will be employed for position and attitude reference along with a laser ranger. The total orbiter GNC mass is 56 kg while the lander is only 5 kg (excluding propulsion).
The electrical power subsystems carry out the power generation, distribution, and storage for the other spacecraft systems. The systems are chosen and sized according to the power requirements at the end of the mission. The orbiter's power system must provide 600W at the end of 1 year, and the lander needs a maximum of 600W during its mission. Both probes will use proven, "off-the-shelf" hardware.
The WORLD-M probes will both use Sun-tracking silicon (Si) solar arrays and nickel-cadmium (NiCd) batteries. In order to generate 600W at the end of the mission, the orbiter will have a 12.8m2 solar array, while the lander's will be 6.0m2. NiCd batteries are the best energy storage system for the relatively short mission duration. The orbiter will have three batteries and the lander will have two batteries. The power distribution system is based on the energy-efficient peak-power tracker, which extracts from the solar arrays the amount of power required by the loads. The common bus voltage is 28Vdc.
For the communications system, it was decided that both the orbiter and the lander will communicate directly with the Earth. In this way, a data storage system is not required on the lander. Using a typical signal frequency in the Ku band, and a beamwidth that allows coverage of the entire Earth, the required diameter of the communications antenna was found to be 0.88 m. This size is dependent upon the signal frequency, but precise frequencies will not be known until they are assigned by the Federal Communications Commission. Therefore, this size was used for both the orbiter and the lander. This antenna size allowed determination of the mass of the communications system to be 15 kg for both the orbiter and for the lander.
The preliminary design of the computer was estimated by using the sizing by similarity method. The method provides conservation computer characteristics. Using this method allowed determination of the initial throughput requirements. The throughput requirements enabled the selection of the computers. It was decided to break-up each task and use multiple computers on the orbiter and lander. One computer was assigned the housekeeping tasks and the other handled the image processing and science instruments tasks. Multiple computers will also increase the reliability of our computer subsystem. The data recorders were selected to maximize storage while minimizing cost. Tape recorders provide by far the most storage for the smallest cost. The draw-back is that they only last up to a month and a half. This does not have any consequences on the lander because its lifetime is short. However, the orbiterŐs lifetime is approximately a year. Therefore, a majority of the orbiter experiments and images will be performed during the first month and a half of the mission. The solid state record will permit images and experiments to be performed after the first month and a half but at a much lower capacity.
Off-the-shelf components were incorporated in the selection of the computers and data recorders. The off-the-shelf components increase mission reliability and cost. Compression will be an important part of reducing and fine tuning the memory storage requirements. It was assumed that images taken by the UV/Vis Camera will be compressed in half. Encryption was not included in the initial sizing estimates but could be added in the future is desired. The total computer system mass for the orbiter is estimated at 50 kg and the mass of the lander at 40 kg. These estimates include a 50% mass margin because the uncertainty associated in preliminary sizing.
The design team feels that this proposal is certainly feasible from a technical standpoint and, indeed, has a high probability of success. Whenever possible, existing technology and components have been specified to increase reliability and cost effectiveness. Important science and technology demonstration studies will be conducted by WORLD-M for a cost that falls within Discovery guidelines. Also, there is no significant restriction on what year and what month this mission can be launched - the worst case lunar planar alignment has been assumed. And, finally, the purpose of this mission is believed to be of significant interest to the general public. For these reasons the team enthusiastically presents this proposal for consideration and eagerly anticipates working with NASA to see WORLD-M become a reality.
CSR/TSGC Team Web