Power requirements and sources have been one of the primary limitations on space travel and design since the beginning of the space age. Initially, batteries were employed, but were limited to a few days duration. Solar panels followed. These were limited power output sources, but had longer lifetimes. The photoelectric array was used in conjunction with batteries for quite a while. For missions beyond the asteroid belt, even these systems were untenable. With outer-planetary missions becoming prevalent in modern space exploration, the need for different power sources that were small, sturdy, and capable of producing a considerable amount of power arose. Radioisotope thermoelectric generators (RTGs) were developed to meet the above requirements. Radioisotope decay produces energy that, through the thermoelectric effect, converts to electricity. For the RTG, power relies primarily on component degradation and radioisotope half-life. RTGs were used in the Viking Mars Lander, Apollo Lunar Surface experiments, and other missions that are quite similar to the Odysseus project we are undertaking. At this point nuclear reactor-based power is a thought, but one that doesnt have a lot of history in space missions to back it up. The U.S. has only flown one reactor test mission (SNAP-10A) in the year 1972. Ultimately, power is a focal point in spacecraft design. It influences and is influenced by virtually every other subsystem in the project.

The purpose of the Power system is to generate and store power for use by the other spacecraft subsystems. Tradeoffs in power functions are therefore specifically important to the success of the overall design. The power system is designed to control, condition, and process the raw power from the power source to meet the needs of the other subsystems and to provide stable, uninterrupted power for the design life of the system. Along with this purpose, there must be redundancies built into the power system in order to prevent catastrophic malfunctions that could cause main bus voltage to drop considerably. For long-duration missions such as the Odysseus Project, large amounts of power will be required, therefore the power system will be very important and must be thoroughly engineered to handle any requirement for power in its mission life.

There are seven basic drivers for power system design. The first driver is the target planet, in our case Mars. The solar distance and corresponding solar energy per unit area are main considerations. The lifetime power requirements and operating environments are of extreme importance. Thermal, radiation, eclipse-area and time constraints, and other environmental factors are all paramount in power level and duration selection and design. Attitude control is also a driver of power systems design. This system may require power inputs or at least certain configuration constraints for size and balance. Orbital parameters such as surface operations, time spent on the dark side of the planet, and orbit inclination are also key factors influencing power systems design. This corresponds closely with the spacecraft operating speeds and orientations. The other main driver for the system design is the mission requirements and experimental considerations for power usage.

In typical design and engineering to date, there are seven common practices that power systems specialists and teams typically follow. Direct current switching is one of these common practices. Relays and switches are connected so that they are in a positive line to a given element, with a direct connection on the negative side being maintained. The purpose of this design practice is to allow power to be removed in case of a high-current flow failure or a short circuit in the element. Arc suppression is another common design parameter. Arc suppression devices should be placed very close to the arc source in order to ensure maximum arc suppression and therefore, maximum effectiveness. Modularity is also a key. Modular design provides for the easy testing, repairing, and replacing of power systems components. Modular pieces make installation and removal easier and precludes the expense of replacing the entire system if only a component or subsystem is damaged. Grounding of the system is also a common practice. As expected, grounding of electrical systems is imperative. In design, a ground cable is usually chosen over a ground structure. This is done to alleviate the difficulty of maintaining high continuity between structural elements. Using the common configuration, there isnt as much variation in ground voltages and the current is less likely to disturb the sensitive electrical components. Continuity must also be maintained so that static electric potential or other voltage differences do not build-up. Therefore, continuity in design is also commonly practiced. Shield continuity is also maintained across all connections when power systems are designed and implemented. The shielding serves two purposes: shield noise-sensitive circuits and shield noise-generating circuits within the system. A common design practice is to use a single point shield ground vice shielding the current flow. The final practice is one that is involved in virtually every engineering project---complexity. Like most other designs, power systems should be as simple as possible to effectively accomplish the mission and power requirements. Simplicity costs less, is easier to test and repair, and saves time and valuable space. Power Systems Engineers goal is to make the power system as complex as necessary, but also as simple as possible.

Solar arrays are made up of many cells aligned on a substrate. Typically, the cells produce small currents and voltages. However, when aligned properly in series or parallel, the array is capable of providing useful power output values. The most common solar cell is rectangular in shape and measures 2" x 4". This allows for tight cell packing on the substrate, a design that reduces mass and size while maintaining a packing density of approximately ninety percent. This packing density can not be further improved upon too much; it has already been optimized to near its fullest extent.

Solar arrays can be found in many forms. The skin of the spacecraft itself may consist of the solar cells or the solar arrays may be deployable in nature. Deployable panels can be found as fold-up type or roll-up type. These lightweight structures have the advantage that they can be packed into a small space due to the usage of softer composite materials in the substrate. Solar array design is driven by rigidity requirements vice strength. Thin-section, built-up structures are possible to meet the design criterion. The only problem with low-strength structures like these is the possibility of damage occurring when handled by crew or engineers. A perfect example of the fold-up/ roll-up type array was the 12.5 kW model demonstrated on a shuttle mission in the past eight years. Because our mission and many like it require high power and voltages, the conductor and insulation masses between circuit elements can be quite large. This problem makes flexible fold-up arrays a "least massive" alternative to full-size deployable arrays.

Because solar arrays depend upon proximity to the sun, there are amplifiers used on solar arrays operated at greater distances from the sun. These amplifiers are called concentrators and are flat arrays that are used with silicon cells having a trough that increases the collection area of the array relative to the cell area. This method is good for distances greater than 1.5 AU from the sun and is estimated to work well up to the

3-4 AU range. This method also keeps the cells from excessive cooling. Concentrators are expensive, especially the gallium arsenide ones. This type prefers a higher temperature than the silicon type concentrator. This is a cost reducing quality that brings the cell to a greater operating temperature than is otherwise attainable. The main disadvantage of this type of concentrator is its cost and complexity. These also create stowage and deployment problems.

When designing solar cells, the main concerns are temperature difference and the current-voltage relationship. Most cells are capable of 2 mV/ C temperature dependence characteristics. As the temperature drops, the cell voltage increases, leading to a decrease in current. Surge potential exists when the cold panels experience to great a temperature increase after emerging from eclipse. Therefore, the cells should be operated for minimum mass and maximum efficiency. To meet these conditions, the cells operate at the MPP or maximum power point. The majority of spacecraft systems are designed to operate at this point. The series/ parallel arrangements are determined by this MPP, which in turn, is driven by the current-voltage relation of the cells. In some cases, the MPP can be driven off of the I-V curve and require a battery on backup to bring the array power system back to its operating point. An online battery helps stabilize and correct this problem that is characterized by significant voltage drops. Even systems that dont typically require batteries may require them as backups to alleviate this potential problem. The batteries can be stored as offline, then brought online to provide the temporary energy boost required to fix the voltage problem.

Another consideration with solar arrays is that current drops as we move away from the sun. The voltage stays the same or increases if the temperature drops. The reverse condition is true when moving toward the sun. The MPP moves slightly as we move toward or away from the sun. for our planetary orbiter this becomes an important consideration if we choose solar arrays as a power source. An important reminder is that we need to plan for Mars to sun distances when we consider our solar array design as we will get less "bang for the buck" from the array when we are further away from the sun. We would also want to design our array such that we can keep the angle of incidence between the sun and array as close to ninety degrees as possible to guarantee maximum power output.

There are many other factors to consider when choosing solar arrays as power sources. Radiation degrades the solar cells. For a long mission in an adverse environment such as Odysseus II, this becomes a real design conflict. The array must be sized for end-life power requirements to ensure that the array provides enough power in the later days of the mission. More information on radiation effects will be presented at a later date. With the introduction of gallium arsenide cells, there has been a movement toward using more radiation-resistant solar cells. Currently, this material is only used when radiation resistance is a prime concern due to the ten times cost over silicon cells. For our project, with its complexity and uncertainty, cost should be minor consideration, though cost is a much more significant factor in most space missions.

The Odysseus I and II missions to Mars are indeed follow-ons to the Mars Pathfinder mission and the deployment of the Sojourner probe. Since the solar arrays used for these missions proved to be successful in terms of power conversion efficiency and in withstanding Mars harsh climatic conditions, it is logical to use them for the Odysseus missions. After all, the array was already designed to account for the solar energy density both near mars and on the surface of the planet. The components have already been engineered and designed to withstand the Mars and low Mars orbit climates. Solar systems are also more cost effective than the nuclear power plant and RTGs.

Our arrays are nearly exact duplicates of the Sojourner array. They have a ninety-one percent packing density and are made up of gallium arsenide cells. The gallium arsenide cells used in this array have an eighteen percent solar energy to electric energy conversion efficiency. This is quite high for solar arrays and is therefore a good choice for our power systems. The solar array consists of photovoltaic cells that convert the solar energy into electricity. The cells are made up of a number of semiconductor layers. Within these layers, an electron current is created. A metallic grid or other electrical contact collects the electrons from the semiconductor material. The energy is then transferred to the external load, the spacecraft subsystem requiring electrical energy. To shield the cells from weather and other environmental elements, a glass-like cover is placed over the top of the cell. This also serves as an anti-reflective coating which keeps the light from reflecting the light back away from the cell. This guarantees that more solar energy is "absorbed" by the cell.

The "cold-side" of the array is made up of Molybdenum or a lightweight Aluminum material. The metal used to form the protective grid contact points conducts the energy very well, accounting for a very small (3-5%) resistance loss. We must have these grids in order to efficiently collect current. The grid is very thin so as not to shade the cells from the incoming solar energy. Yet, the grid is thick enough to conduct well, with minimum resistance loss. Typically, this thickness is on the order of a few microns thick.

To further boost the usable power from each cell, we can connect the cells to form modules, which are in turn connected to form the complete array. At all times (when not in black planet conditions), at least half of our array will be exposed to the solar energy. This is based upon the cylindrical configuration of the solar array mounting. When behind the planet, the batteries and fuel cells discharge and the stored power is used to maintain power requirements. Finally, for thermal shielding of the array itself, we will use Silica Aerogel. This is the glass-like substance discussed earlier. Besides being used as an anti-reflection mechanism, it protects the array against thermal damage too. This Silica Aerogel is roughly one fiftieth the density of water, therefore it is extremely lightweight too.

The surface array will be mounted on the Odysseus II spacecraft skin. The array carried on Odysseus I will be somewhat different in that it must have more structural weight since it will not be supported by the aircraft structure. Still, the structural weight will not be much because we will mount the arrays on the bio-dome, which is a structural element itself. The surface area of the dome will not be completely covered by the solar arrays, but the array will be large enough so that the capture area is sufficient to harness 75 kW of power. This should account for 1545.45 kg of solar array mass.

The Odysseus mission solar arrays match the Pathfinder arrays nearly perfectly. For the Odysseus missions, the following technical information is quite significant:

Cell Type: Gallium Arsenide on Germanium

Cell Size: 2 x 4 cm each, 5.5 mil thick.

Coverglass: CMG material (Silica Aerogel), 3 mil thick

Substrate: Nomex honeycomb

Conversion Efficiency: 18 % +

Power: 45 W peak power at noon

Operating Voltage: 14-18 Volts

Survival Temperature: -140 +110 C

Specific Power: 132.35 W/kg Peak power

Power After Reclamation from Battery: 80% of Input

Ideal power densities for solar arrays are roughly 300 W/m². At peak power, we will attain a power density that is roughly 67% of the ideal. For our purposes and expected power output of the Odysseus solar array, this is sufficient. At present, ideal specific power levels for solar arrays fluctuate between 100 and 300 W/kg. At peak power, we only reach 132.35 W/kg. This is much lower than we would hope for, but it is still tolerable for our mission. In the future, lighter weight materials and more efficient cells can be used. This will drive our number up closer to the 300 W/kg level. However, at this time, for the power output we expect of our missions solar arrays, the 44% of maximum levels will be fine.

In order to get 30 kW of usable, onboard electrical power, we had to account for the fact that only 50% of the array would be exposed to the sun at any time because of the cylindrical configuration of the spacecraft. We further concluded that 20% of this power garnered would be lost when we reclaimed it from the battery systems. Therefore, we computed that we needed the array to generate 75 kW of energy at 100% sun exposure. This allowed for 37.5 kW of power to be harnessed into the batteries with a 50% array exposure. 20% of this 37.5 kW will be lost between the batteries and the systems requiring power, therefore 30 kW of power will be the final output.

Approximately 1500 m² of surface area was available for the array. We decided that 30 kW of usable, onboard electrical power would be a sufficiently large output for the solar array. We chose this value because it is quite large and will provide a great deal of power to the ships internal electronics and other systems. This value also is small enough that the power converters and batteries will not be overloaded or made to be excessively big and bulky.

The final onboard surface array will be 1000 m² in area and will weigh

1545.45 kg. This will fit into the size and area allotted us by the Configuration Systems team.

Usable Power after Batteries (kW)	Gross Power Required if 50% array exposure	Solar Array Area (m2)	Solar Array Mass (kg)
30 kW	75 kW	1000 m2	1545.45 kg

RTGs are independent of solar distances. However, there is limited use for RTGs at the present time. The radioisotope thermoelectric generator converts heat energy generated by radioisotope decay into direct current electricity by means of the thermoelectric effect. A central core a radioisotope material is surrounded by a thermocouple array in a parallel series. The hot junction contacts a canister of radioisotope while the cold junction attaches to the external generator wall to provide a heat sink (radiator) to space. RTG efficiency is limited by the conversion capability of the thermoelectric elements in the system. At present, the semiconductor thermoelectric elements allow for a 10 - 11 % conversion efficiency. The RTG is also limited by the internal thermal conductivity of the radioisotope. It is very difficult to design the thermal pathway such that a minimum temperature drop occurs from the isotope near the hot junction to the cold junction outer casing. This is a primary design goal, along with minimizing the heat leakage between the points that bypass the thermoelectrics. This is further limited by the material limits on the hot side (source) and the radiator size on the cold side (sink).

The RTG used on Galileo produced 298 W of power (+/- 10%) for a weight of

56 kg. For a larger system, the power to weight ratio will improve somewhat. Most commonly, Plutonium 238 is the radioisotope of choice. The Plutonium is a good choice of radioisotope. Over time, however, there will be a loss in energy output due to degradation of the thermoelectric elements which is caused by dopant migration at higher temperatures. Insulation breakdown due to temperature and radiation effects is another performance reducing condition in the RTG. For our mission length and power requirements, Strontium 90 appears to be the only viable option for the radioisotope. It has the longest half-life with the best power output capabilities of the choices available (28 yr half-life, 0.93 W/gm, 250$ W).

The RTG cant be turned off. If a temperature differential exists, the thermoelectrics produce energy. The radioisotope continues to decay. Therefore, the RTG is stored in a shorted condition so that the temperature differential is a minimum to keep energy production at a minimum. These conditions make handling difficult. The high temperature and radiation effects make installation, repair and replacement an arduous task. If a great deal of work needs to be done on the RTG, the crew needs to be larger in order to reduce exposure to the ionizing radiation. Due to the low earth orbit installation location, problems about radiation exposure at reentry or aborted takeoff can be neglected for the Odysseus project. However, the radiation must be considered for its detrimental effect on electronics and instruments. To limit radiation exposure, it is possible to mount the RTG on a boom that extends away from the spacecraft main body or to shield it with a cladding material.

As the name of the apparatus implies, the RTG consists of radioisotopes undergoing a decay process. The decay of the radioisotopes produce large amounts of thermal energy. Using thermocouples, the device converts the thermal energy to electric energy via the thermoelectric effect. The electrical energy created can then be transferred to the spacecraft subsystems requiring power. Radioisotope thermoelectric generators operate much like nuclear reactors. A central core of radioisotope material is surrounded by a thermocouple array in a parallel series. The hot junction contacts a canister of radioisotope while the cold junction attaches to the external generator wall to provide a heat sink (radiator) to space. The radioisotope fuel is compressed into seventy-two ceramic-like cells. Each heat source consists of eighteen separate modules, each encasing four of the fuel pellets. Typically, Plutonium 238 is the radioisotope fuel of choice for the decay process. It is one of the best choices of radioisotope although Strontium 90, Cesium 144, and several other radioisotopes also provide good material properties. Still, Plutonium 238 has the more ideal qualities in terms of half-life, power production per gram, and cost for most space missions (Strontium 90 also shows great promise for usage in long-duration missions like the manned Mars mission.) The radioisotope fuel is found in the form of plutonium dioxide solid fuel. As the radioisotope decays, heat is released. The heat is then converted to electricity by a thermoelectric converter. An electromotive force is produced from the electron diffusion across the cold and hot junctions. This occurs because the junctions are at different temperatures (approximately 700 K difference) and are made of different materials. The materials are joined together to form a circuit. These junctions, consisting of different metal wires, are thermocouples. At present, the semiconductor thermoelectric elements allow for a ten to eleven percent conversion efficiency from thermal to electrical energy. RTG efficiency is limited by the conversion capability of the thermoelectric elements in the system. The RTG is also limited by the internal thermal conductivity of the radioisotope. The waste heat can be radiated throughout the ship to provide heat for the crew or other systems or it can be radiated into space using a large heat radiator surface on the skin of the ship or attached to the power plant itself if it is mounted remote of the main spacecraft.

The radioisotope thermoelectric generator is a superior power source for spacecraft design. The RTG has no moving parts, making it a highly reliable power source. This helps to eliminate problems relating to component mechanical wear and deterioration. RTGs do not require bulky dynamic machinery to produce energy. RTGs are independent of solar distances. Unlike solar arrays, they do not require solar input to function properly. RTGs are primarily limited by the half-life of the radioisotope used and the thermal-to-electric energy conversion efficiency. RTGs are also small in size for the relative amount of energy produced. The RTG used on Galileo produced 298 W of power (+/- 10%) for a weight of fifty-six kilograms. For a larger system, the power to weight ratio will improve somewhat. Of this weight, only eleven kilograms of the weight was in the form of the plutonium dioxide fuel which was pressed into seventy two solid ceramic-like cylindrical pellets. The rest of the weight was comprised of the shielding and containment devices for the RTG. We believe that by redesigning the thermocouples somewhat, we can increase the thermal-to-electric power conversion efficiency to near twenty percent. This large increase in efficiency would provide electric power levels that are sufficient for the Odysseus I and II missions.

The RTG itself has multiple layers of protective material to shield the other spacecraft components from the high temperature levels and the radiation output. An aeroshell heat shield contains the carbon bonded carbon fiber sleeves and disks. These components have graphite impact shells which contain the fuel pellets and the rest of the internal RTG instruments. Since the fuel is stored in independent modular units, each having its own heat shield and impact shell, the chance for fuel release is minimized. This is because the design limits the impact distribution on each vessel. The multiple layers of protective materials , namely iridium capsules and high strength graphite blocks, help in the prevention of fuel leaks. These two materials are both corrosion resistant and strong. They also are compatible with the fuel pellet material. The graphite outer coverings protect against structural, thermal, and post-impact situations. The iridium cladding helps contain the fuel pellets in the event of a crash or other destructive end. Finally, the graphite and iridium are very heat resistant, making them ideal choices to protect in a high-heat environment.

When being implemented, the radioisotope thermoelectric generators are designed to withstand both impact and other forms of damage and destruction. The ceramic-form plutonium fuel is heat resistant, thus making it more difficult to be vaporized in case of fire or reentry environmental exposure. The fuel is also very insoluble. It has a low chemical reactivity and breaks in large pieces, not small parts that can be inhaled or ingested. Unlike in nuclear accidents, RTGs cannot explode because no fusion or fission processes are occurring. Hence, the acute radiation sickness associated with nuclear explosions wont be witnessed in an RTG accident..

As mentioned previously, the mechanical containment vessels protect the RTG from impact damage and excessive heat. To protect the crew and the electronic instruments that can be damaged, the RTG can be mounted on a boom that extends away from the main body of the spacecraft. The separation distance does a lot to limit exposure to alpha, gamma, and neutron radiation. Shielding or cladding can be used to further protect the spacecraft hardware and personnel. The shield can either contain the entire RTG internally, or a shadow shield can be used. A shadow shield is lighter in weight and smaller in size. It only blocks radiation exposure over a small area of importance. This keeps the radiation from penetrating vital areas of the spacecraft, while saving precious space and weight. If a shadow shield is used though, it must be large enough to properly shield the crew and the sensitive electronic instruments like the telecommunications and solar arrays.

Other safety factors have been built in to protect the ship and personnel from danger. To keep the worlds citizens safe, as well as protect the ship, RTGs are tested often. Detailed inspection and analysis are performed, along with safety briefings and a great deal of training involving the operation and care of radioisotope thermoelectric generators. All parties working with RTGs are well trained and equipped. Before launch is even approved, an Interagency Nuclear Safety Review Panel (INSRP) checks out the system. This group consists of academics, industry officials, government personnel, and environmental specialists to ensure all facets of system operations are safe. Excluding all of these inspections, the RTGs are initially designed to successfully pass a number of "crash tests." The RTG is engineered to withstand launch vehicle explosion/ fire, land or water impact following reentry, and a number of other post-impact conditions.

The RTG does, however, have some complications associated with its use. The RTG cant be turned off once it is activated. If a temperature differential exists, the thermoelectric elements will continue to produce energy. The radioisotope continues to decay. Therefore, the RTG is stored in a shorted condition so that the temperature differential is kept at a minimum to keep energy production at a minimum. It is very difficult to design the thermal pathway such that a minimum temperature drop occurs from the isotope near the hot junction to the cold junction outer casing. This is a primary design goal, along with minimizing the heat leakage between the points that bypass the thermoelectrics. This is further limited by the material limits on the hot side (source) and the radiator size on the cold side (sink).

Another problem with the RTG occurs only over very long periods of lifetime usage. There will be a noticeable loss in energy output due to degradation of the thermoelectric elements. This is caused by dopant migration at higher temperatures. The semiconductor materials are often laced with small amounts of dopant material like the Boron or Phosphorous that are added to Silicon-Germanium semiconductors. This is done because it produces an excess or deficiency of electrons. This makes the semiconductor a more efficient power converter than conventional metals. Insulation breakdown due to temperature and radiation effects is another performance reducing condition in the RTG. Also, as the radioisotope decays over an extended period of time, the rate of heat release decreases somewhat. Therefore, when designing a power system requiring RTGs, the designer should plan for the end of mission power outputs of the system rather than the optimal energy output at the beginning of radioisotope decay.

Radioisotope thermoelectric generators are also very dangerous to handle. The high temperature and radiation effects make installation, repair and replacement an arduous task. If a great deal of work needs to be done on the RTG, the crew needs to be larger in order to reduce exposure to the ionizing radiation. To limit radiation exposure, it is possible to mount the RTG on a boom that extends away from the spacecraft main body or to shield it with a cladding material. Since the RTGs Plutonium 238 or other primary radioisotopes generally emit short-range alpha particles, the power source is often mounted on a five meter boom that extends away from the spacecrafts main body and electronics suites. The alpha particles arent very dangerous and only travel a few inches in air. When working on the RTG, personnel can be protected from these particles if wearing a protective suit. The alpha radiation only becomes a problem when it is deposited inside of the human body. One of the reasons the plutonium dioxide is in a solid form is to prevent the easy inhalation of small particles should the radioisotope be destroyed or dropped. The neutron and gamma radiation from an RTG is also quite small, and poses no significant health hazard if the external dose is limited as much as possible.

The RTG final mass and volume, along with the anticipated power outputs are shown in the tables below:

The system mass is based on data previously supplied in the reports. For an RTG using Strontium 90, the total power/ system weight was 0.93 W/g. Taking the total power available and this value, allowed us to calculate the system mass. The volumes are approximations. We found the average density of the material components used in the RTG and scaled it as 3000 kg/ m³. The volume was found by dividing the system mass by the average system density. These estimates for volume are on the large side, but provide a good basis for later analysis. The power available from the RTGs follows:

This accounts for all of the energy generated by the RTGs of the ship. The Thermal Systems Specialist has been given our suggestions for the life support heating system using the RTG waste heat along with sketches we have prepared. He will most likely discuss the heat output of the RTGs for use in thermal systems in his work, so we excluded that data from this report.

Previously, we suggested that changing the dopant material in the semiconductor, or the semiconductor itself might increase the thermocouples energy conversion efficiency. We do not have the materials/metallurgical engineering knowledge to choose a better material at this time. We have found, however, from consulting with metallurgists and solid state physicists, that there are more efficient semiconductors on the market that can bring our conversion efficiency up by a few percentage points to approximately 10%. It seems that NASA did not use these materials in the past because the purifying methods either did not exist at the time or were too expensive to be feasible for that particular project. Unfortunately, none of the sources we consulted had enough information available to choose a better semiconductor material or dopant. The physicists and metallurgists did confirm that Improving the thermocouple would be the most likely way to improve the energy conversion efficiency from its current 7% level. Unfortunately, without the new semiconductor/ dopant available, our RTGs will still operate at 7% efficiency.

The Odysseus II mission attempts to utilize technologies from many different countries since the project is built on the premise that it is an international cooperative effort. For Odysseus though, we will not use nuclear reactors, even though we initially estimated that we would employ six Russian TOPAZ reactors. The Russian TOPAZ design is discussed below, but is not used in the final project.

The reactors use a fissioning core to provide heat for the direct conversion into electrical power. The Russian model operates using the principle of thermionics. The thermionic converter has two metallic conducting elements, an emitter and a collector. Electrons will flow between the two devices when the emitter is heated to 1000-2500 K, establishing a current flow between the two.

The basic unit of the reactor is the Electro-Generating Channel (EGK). It consists of a series of cylinders, one inside another. In the center are the Uranium Dioxide fuel pellets. The are enriched Uranium 235, enriched to 93 % and weighing 50 kg (total fuel pellet weight) respectively. The series of cylinders is surrounded by a Beryllium moderator, which is in turn encased by the emitter. The emitter is a cylinder of Molybdenum that is maintained at a temperature of 1773 K. Outside the emitter is the collector cylinder, which is made of Niobium. A Cesium vapor-filled gap is between the emitter and collector. The collector is maintained at 773 K by a flow of Sodium-Potassium ectectic. in order to create the temperature differential between the emitter/collector assembly. There are 79 of these EGK elements in the core.

The core is surrounded by a Beryllium neutron reflector with 12 control drums in the reflector. The criticality of the reaction is controlled by the amount of absorber facing the reactor. To perform a comparable operation to control rod insertion for the slowing and halting of the reaction, similar to the American control and SCRAM methods, the Russian TOPAZ system uses a positive temperature coefficient of reactivity to monitor the reactivity and criticality levels of the reactor system. Because of this crude system, the alignment of the internal control drums must be very precise, otherwise there could be a core meltdown.

The primary reason we will not be using the Russian TOPAZ system instead of a U.S. system that has been better engineered for safety is because the U.S. doesnt have a proven space-borne and tested nuclear power source. Because the TOPAZs reliability is somewhat questionable, the nuclear power source will not be used. Political and environmental concerns also preclude its use on this project.

This fact justifies paying the additional cost of the RTG over the nuclear reactors. The RTGs have not experienced the space-borne failures that TOPAZ has. The RTGs are also better engineered for safety and provide much less of a health and safety hazard.

If TOPAZ were used, the average electrical power output of the reactor over its lifetime is 10.2 kW. This is based off of the total thermal energy output of 85 kW at a conversion efficiency of 12%. The total reactor weight is 105 kg. This includes the 50 kg fuel pellet weight of enriched Uranium 235. The core measures 28cm x 26cm x 45cm. The total reactor size is then estimated at 0.033 m³. In actuality, the volume will be somewhat larger to account for the outer core components. A better approximation is a volume of 0.04 m³. The table below describe the pertinent sizing and capacity data.

Fuel cells are responsible for the direct conversion of chemical energy to electrical energy. Oxidizer and fuel are fed into the cell (like the battery in its internal configuration). The oxidation within the cell produces the requisite energy, but without the high temperatures typical of combustion reactions. Hydrogen and oxygen are the most common reactants in the current cells being used. The output is electrical energy and pure water. This can be used for drinking water by the astronauts. Energy is stored in the fuel cell to be used during eclipse or when solar arrays cant be used. When the arrays are used, they provide power for their normal functions and for re-energizing the fuel cells. The fuel cell mass varies with time since the mass of the reactant must be included in the equation. 2.6 kW of power can indicate a 1.1 kw hr/ lb.

The fuel cells used on Odysseus II will be 12kW capacity cells that rely on fueling from Oxygen and Hydrogen that can be stored in cryogenic tanks connected to the cells. Each cell is 14"x15"x45" in size and weighs nearly 260 lbs. The cells are electrically connected to a 28 V output. They use alkaline electrolyte technology to provide electrical energy and potable water. The cells employed on Odysseus II operate at a temperature of 200 ⁰F, a pressure of 4 atm, and are 70% efficient.

The Oxygen-Hydrogen reactants are stored in spherical tanks made from filament-wound Kevlar 49/ epoxy matrix. This material has a rupture stress well beyond expected tank stresses. A 10mm Titanium liner is included to reduce diffusion through the tank wall. The reactants were stored at an electrolyzer operating pressure of 2.2 MPa. As the reactants combine to form the water and electricity, the water is drained off into a potable water storage tank. The electricity is stored for later use. Heat in the fuel cells is dissipated through a system radiator that is standard equipment on the fuel cells. The heat sink operates at 250 K.

We wish the fuel cells to provide 84 kW of electrical energy. Based on this total power value, the following table describes the fuel cell quantity, mass, weights, and power outputs.

The power processor regulates the voltage supplied to the remainder of the spacecraft to within a specified tolerance of a fixed level. This serves to protect the subsystems from dangerous voltage fluctuations. Electrical noise from the main power source or the control functions is isolated from the main bus. The main bus should always be isolated from any power source faults, big or small, to better protect all of the systems. The processor functions to protect every system from a fault or failure in any other system.

Power conditioners process the electricity flowing from the batteries or other power systems into the subsystem requiring electrical power. The conditioners can convert AC power into DC power and vice versa. They also convert the power frequency, current, voltage, and maximum load to match the system to which the power is flowing. A configuration of inverters and regulators is responsible for converting AC and DC loads. The power conditioner can also limit current and voltage in order to maximize the power output. Finally, the power converter safeguards the utility network system and its operators/ maintenance personnel from possible harm during repairs by utilizing a system of breakers and parallel flow channels. The power converters can be stand alone systems or combined with the battery systems.

Since the total power available after all craft-borne systems is expected to be slightly less than 250 kW, we need a power conditioning system capable of handling at least a 500 kW load. We will once again employ a fully redundant system. Each 100 kW bank will be a composite of a series of smaller 25 kW systems. Each power source will be connected to each conditioning subsystem. Each conditioning subsystem will be connected in a parallel series to form the entire system of power conditioners. This will provide a failsafe should any one component fail. There will be a total of 10 power conditioners, each having a 50 kW capacity. The individual conditioners will weigh 85 kg and have an approximate volume of 0.02 m³. The total mass of the system is 850 kg, while the total volume it accounts for is 0.20 m³.

For systems that require processing of power loads above 25 kW, the power can be conditioned via controlled combinations of the power conditioners. In other words, the conditioners themselves can be connected in series or parallel to provide the voltage, current, or frequency combinations that would allow the components to handle the greater applied loads. In most cases though , this does not seem to be a problem. Very few, if any, of the ships systems require more than 25 kW to operate effectively. Therefore, we do not view this as being a matter of extreme consequence for the Odysseus project.

Without knowing the complete descriptions of the amount, type, and characteristics of the power each spacecraft system will need, it is impossible to determine the exact composition of each conditioning subsystem. However, the size estimations remain quite accurate for a wide range of combinations.

Batteries were and remain the primary means of electric energy storage. The two major categories of batteries are primary and secondary. Primary batteries are characterized by higher power densities, but are generally not rechargeable. They are typically good for one-time events that require large power with minimum battery mass. Disposable stage platforms often use primary batteries. These batteries are "dry" prior to activation. The batteries are activated when the electrolites "wet" the battery. The electrolites are stored in a connecting container. Secondary batteries have smaller lower energy densities and are more often rechargeable. They are more easily reused. There are several different types of batteries currently used today. Silver-Zinc (AgZn), Nickel-Cadmium (NiCd or NiCad), NiH2, and Lithium. Table 10.3 summarizes some of the characteristics of each battery type. For the Odysseus missions, Lithium batteries or Nickel-Cadmium would probably serve us best. They have the longest lifetime and highest energy storage capacity.

Batteries can be split into multiple battery packs for better packaging, placement, and balance. The only hard-fast requirement is that each battery pack must have twenty-two series-connected cells in order to keep the voltage level where it should be. Another important thing to remember is that we must also include redundant systems for the power system to be complete. Using Nickel-Cadmium batteries and reconditioning them throughout the mission, will allow us to exploit the power from the batteries for the longest amount of time possible. Reconditioning consists of a very high depth of discharge to a point of voltage reversal followed by a tightly controlled recharge.

For the Odysseus I and II missions, batteries will most assuredly be used. However, these will not be the only source of power on the spacecraft. The batteries will be used for power storage and backup power for most situations.

To store the power gathered from the solar array, we needed to devise a battery storage system that stores the solar energy collected during the day for use at night. We also needed a set of power conditioners that convert DC power from the batteries into AC power for other ship systems. The power converters also have the ability to convert the electricity from one voltage/ current level to another.

Initially, Nickel-Cadmium batteries seemed to be the most common, most efficient battery to use. New information has lead us to believe that a commercial Sodium-Sulfur battery may work even better. All indications are that these batteries have higher storage capacity for their mass than NiCad batteries. Both the Nickel-Cadmium and Sodium-Sulfur batteries have estimated efficiency lifetimes of ten years or more. This will be better than what is required for the Odysseus missions.

The batteries will account for the largest mass percentage of the power system. Each battery will weigh approximately 145 lbs and be sized at 33.5"x10"x11". This equilibrates to 0.0604 m³ and 65.8 kg. This data is for the Nickel-Cadmium variety of battery. The Sodium-Sulfur batteries provide comparable power storage capacity and volume, but weigh less. In most ways, the NaS are exact duplicates of the NiCad ones. The differences between the two are insignificant enough that both systems would receive fair consideration for use in this project. We would choose the NaS batteries to save weight. The system will provide power at a 28 V driving force. Each assembly will operate at 40 amp-hrs, delivering 144000 Coulombs of charge. Each battery can store up to 1.44 kW of power. We plan on having a total battery storage capacity of 100 kW (to account for the 80% reclamation efficiency) so that there is a full redundant system for the solar array energy. This accounts for the following total system amounts:

This will account for double the output of the solar arrays and can be used to store energy that is not needed from the fuel cells or other systems. The additional storage capacity also accounts for the redundant system for the solar array energy collection. Four central power busses will be provided power via the power control unit. This system serves as a means of charging and discharging the batteries, as well as monitoring the depth of discharge. This ensures that the batteries do not become overloaded or do not go dead.

Each 50 kW bank of batteries will be subdivided into 10 kW modules so that a failure in a single battery grid will not affect the entire network. This serves to contain failures and distribute the power load evenly. Each grid will be made up of a parallel series of connections, so that this can be accomplished. The batteries provide available DC current only. The power control unit (PCU) serves as a charge controller and can be likened to a power converter. However, the batteries we will use have a separate AC/DC power conversion system. The PCU and charge controller will only be responsible for monitoring depth of discharge and the charging process. They also prevent current from flowing backwards from the battery into the array at night.

The systems for these parts of the mission will follow the same design as the main spacecraft systems. The systems, capacities, and pertinent sizing information will be based completely upon the systems employed in Odysseus II. The final data is highlighted in the following:

TOTAL COST

$107,877,750.00

The lander will supply 30 kW of constant power via the RTG and another 36 kW from the fuel cells, for a total of 66 kW. The power conditioners will be built to handle a

150 kW total load. These power conditioners can be used for both the lander and for the ground team that is using power systems transported from Odysseus I. The rest of the power systems used for the ground operations will be carried on Odysseus I. Once on the planet surface, the ground crew can hook up surface operating systems that require electricity to the lander itself.

The Odysseus I mission will carry the following power system components in support of the surface mission:

The batteries and power conditioners were chosen to support the combined power systems of the surface mission and the lander, whose power will be used when on the surface of Mars. The solar array will partially cover the survival tent that will be built as a domicile and lab for the ground team on Mars. Between the solar array and fuels cells from Odysseus I and the RTGs and fuel cells from the lander, a total of 122 kW of power will be available when surface missions commence.

The total cost analysis is included at the end of the report, as is the solar array sizing chart. The Odysseus missions utilize several different power generation sources and a number of complimentary systems. The power available for each aspect of the mission is more than adequate. There is redundancy in every aspect of the power design. Each system is at least double redundant, with some systems being triple redundant. Modularity is found in many of the systems and subsystems. We have the highest confidence that the power system for the manned Mars mission is comprehensive, safe, and efficient. The last point of interest is the amount of power allocated to each specialty system for the Odysseus II mission. The power allotted does not add up to the total power available. It does, however, account for the sum of the power requested by the other mission design specialists. Note that the pump startup power of 105 kW will only be used until the nuclear reaction feed takes over the electric load for the pump.-*

The total power available for Odysseus II is 214 kW. The lander has a total of 66 kW of power available for total use. There is no subsystem allotment for this system since the lander has not been fully designed. The same is true of Odysseus I. The power has not been divvied into individual system allotments, though a total of 36 kW of power will be usable on the craft itself. A 20 kW solar array will be carried for use on the Mars surface. However, this power will not be available for use by the craft as it transits from low Earth orbit to low Mars orbit.