CDR - Power Subsystem - Lamar

Conceptual Design Report


Power Group - Lamar University

Curtis A. Stephenson - Team Leader

Clay Naiser
Jeff Stinnett
Frederick Pouncey
Tariq Naviwala

Conceptual Design Report - Power

Power sources in space may be self-contained or they may draw energy from solar radiation. Three fundamental sources of energy are: chemical, solar, electric and nuclear. Auxiliary power may be derived from batteries, solar cells, nuclear energy, or high-energy fuels. The variety of power sources available and in use confirms that each of the various methods of power generation will have its advantages and disadvantages in the various applications, and that one should attempt a comparative evaluation of the current states-of-the-art in the various power generation fields before selecting a power system for a particular application.

A direct comparison in terms of total energy available per unit weight and volume of various systems is one method of comparison. Solar radiation is available a conservative cost, and the time of usage of a fixed energy supply must be considered in the comparison. Also, the rate of maximum energy expenditure, the duration of energy expenditure, the regulation required, the operating environment, and the power system weight budget will be determining factors.

[Figure not Avaliable]

Figure 1 illustrates the general areas of usefulness of the various space power systems which might be used to generate electrical power. The boundaries indicated are rather indefinite, and change as the various states-of-the-art improve. However, this illustration provides general guidelines for the selection of electrical power systems for various types of projects based on power requirements and mission duration. It is evident from this figure that fuel cells would be an ideal candidate for a two-week mission of processing lunar materials. Re-use in subsequent lunar day periods would pose additional problems, such as reactant (fuel and oxidizer) storage, additional thermal protection, etc. Since Figure 1 is for “primary” power systems, it is not evident from this figure, but another possible combination would be a photovoltaic/battery (secondary, or rechargeable) system. A small radioisotope, thermionic, or thermoelectric generator would also seem to fit this project.


Batteries or electrochemical cells consist of an anode, a cathode, and an electrolyte. The anode is a reducing agent which gives up electrons, the cathode is an oxidizing agent which receives electrons, and the electrolyte serves as an electron barrier which nevertheless permits current flow by ionic conductance within the cell.

The best current batteries are generally held to be nickel-cadmium which produce about 35 Watt-hours per kilogram (Wh/kg). Silver-cadmium batteries are being developed which are expected to produce 52.8 Wh/kg. A study of new materials for battery development has indicated that a magnesium/magnesium-perchlorate/mercuric-oxide cell is capable of 110.2 Wh/kg and 195.2 Wh/ L, and that a magnesium/magnesium-perchlorate or magnesium-bromine/magnesium-dinitrobenzene (DNB) is capable of producing 198.4 Wh/kg. Lozier has estimated the current experimental and maximum electro-mechanical systems, as indicated in Figure 2.

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A fuel cell is essentially a battery with replenishable cathodes and anode materials. This permits the use of light-weight, very active cathode and anode materials. As indicated in Table 1, a hydrogen-oxygen fuel cell is theoretically capable of delivering better that 2.2 kilowatt-hour per kilogram (weight includes fuel). The energy/weight ratio improves with temperature as indicated by Table1, and some provision for thermal agitation (heating) may be desired.

A simplified fuel-cell is portrayed schematically in Figure 3 to illustrate the cell operation. Both anode and cathode are of porous, waterproof material (e.g., carbon on a membrane). Fuel pressurization is commonly used. Sodium amalgam and metallic hydrides are being studied as possible fuels for various handling and ion-exchange reaction reasons. Regenerative techniques and radioisotope activation are being studied as improvements as well. Fuel cells offer a great development potential at the present time, and it is expected that they will come into greater use as their characteristics stabilize.

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The advent of space operations has made feasible the direct tapping of solar radiation energy through the use of solar cells. Solar cells are essentially photo-electric converters of various types which can deliver electrical power in proportion to the intercepted solar radiation power. The characteristics of solar cells may be described in terms of spectral response, output power per unit radiation-collecting area, and life expectancy. Special consideration must be given to the solar eclipse during each orbit period in planning solar cell use. As a special problem, solar cell life and efficiency radiation belts.

Silicon solar cells are usually made from 0.5 mm lapped slices of single-crystal silicon. A typical solar cell is illustrated in Figure 4. The spectral response of silicon solar cells is generally in the 0.5 to 1.0 micron region. Fifteen percent efficiency has been attained. Gallium-arenide solar cells have a spectral response generally in the 0.6 to 0.9 micron range, and provide efficiencies on the order of 10 to 12 percent. Life characteristics are expected to run to ten years. Gallium-arsenide cells are much more (10 to 20 times) resistant to proton bombardment that silicon cells.

A solar cell array must be oriented to face the Sun or else the radiation intercepted will be reduced, as a cosine function. The output of solar cells may be increased by employing large focusing mirrors to collect a much greater amount of solar radiation energy and to concentrate this energy on the photo electric cell. This has the advantage of working the cell at a higher efficiency because the higher density energy conversion will cause a higher operating temperature. Even degraded by reflection and focusing factors, one may expect to obtain more power per mass than if the total collector area were composed of solar cell material for direct energy conversion, and had to support the penalty of interconnecting circuit electrical losses and weight.


Radioisotope power units will probably always use thermoelectric or thermionic conversion devices. They are most attractive in the power range of several kilowatts, and in the lifetime from around one week to a year. Power output decreases exponentially with time, depending on the isotopes half life.

The biggest problem in these power sources is that radioisotopes will probably never be available in sufficient quantity to produce a large number of power units. Another problem is the radiation hazard during launch operations. Also, there is the danger of the isotope contamination the Moon or planets if landed there. Besides the availability and radiation hazard factors, consideration must be given to the half-life, type of decay particles, and cost factors when choosing an isotope for use in a power source. After researching the possible use of radioisotopes, it has become evident to the power group that this particular candidate system would not be very feasible for the lunar Lander mission. The reasons for this are that radioisotopes are dangerous and very costly. Also, radioisotopes seem to be used for missions that will last for up to 10 years, which is entirely too long for this mission.

After completing the above study of various types of power sources, the subsystems to which power will be provided were defined. The first is communications, which will require between 20 to 25 Watts. Next, computers will need approximately 16 to 25 Watts with an additional 6 Watts for information storage. Navigation and guidance will require 40 to 50 Watts, while attitude control will need 25 Watts. Upon arrival at the moon, the oxygen plant will require at least 250 Watts to perform its tasks. Also, a camera, which would require approximately 6 Watts, may be desired to record certain parts of the mission. Other subsystems of which the power requirement are not known include propulsion, landing gear, structures, and thermal control. As these numbers are not set in stone, an assumption was made that a total of 600 Watts of total power will be needed for the mission.

Next, a study of various battery types was made assuming a total load of 600 Watts. The findings of this research will be given as follows:

Primary Batteries: - Batteries that are used for one time use, for they cannot be recharged. The first primary battery type is a Mercad. This type of battery has a capacity of 50 Watt-hour/kg or 180 Watt-hour/L, and has a working voltage of .85 Volts. Another battery type is the Alkaline. Its capacity is 95 Watt-hour/kg or 210 Watt-hour/L, and works at 1.15 Volts. A third battery that was researched was the Solid Electrolyte. This type has a capacity of 150 Watt-hours/kg or 400 Watt-hours/L, with a working voltage of 2.8 V. The Silver Oxide battery is another candidate which has a capacity of 130 Watt-hour/kg or 515 Watt-hour/L. It works at 1.5 Volts.

Secondary Batteries: - Batteries that can be recharged by the use of some other device such as a solar array. The first type of secondary battery is the Nickel - Cadmium. Its capacity is 35 Watt-hour/kg or 80 Watt-hour/L. The working voltage is 1.2 Volts. The next type of battery which was researched is the Zinc Chlorine type. It is one whose capacity is 100 Watt-hour/kg or 130 Watt-hour/L, and a working voltage of 1.9V. Another type is the Edison. This battery has a capacity of 30 Watt-hours/kg or 60 Watt-hours/L. It has a working voltage of 1.2 V. The final two batteries considered were the Silver Zinc and Nickel Zinc. These two batteries have a capacity of 90 Watt-hours/kg (180 Watt-hours/L) and 60 Watt-hours/kg (120 Watt-hours/L) respectively. Their working voltages are 1.5 and 1.6 V.

Working temperature of batteries is typically in the range of 5 to 20 C and solar arrays normally operate between -100 to 100 C. Batteries are an excellent option for short term missions such as this one, especially if accompanied by solar cells. They are small and lightweight, thus making them a prime candidate for this project.

The next power source that was researched was the solar array. This source would be used to recharge batteries if they were chosen for use in this mission. The first type is the Multi-junction Cascade GaAs Cell. It has a nominal thickness of .1397 mm. It has an efficiency of 22% at 1.5 AMO and provides 290 Watts per square meter at 190 Watts per kilogram. The estimated power of this type of cell is 95 Watts. The next type of solar array is GaAs/Ge cells. These cells have a nominal thickness of .1016 mm and have an efficiency of 19.3% at 1.5 AMO. It can provide 245 Watts per square meter at 197 Watts per kilogram. The estimated power output of this type of cell is 116 Watts. A final type of solar array is the AmSi cell. It has a nominal thickness of .0508 mm and an efficiency of 6% at 1.5 AMO. It can provide 80 Watts per square meter at 55 Watts per kilogram. The estimated power output of this type of cell is 25 Watts.

The final item in the research of these different power systems is the fuel cell. One specific cell was found, the Gemini System, that would be more than adequate for this mission. Its mass of 30 kg falls well within the allotted 50 kg. It produces a total of 1000 Watts of power, which would provide a larger power allotment for each subsystem. The fuel cell is 63.5 cm long and 31.7 cm in diameter. Its working voltage is in the range of 26.5 to 23.3 Volts. The reactant consumption rate of the cell is 400 grams per kilowatt-hour. One drawback to this system is that it produces 0.5 Liters of water per kilowatt-hour, thus requiring a means of disposing the water. Finally, it has a decay rate of .004 Volt/hour at 10 to 24 Amps. At this rate, the fuel cell has the ability to last for 260 days.


Space Science and Engineering, Ernst Stuhlinger and Gustav Mesmer, McGraw-Hill Book Company, New York, 1965.

Handbook of Batteries and Fuel Cells, David Linden, McGraw0Hill Book Company, New York, 1984.

Structural Design of Missile and Spacecraft, Lewis H. Abraham, McGraw-Hill Book Company, New York, 1962.

Manned Spacecraft Electrical Power Systems, W. E. Simon and Donald L. Nored, March, 1987.