Electricity: battery or capacitor charging or discharging – Wind – solar – thermal – or fuel-cell source
Reexamination Certificate
2002-02-20
2002-10-01
Tso, Edward H. (Department: 2838)
Electricity: battery or capacitor charging or discharging
Wind, solar, thermal, or fuel-cell source
Reexamination Certificate
active
06459232
ABSTRACT:
BACKGROUND
1. Field of the Invention
The present invention relates to the field of spacecraft power sources and charging effects on spacecraft, and in particular, the discharging of spacecraft and charging of batteries in a plasma field.
2. Description of the Related Art
The total weight of any satellite or other spacecraft is proportional to the launch cost and must therefore be minimized. Spacecraft power systems, therefore, typically rely primarily on solar arrays. However, spacecraft within the solar system may travel through two distinct phases—a non-eclipse phase, where the sun transfers energy directly to the spacecraft, and an eclipse phase, where the spacecraft is fully or partially obscured from the sun. During the non-eclipse phase, power requirements are generally met or exceeded through the solar arrays. However, during the eclipse phase, where the solar arrays are ineffective at supplying power, at least one battery is often needed to meet power requirements.
The use of rechargeable batteries on spacecraft is known in the art. See, for example, U.S. Pat. No. 6,157,161, issued Dec. 5, 2000 to Canter et al. Such systems typically utilize energy derived from the solar arrays during non-eclipse phases, in excess of power requirements, to recharge the batteries for later use during eclipse phases. The solar array panels must therefore be larger than necessary to meet typical power requirements during the eclipse phases. This increases the weight and often the complexity of the spacecraft, resulting in higher costs. In addition, solar array panels are exposed to potentially harmful elements, such as space debris, leading to arcing and may potentially be destroyed or damaged. Spacecraft anomalies due to solar array degradation, which often occurs suddenly, are commonly reported. By contrast, batteries located inside spacecraft are better protected by the spacecraft; therefore, spacecraft components dependent on battery power are more likely to receive the power needed for longer periods of time than those dependent on solar arrays.
Spacecraft are surrounded in space by plasma in which atoms and molecules have been dissociated into positively charged ions and negatively charged electrons. In other words, plasma is generally composed of ions and “free electrons,” which are no longer bound to their atoms or molecules.
Ions are considerably heavier and slower than electrons. This accounts for the fact that while space plasma is neutral, the ion current at a given location will be smaller than the electron current. At geosynchronous altitudes, for example, the ambient electron flux is approximately two orders of magnitude (100 times) larger than that of ions.
A spacecraft initially exposed to space plasma will therefore begin to accumulate a negative charge. This negative charge will then begin to repel further electrons while attracting the positive ions. At some point, the spacecraft reaches equilibrium where no further charging of the spacecraft occurs and the surface potentials are in equilibrium. This process occurs rapidly, typically over a period of milliseconds, and generally results in a net negative potential of the spacecraft. By way of example, the time required to reach equilibrium—which is mostly a function of surface capacitance—is approximately one millisecond for typical satellites at geosynchronous altitude.
The negative potential may reach thousands of volts during active periods, such as geomagnetic storms and substorms. During geomagnetic substorms, which occur almost daily, the electrical current intercepted by typical spacecraft is typically two orders of magnitude higher than during quite periods. The current rises even higher during geomagnetic storms—which occur several times each year, depending on the sun's activity cycle, and likely follow significant solar eruptions of hot plasmas or coronal mass ejections toward the Earth. The spacecraft surface charging effect occurs in both eclipse and non-eclipse phases.
The current intercepted is proportional to the exposed surface area of the spacecraft. This negative charging is often harmful to onboard electronics, telemetry, and spacecraft survivability. Sudden onset of spacecraft charging is particularly harmful, as it is more difficult if not impossible to take appropriate action to mitigate the effects.
In addition to “absolute charging,” where the entire spacecraft potential differs from the ambient plasma, “differential charging” frequently occurs where parts of the spacecraft are charged to different potentials relative to each other. Both absolute and differential charging are potentially harmful to spacecraft components and even to the survivability of the spacecraft. It is therefore desirable to mitigate the harmful effects of spacecraft charging by discharging the spacecraft.
Accordingly, there is a need for a lightweight spacecraft power system, protected from the harmful effects of space, that provides sufficient power in both eclipse and non-eclipse phases and that mitigates harmful spacecraft charging effects on system components.
SUMMARY
To achieve these and other objects, the present invention provides a spacecraft power system that discharges excess voltage potential on the spacecraft while re-charging on-board batteries, thereby resulting in both an increased spacecraft power supply and a reduction in the potential harm to spacecraft components.
Empirical evidence from a variety of satellites (e.g., ATS-5, ATS-6, SCATHA, and Los Alamos National Laboratory geosynchronous satellites LANL-1989-046, LANL-1990-095, LANL-1994-084, and LANL-97A) indicates the existence of a critical or threshold temperature (T
c
,) at which spacecraft charging onsets. The concept of critical temperature, developed by the inventor, may be determined for any given material by methods known in the art. It can be further shown that the critical temperature is independent of the plasma density in a Maxwellian space plasma. See “Spacecraft Charging at Geosynchronous Altitudes: New Evidence of Existence of Critical Temperature,” S.T. Lai et al.,
Journal of spacecraft and Rockets
, vol. 38, no. 6, pp. 922-928, November-December, 2001, which is herein incorporated by reference herein. The critical temperature depends on the surface properties. Once the critical temperature is reached, there is a strong correlation between electron temperature and spacecraft potential.
By utilizing different materials on a spacecraft surface, each having corresponding differences in the critical temperature, each material surface will generally have an electrical potential relative to the other material surfaces. For example, one embodiment includes a spacecraft with an aluminum oxide (AL
2
O
3
) surface on one side and a silicon dioxide (SiO
2
) surface on the other side. Above the critical temperature, these materials will exhibit a potential relative to each other. See
FIG. 3
, which illustrates the approximate relative potential between these two surface materials, as well as for Magnesium Fluoride (MgF
2
) and Kapton, for electron temperatures up to 6.0 keV. The graph presented in
FIG. 3
assumes a spherical satellite positioned in the eclipse phase where the electron temperature is equal to the ion temperature and where the ambient electron flux is ten times the ambient ion flux.
By connecting each surface material to a battery or batteries, the excess electrical charge can be used to recharge the battery power supply. As the electrical charges on the surfaces are channeled to the battery power supply, the surface potential decreases accordingly. In this manner, the natural space plasma energy can be harnessed to provide power for the spacecraft while the potential harm from differential surface charge potentials is substantially reduced.
The material surfaces may be selected and designed to best accommodate any particular spacecraft, depending on the spacecraft environment, orientation with respect to the sun and other bodies, the potential harm due to spacecraft charging, and on the particular power needs of the spacec
Sayeedi Stephen A.
The United States of America as represented by the Secretary of
LandOfFree
Spacecraft power system does not yet have a rating. At this time, there are no reviews or comments for this patent.
If you have personal experience with Spacecraft power system, we encourage you to share that experience with our LandOfFree.com community. Your opinion is very important and Spacecraft power system will most certainly appreciate the feedback.
Profile ID: LFUS-PAI-O-2998826