Patent Description:
Thousands of spacecraft orbit the Earth for performing various functions including, for example, telecommunication, GPS navigation, weather forecasting, and mapping. More complex large spacecraft are also in orbit, including the International Space Station, to which nations throughout the world send crew and supplies for scientific investigation and research. However, spacecraft periodically require servicing to extend their functioning life span. Servicing may include, for example, component repair, refueling, orbit raising, station-keeping, momentum balancing, or other maintenance. Without life extension maintenance, these spacecraft may fall out of service, and replacement is generally extraordinarily expensive and can have a lead time of years. In the case of unmanned spacecraft, to accomplish such servicing, a servicing spacecraft may be sent into orbit to dock with a client spacecraft requiring maintenance, and subsequent to docking, perform life-extending maintenance on the client.

However, spacecraft or other bodies in orbit often possess different electrical potentials. When two spacecraft approach each other, a significant risk arises that an electrostatic discharge could occur between the two spacecraft. Spacecraft contain numerous electronic systems that could be damaged or destroyed by such an electrostatic discharge event. Various patents and publications have considered how to mitigate the risk of an electrostatic discharge event, including <CIT>, <CIT>,<CIT>, <CIT>, <CIT>, <CIT>, <CIT>, <CIT>, <CIT>, <CIT>, and <CIT>. However, an improved system and method for mitigating electrostatic discharge between a first spacecraft and a second spacecraft is desirable.

<CIT> discloses an ion engine <NUM> for plasma generation installed in a first aerospace vehicle <NUM>. The ion engine <NUM> is driven and controlled based on the potential difference in the space between the vehicle <NUM> and a second aerospace vehicle <NUM>, and plasma is discharged toward the second aerospace vehicle <NUM>.

The invention proposes a system for mitigating electrostatic discharge between a first space vehicle and a second space vehicle as presented in claim <NUM> and a method for mitigating electrostatic discharge between a first space vehicle and a second space as presented in claim <NUM>.

Other advantageous and non-limiting features of the invention are presented in the dependent claims.

The drawings included in the present application are incorporated into, and form part of, the specification. They illustrate embodiments of the present disclosure and, along with the description, serve to explain the principles of the disclosure. The drawings are only illustrative of certain embodiments and do not limit the disclosure.

Although embodiments of the disclosure disclosed herein are amenable to various modifications and alternative forms, specifics thereof have been shown by way of example in the drawings and will be described in detail. On the contrary, the intention is to cover all modifications and alternatives falling within the scope of the invention as defined by the claims.

As used herein, the term "substantially" in reference to a given parameter means and includes to a degree that one skilled in the art would understand that the given parameter, property, or condition is met with a small degree of variance, such as within acceptable manufacturing tolerances. For example, a parameter that is substantially met may be at least about <NUM>% met, at least about <NUM>% met, or even at least about <NUM>% met.

The inventors have recognized the risk of damage from electrostatic discharges due to static charge differentials associated with an approach of a first spacecraft to a second spacecraft. In some embodiments, the first spacecraft may comprise a capture assembly that beneficially provides electrostatic mitigation to protect electronic components in the first spacecraft, the second spacecraft, or both. Some embodiments provide systems and methods for reducing the static potential between a first and second spacecraft in a manner that protects the components of both spacecraft.

<FIG> is a side elevation view of two spacecraft in proximity in space according to one embodiment. In some embodiments, first spacecraft <NUM> may be designed to dock to second spacecraft <NUM>. First spacecraft <NUM> may be a servicer spacecraft designed to provide service to second spacecraft <NUM>. According to some embodiments, second spacecraft <NUM> may be a satellite in orbit around a body such as the Earth. If second spacecraft <NUM> is in orbit around Earth, second spacecraft <NUM> may be in low or medium Earth orbit, geosynchronous or above-geosynchronous orbit, or any other orbit.

First spacecraft <NUM> may have a capture apparatus <NUM> with a probe and a propulsion system. The propulsion system of first spacecraft <NUM> may include one or more main thrusters <NUM>, one or more gimbaled thrusters <NUM>, or both. Main thruster <NUM>, gimbaled thrusters <NUM>, or both may be electric propulsion apparatuses. Second spacecraft <NUM> may have an engine <NUM>. Engine <NUM> can be any type of suitable engine or motor for a spacecraft, including a liquid apogee engine or a solid fuel motor. First spacecraft <NUM> may have a first static potential <NUM>, and second spacecraft <NUM> may have a second static potential <NUM>. Upon approach or contact of first spacecraft <NUM> to second spacecraft <NUM>, a differential between first static potential <NUM> and second static potential <NUM> may cause an electrostatic discharge. Such an electrostatic discharge may cause damage to first spacecraft <NUM>, second spacecraft <NUM>, or both, unless the differential between first static potential <NUM> and second static potential <NUM> is mitigated.

<FIG> is a perspective diagram of a passive electrostatic discharge mitigation system <NUM> according to one embodiment. A circuit <NUM>, as show in <FIG>, of passive electrostatic discharge mitigation system <NUM> may be housed within a housing or box <NUM>. Passive electrostatic discharge system <NUM> may be electrically connected to capture apparatus <NUM> or another portion of first spacecraft <NUM>. Passive electrostatic discharge system <NUM> may also be electrically connected to a first electrical contact apparatus <NUM>. First electrical contact apparatus <NUM> may include one or more compliant members (e.g., whiskers <NUM>). Whiskers <NUM> comprise an electrically conductive material. Whiskers <NUM> may be comprised at least in part of beryllium copper.

<FIG> is a diagram of a circuit <NUM> of a passive electrostatic discharge mitigation system <NUM> according to one embodiment. Passive electrostatic discharge mitigation system <NUM> may be configured as a resistance inductance, or RL, circuit comprising one or more resistive elements <NUM> and one or more inductive elements <NUM>. In some embodiments, the one or more inductive elements <NUM> may be an inductor, or one or more ferrite beads, one or more chokes, or another inductive element. The one or more resistive elements <NUM> may be one or more resistors and, in some embodiments, may be configured to provide a resistance of more than <NUM> megaohm and, in some embodiments, may be configured to provide a resistance of greater than or equal to <NUM> megaohms. When first spacecraft <NUM> and second spacecraft <NUM> make contact or come in close enough proximity for a static electric arc to occur between the first spacecraft <NUM> and second spacecraft <NUM>, the passive electrostatic discharge mitigation system <NUM> provides an equalization path for the voltage differential between the two spacecraft and allows the different static charges to equalize.

As a result of the passive electrostatic discharge mitigation system <NUM>, static voltage differential between the two spacecraft <NUM>, <NUM> may be converted into heat to remove energy. This dissipation will reduce, or in some instances eliminate, electrostatic discharges and the amplitude and rise time of any associated voltage spikes that may be detrimental to either spacecraft. In some embodiments, the voltage differential may be discharged over a period of time, for example <NUM>-<NUM> nanoseconds or more. In some embodiments, discharge current may be reduced below <NUM> milliamps by passive electrostatic discharge mitigation system <NUM>. According to certain embodiments, the one or more inductive elements <NUM> and one or more resistive elements <NUM> may be selected to accommodate a transient static potential difference between first spacecraft <NUM> and second spacecraft <NUM> of up to or more than <NUM> kilovolts. In some embodiments, the passive electrostatic discharge mitigation system <NUM> may be configured to have parallel circuit paths that may mitigate the risk of individual component failures.

<FIG> is a perspective view of a housing <NUM> for a passive electrostatic discharge mitigation system <NUM> (<FIG>). Insulated conductor <NUM> provides electrical connection between the passive electrostatic discharge mitigation system <NUM> and at least one first electrical contact apparatus <NUM> (<FIG>), wherein first electrical contact apparatus <NUM> may comprise a compliant member which may be in the form of whisker <NUM> (<FIG>). Insulated grounding conductor <NUM> provides electrical connection between the passive electrostatic discharge mitigation system <NUM> and capture apparatus <NUM> or elsewhere on a body of first spacecraft <NUM> (<FIG>).

<FIG> is a perspective view of a housing <NUM> for a passive electrostatic discharge mitigation system <NUM> (<FIG>) mounted to capture apparatus <NUM>. Insulated conductor <NUM> provides electrical connection between the passive electrostatic discharge mitigation system <NUM> and at least one first electrical contact apparatus <NUM>, wherein first electrical contact apparatus <NUM> may comprise a compliant member such as whisker <NUM>.

<FIG> is a perspective view of first electrical contact apparatus <NUM>. First electrical contact apparatus <NUM> may include one or more compliant members such as whiskers <NUM>. Whiskers <NUM> may comprise a spring element <NUM> that may increase compliance of whiskers <NUM>. Spring element <NUM> may be a torsion spring. Spring element <NUM> may allow whiskers <NUM> to move in a substantially rotational manner when whiskers <NUM> contact engine <NUM> (<FIG>) or another physical structure on the second spacecraft <NUM> (<FIG>). First electrical contact apparatus <NUM> may be designed to be electrically isolated from capture apparatus <NUM>, for example, by one or more insulated posts <NUM> that electrically isolate the conductive components, such as the whiskers <NUM>, from the capture apparatus <NUM>. Insulated posts <NUM> may be comprised of a machinable glass ceramic or other insulating material sufficient to electrically isolate the conductive components. In some embodiments, the conductive components of first electrical contact apparatus <NUM> may be positioned <NUM> inch (<NUM>) or more from the closest conductive component of capture apparatus <NUM>, or another suitable distance to prevent charge creep or arcing.

<FIG> is a perspective view of a capture apparatus <NUM> with a passive electrostatic discharge mitigation system <NUM> mounted thereon approaching the engine <NUM> of the second spacecraft <NUM>. Whiskers <NUM> may be designed to be of a sufficient length to ensure that at least one whisker <NUM> provides the first point of physical contact between first spacecraft <NUM> and second spacecraft <NUM>. Whiskers <NUM> may be designed to be of a sufficient length to ensure that at least one whisker <NUM> is the only physical structure on the first spacecraft <NUM> to come within a distance that would allow a static electric arc between first spacecraft <NUM> and second spacecraft <NUM> before any portion of first spacecraft <NUM> physically contacts second spacecraft <NUM>. In some embodiments, whiskers <NUM> may be at least <NUM> inches (<NUM>) in length.

<FIG> depicts charge potential differentials of a first spacecraft in relation to a second spacecraft for use with an active electrostatic discharge mitigation system, according to one embodiment. <FIG> graphically represents sample anticipated static potential, or charge, differences between various portions of first spacecraft <NUM> and second spacecraft <NUM>. In some embodiments, static potential differences may be on the order of <NUM> kilovolts or more and capacitance between the vehicles may be on the order of <NUM> picofarads or more.

<FIG> and <FIG> depict an active electrostatic discharge mitigation system <NUM> using plasma. The active electrostatic discharge mitigation system <NUM> creates a plasma field that engulfs both the first spacecraft <NUM> and the second spacecraft <NUM>. The active electrostatic discharge mitigation system <NUM> creates the plasma field using one or more electric propulsion engines of first spacecraft <NUM>, which may be main thruster <NUM>, one or more gimbaled thrusters <NUM>, both, or another engine. The one or more electric propulsion engines may be Hall Effect Thrusters. The plasma field created by the active electrostatic discharge mitigation system <NUM> may be low temperature plasma. Active electrostatic discharge mitigation system <NUM> can be operated to reduce the static potential measured to ground reference of each of first spacecraft <NUM> and second spacecraft <NUM>. The reduction of static potential differential between first spacecraft <NUM> and second spacecraft <NUM> may be to a level less than about <NUM> kilovolts, less than about <NUM> kilovolt, less than about <NUM> volts, or less than about <NUM> volts in various embodiments. In addition, use of the active electrostatic discharge mitigation system <NUM> may reduce potential ground bounce between the first spacecraft <NUM> and second spacecraft <NUM>.

In some embodiments, the first spacecraft <NUM> may have both a passive electrostatic discharge mitigation system <NUM> and an active electrostatic discharge mitigation system <NUM>. In such embodiments, active electrostatic discharge mitigation system <NUM> may reduce differential static potential between first spacecraft <NUM> and second spacecraft <NUM> before contact, and passive electrostatic discharge mitigation system <NUM> to mitigate remaining differential static potential between first spacecraft <NUM> and second spacecraft <NUM> upon contact or approach sufficient to permit electrostatic arcing. In such embodiments, passive electrostatic discharge mitigation system <NUM> and active electrostatic discharge mitigation system <NUM> provide redundancy upon component failure of either system.

Claim 1:
A system for mitigating electrostatic discharge between a first space vehicle (<NUM>) and a second space vehicle (<NUM>), the system comprising:
an active electrostatic discharge system (<NUM>) configured to be situated on the first space vehicle (<NUM>), wherein the active electrostatic discharge system (<NUM>) is configured to use an electric propulsion apparatus (<NUM>, <NUM>) on the first space vehicle (<NUM>) to reduce an electric potential between the first space vehicle (<NUM>) and the second space vehicle (<NUM>) by engulfing both the first space vehicle (<NUM>) and the second space vehicle (<NUM>) in a plasma field.