Electrostatic discharge mitigation for a first spacecraft operating in the vicinity of a second spacecraft and related methods

Abstract

Methods and systems can be used to mitigate or reduce the risk of electrostatic discharge due to electrostatic charge differences between a first space vehicle (10) and a second space vehicle (11) when the first space vehicle approaches the second space vehicle using passive electrostatic discharge mitigation devices (20). In some embodiments, mitigation of electrostatic potential between the first space vehicle and the second space vehicle can be actively achieved by an electric propulsion system (31) disposed on the first space vehicle. In some embodiments, mitigation can be provided by actively and passively mitigating electrostatic potential between the first space vehicle and the second space vehicle.

Description

[0001] This divisional application is based on Chinese Invention Patent Application No. 201880033856.5 (International Application No. PCT / US2018 / 017852), entitled "Method for mitigating electrostatic discharge of a first spacecraft operating near a second spacecraft and related methods", filed on February 12, 2018.

[0002] Priority requirements

[0003] This application claims the benefit of U.S. Provisional Patent Application Serial No. 62 / 484,969, filed April 13, 2017, entitled “Electrostatic Discharge Mitigation for a First Spacecraft Operating in Proximity to a Second Spacecraft,” and U.S. Provisional Patent Application Serial No. 15 / 829,758, filed December 1, 2017, entitled “Electrostatic Discharge Mitigation for a First Spacecraft Operating in Proximity to a Second Spacecraft,” the disclosures of which are hereby incorporated by reference in their entirety. Technical Field

[0004] This disclosure relates to systems and methods for addressing the challenges of electrostatic discharge that arise when a first spacecraft is operating near a second spacecraft (particularly when the first spacecraft approaches the second spacecraft to dock or otherwise contact the second spacecraft). Background Technology

[0005] Thousands of spacecraft orbit the Earth to perform a variety of functions, including telecommunications, GPS navigation, weather forecasting, and mapping. Larger, more complex spacecraft, including the International Space Station, also operate in orbit, and countries around the world send personnel and supplies to these large spacecraft for scientific investigation and research. However, spacecraft require regular maintenance to extend their service life. Maintenance can include, for example, component repair, refueling, orbital lifting, position holding, momentum balancing, or other maintenance. Without life-extending maintenance, these spacecraft may become out of service, and replacements are typically extremely expensive and can take years to arrive. In the case of unmanned spacecraft, to perform such maintenance, a maintenance spacecraft can be sent into orbit to dock with the customer spacecraft requiring maintenance, and life-extending maintenance can be performed on the customer spacecraft after docking.

[0006] However, spacecraft or other objects in orbit typically have different electrical potentials. A significant risk arises when two spacecraft approach each other: electrostatic discharge (ESD) can occur between them. Spacecraft contain numerous electronic systems that could be damaged or destroyed by such ESD events. Various patents and publications have addressed how to mitigate the risks of ESD events, including U.S. Patents Nos. 7,070,151, 7,216,833, 7,216,834, 7,461,818, 7,484,690, 7,575,199, 7,588,213, 7,611,096, 7,611,097, 7,624,950, and 8,205,838, the disclosures of which are incorporated herein by reference in their entirety. However, an improved system and method are needed to mitigate ESD between a first and a second spacecraft. Summary of the Invention

[0007] This document discloses methods and systems for mitigating the risk of electrostatic discharge (ESD) caused by a difference in electrostatic charge between a first spacecraft and a second spacecraft. Various embodiments teach a passive ESD mitigation device that helps to safely reduce the electrostatic potential between the first and second spacecraft before and / or after contact. Some embodiments provide a means for guiding the flow of ESD current in a manner that minimizes the risk to the electronic components of the first and second spacecraft, which can be achieved by providing one or more tentacles electrically connected to the passive ESD mitigation device. Some embodiments provide active mitigation of the ESD between the first and second spacecraft by means of an electric propulsion system disposed on the first spacecraft. Some embodiments provide both active and passive mitigation of the ESD between the first and second spacecraft.

[0008] Some embodiments teach a system and method for mitigating electrostatic discharge (ESD) between a first space launch vehicle and a second space launch vehicle, comprising a passive ESD mitigation system located on the first space launch vehicle, wherein the passive ESD mitigation system includes one or more resistors, and wherein the passive ESD mitigation system includes one or more ferrite beads. In some embodiments, the passive ESD mitigation system reduces the ESD current between the first and second space launch vehicles to less than or equal to about 800 mA for a time period of less than or equal to about 90 nanoseconds. Some embodiments also include one or more tendrils as part of the passive ESD mitigation system.

[0009] Some embodiments teach a system and method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising an active electrostatic discharge system located on the first space launch vehicle, wherein the active electrostatic discharge system manipulates a plasma field. In various embodiments, the controlled plasma field is the plasma field of the first space launch vehicle, the plasma field of the second space launch vehicle, or the plasma field of both the first and second space launch vehicles. The active electrostatic discharge system may use an electric propulsion device to manipulate the plasma field. Some embodiments specify that the plasma field generated by the electric propulsion device is sufficient to reduce the potential between the first and second space launch vehicles to less than approximately + / - 200 volts. The electric propulsion device may be one or more Hall effect thrusters.

[0010] Some embodiments teach a system and method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising: a propulsion mechanism for maneuvering the first space launch vehicle in space; a capture mechanism positioned on the first space launch vehicle for at least temporary engagement of the first and second space launch vehicles; and a mechanism for passively mitigating electrostatic discharge. The mechanism for passively mitigating electrostatic discharge may include one or more compliant members, which in some embodiments may be one or more tentacles. In some embodiments, the compliant member provides a first physical contact between the first and second space launch vehicles. The compliant member may contact the engine of the second space launch vehicle, and in some embodiments, the compliant member always contacts the engine before any other physical structure on the second space launch vehicle. The engine of the second space launch vehicle may be a liquid apogee engine. According to some embodiments, the capture mechanism includes a probe. The compliant member may be located on the probe. Some embodiments specify that the compliant member extends from the first space launch vehicle in a direction toward the second space launch vehicle. The compliant member may be highly compliant. In some embodiments, the compliance of the compliance member is partially provided by a torsion spring. According to some embodiments, the compliance member comprises a beryllium copper alloy.

[0011] Some embodiments teach a system and method for mitigating electrostatic discharge (ESD) between a first space launch vehicle and a second space launch vehicle, comprising a passive ESD mitigation system located on the first space launch vehicle and an active ESD mitigation system located on the first space launch vehicle. In some embodiments, the passive ESD mitigation system includes one or more resistors, and wherein the passive ESD mitigation system includes one or more ferrite beads. In some embodiments, the active ESD system manipulates a plasma field. The passive ESD mitigation system can reduce the ESD current between the first and second space launch vehicles to less than or equal to about 800 mA for a time period of less than or equal to about 90 nanoseconds. The system may also include one or more tentacles. The tentacles can provide a first physical contact between the first and second space launch vehicles and can contact the engine of the second space launch vehicle. The tentacles can extend from the first space launch vehicle in a direction toward the second space launch vehicle. In some embodiments, the tentacles always contact the engine before contacting any other physical structure on the second space launch vehicle. In some embodiments, the capture mechanism may include a probe, and the tentacles may be located on the probe. The tentacles can be highly compliant, and in some embodiments, compliance is partially provided by a torsion spring. In some embodiments, the tentacles comprise a beryllium copper alloy. In some embodiments, the controlled plasma field is the plasma field of a first space launch vehicle, the plasma field of a second space launch vehicle, or the plasma field of both the first and second space launch vehicles. An active electrostatic discharge system can use an electric propulsion device to manipulate the plasma field, which in some embodiments may be one or more Hall effect thrusters. The electric propulsion device can generate a plasma field sufficient to reduce the potential between the first and second space launch vehicles to less than approximately + / - 200 volts.

[0012] The above overview is not intended to describe every illustrated embodiment or every implementation of this disclosure. Attached Figure Description

[0013] The accompanying drawings included in this application are incorporated in and form a part of this specification. They illustrate embodiments of the present disclosure and, together with the description, serve to explain the principles of the disclosure. The drawings are merely illustrative of certain embodiments and do not limit the scope of the disclosure.

[0014] Figure 1 It is a side view of a first and second spacecraft close together according to one or more embodiments.

[0015] Figure 2 This is a perspective view of a passive electrostatic discharge mitigation system according to one or more embodiments.

[0016] Figure 3This is a circuit diagram of a passive electrostatic discharge mitigation system according to one or more embodiments.

[0017] Figure 4 This is a perspective view of a housing for a passive electrostatic discharge mitigation system according to one or more embodiments.

[0018] Figure 5 This is a perspective view of the housing of a passive electrostatic discharge mitigation system for mounting on a capture device, according to one or more embodiments.

[0019] Figure 6 This is a perspective view of a first electrical contact device according to one or more embodiments.

[0020] Figure 7 It is a perspective view of a capture device according to one or more embodiments, wherein a passive electrostatic discharge mitigation system mounted on the capture device is close to the engine of a second spacecraft.

[0021] Figure 8 It is a graphical representation of the sampled expected electrostatic potential difference between various parts of a first spacecraft and a second spacecraft according to one or more embodiments.

[0022] Figure 9A and Figure 9B It is a graphical representation of the operation of an active electrostatic discharge mitigation system generating plasma according to one or more embodiments.

[0023] While the embodiments disclosed herein can be modified and alternatively implemented in various ways, their details have been shown by way of example in the accompanying drawings and will be described in detail. However, it should be understood that this disclosure is not intended to limit it to the specific embodiments described. Rather, it is intended to cover all modifications, equivalents, and alternatives that fall within the scope of this disclosure. Detailed Implementation

[0024] As used herein, the term "substantially" for a given parameter means, and to a certain extent includes, the degree of variation of a given parameter, characteristic, or condition as understood by those skilled in the art, such as being within acceptable manufacturing tolerances. For example, a substantially satisfied parameter could be satisfied at least about 90%, at least about 95%, or even at least about 99%.

[0025] The inventors have recognized that electrostatic discharge resulting from a difference in electrostatic charge associated with the proximity of a first spacecraft to a second spacecraft poses a risk of damage. In some embodiments, the first spacecraft may include a trapping assembly that advantageously provides electrostatic mitigation to protect the electronic components of the first spacecraft, the second spacecraft, or both. Some embodiments provide systems and methods for reducing the electrostatic potential between the first and second spacecraft in a manner that protects components of both spacecraft.

[0026] Figure 1 This is a side view of two spacecraft approaching each other in space according to one embodiment. In some embodiments, the first spacecraft 10 may be designed to dock with the second spacecraft 11. The first spacecraft 10 may be a service spacecraft designed to provide services to the second spacecraft 11. According to some embodiments, the second spacecraft 11 may be a satellite in orbit around a host (such as Earth). If the second spacecraft 11 is in orbit around Earth, it may be in low Earth orbit or medium Earth orbit, geosynchronous orbit or above-geosynchronous orbit, or any other orbit.

[0027] The first spacecraft 10 may have a capture device 23 with a probe and propulsion system. The propulsion system of the first spacecraft 10 may include one or more main thrusters 17, one or more gimbal thrusters 18, or both. The main thrusters 17, gimbal thrusters 18, or both may be electric propulsion devices. The second spacecraft 11 may have an engine 19. The engine 19 may be any type of suitable engine or motor for a spacecraft (including liquid apogee engines or solid fuel engines). The first spacecraft 10 may have a first electrostatic potential 12, and the second spacecraft 11 may have a second electrostatic potential 13. When the first spacecraft 10 approaches or contacts the second spacecraft 11, the difference between the first electrostatic potential 12 and the second electrostatic potential 13 may cause an electrostatic discharge. Unless the difference between the first electrostatic potential 12 and the second electrostatic potential 13 is mitigated, such an electrostatic discharge may damage the first spacecraft 10, the second spacecraft 11, or both.

[0028] Figure 2 This is a perspective view of a passive electrostatic discharge mitigation system 20 according to one embodiment. Figure 3 As shown, the circuitry 32 of the passive electrostatic discharge mitigation system 20 can be housed within a housing or box 24. The passive electrostatic discharge system 20 can be electrically connected to the capture device 23 or another part of the first spacecraft 10. The passive electrostatic discharge system 20 can also be electrically connected to the first electrical contact device 25. The first electrical contact device 25 may include one or more compliant members (e.g., whiskers 26). Whiskers 26 comprise a conductive material. Whiskers 26 may be at least partially constructed of a beryllium copper alloy.

[0029] Figure 3This is a diagram of circuit 32 of a passive electrostatic discharge (ESD) mitigation system 20 according to one embodiment. The passive ESD mitigation system 20 can be configured as a resistive-inductor (RIL) circuit, which includes one or more resistive elements 27 and one or more inductor elements 28. In some embodiments, the one or more inductor elements 28 may be inductors, or one or more ferrite beads, one or more chokes, or another inductor element. The one or more resistive elements 27 may be one or more resistors, and in some embodiments may be configured to provide a resistance greater than 1 megohm, and in some embodiments may be configured to provide a resistance greater than or equal to 15 megohms. When the first spacecraft 10 and the second spacecraft 11 come into contact or are sufficiently close to generate an electrostatic arc between the first spacecraft 10 and the second spacecraft 11, the passive ESD mitigation system 20 provides an equalization path for the voltage difference between the two spacecraft and allows the different electrostatic charges to reach equilibrium.

[0030] As a result of the passive electrostatic discharge mitigation system 20, the static voltage difference between the two spacecraft 10, 11 can be converted into heat to remove energy. This dissipation will reduce or, in some cases, eliminate the magnitude and rise time of electrostatic discharge and any associated voltage spikes that may be harmful to either spacecraft. In some embodiments, the voltage difference can discharge over a period of, for example, 50 to 90 nanoseconds or longer. In some embodiments, the discharge current can be reduced to below 800 mA by the passive electrostatic discharge mitigation system 20. According to some embodiments, the one or more inductor elements 28 and one or more resistor elements 27 may be selected to accommodate transient static potential differences of up to or greater than 10 kV between the first spacecraft 10 and the second spacecraft 11. In some embodiments, the passive electrostatic discharge mitigation system 20 may be configured to have parallel circuit paths that can mitigate the risk of individual component failures.

[0031] Figure 4 It is used in passive electrostatic discharge mitigation system 20 ( Figure 2 A perspective view of the housing 24. The insulating conductor 21 provides the passive electrostatic discharge mitigation system 20 with at least one first electrical contact device 25. Figure 2 The electrical connection between the first electrical contact device 25 may include a compliant member, which may be in the form of a whisker 26. Figure 2 The insulated grounding conductor 22 provides an electrical connection between the passive electrostatic discharge mitigation system 20 and the capture device 23 or other locations on the body of the first spacecraft 10. Figure 1 ).

[0032] Figure 5 It is a passive electrostatic discharge mitigation system 20 for installation on the capture device 23. Figure 2A perspective view of the housing 24. An insulating conductor 21 provides an electrical connection between the passive electrostatic discharge mitigation system 20 and at least one first electrical contact device 25, wherein the first electrical contact device 25 may include compliant members, such as whiskers 26.

[0033] Figure 6 This is a perspective view of the first electrical contact device 25. The first electrical contact device 25 may include one or more compliant members, such as a whisker 26. The whisker 26 may include a spring element 29, which may increase the compliance of the whisker 26. The spring element 29 may be a torsion spring. When the whisker 26 contacts the engine 19 ( Figure 2 ) or another physical structure on the second spacecraft 11 ( Figure 1 When the spring element 29 is engaged, the tendril 26 can be moved in a substantially rotational manner. The first electrical contact 25 can be designed to be electrically isolated from the capture device 23, for example, by one or more insulating posts 30, which electrically isolate conductive parts such as the tendril 26 from the capture device 23. The insulating posts 30 can be made of machinable glass-ceramic or other insulating materials sufficient to electrically isolate conductive parts. In some embodiments, the conductive parts of the first electrical contact 25 can be positioned 0.25 inches (0.635 cm) or more away from the nearest conductive component of the capture device 23, or at another suitable distance to prevent charge creep or arcing.

[0034] Figure 7 This is a perspective view of the capture device 23, on which a passive electrostatic discharge mitigation system 20 mounted is positioned close to the engine 19 of the second spacecraft 11. Tentacles 26 may be designed to be of sufficient length to ensure that at least one tentacle 26 provides the first point of physical contact between the first spacecraft 10 and the second spacecraft 11. Tentacles 26 may be designed to be of sufficient length to ensure that at least one tentacle 26 is the only physical structure on the first spacecraft 10 that enters within a distance that would allow an electrostatic arc to be generated between the first spacecraft 10 and the second spacecraft 11 before any part of the first spacecraft 10 makes physical contact with the second spacecraft 11. In some embodiments, the length of tentacle 26 may be at least 6 inches (15.24 cm).

[0035] Figure 8 The charge potential difference between a first spacecraft and a second spacecraft, used in conjunction with an active electrostatic discharge mitigation system according to one embodiment, is depicted. Figure 8 The sampled expected static potential difference or charge difference between the various parts of the first spacecraft 10 and the second spacecraft 11 is represented graphically. In some embodiments, the static potential difference may be approximately 10 kilovolts or greater, and the capacitance between the launch vehicles may be approximately 100 picofarads or greater.

[0036] Figure 9A and Figure 9B An active electrostatic discharge mitigation system 31 using plasma is depicted. In some embodiments, the active electrostatic discharge mitigation system 31 can generate a plasma field that engulfs the first spacecraft 10 and the second spacecraft 11. In some embodiments, the active electrostatic discharge mitigation system 31 uses one or more electric propulsion engines of the first spacecraft 10 to generate the plasma field, which can be a main thruster 17, one or more gimbal thrusters 18, both of these, or another engine. The one or more electric propulsion engines can be Hall effect thrusters. The plasma field generated by the active electrostatic discharge mitigation system 31 can be cryogenic plasma. The active electrostatic discharge mitigation system 31 can be operated to reduce the static potential measured relative to a ground reference for each of the first spacecraft 10 and the second spacecraft 11. In various embodiments, the reduction in the static potential difference between the first spacecraft 10 and the second spacecraft 11 can reach levels of less than about 5 kW, less than about 1 kW, less than about 200 volts, or less than about 100 volts. Additionally, the use of the active electrostatic discharge mitigation system 31 can reduce potential ground bounce between the first spacecraft 10 and the second spacecraft 11.

[0037] In some embodiments, the first spacecraft 10 may have both a passive electrostatic discharge (ESD) mitigation system 20 and an active ESD mitigation system 31. In such embodiments, the active ESD mitigation system 31 can reduce the differential electrostatic potential between the first spacecraft 10 and the second spacecraft 11 before contact, and the passive ESD mitigation system 20 can alleviate the residual differential ESD potential between the first spacecraft 10 and the second spacecraft 11 when they come into contact or are sufficiently close to generate an electrostatic arc. In such embodiments, the passive ESD mitigation system 20 and the active ESD mitigation system 31 provide redundancy in the event of a component failure in either system.

[0038] Other example embodiments are disclosed below.

[0039] Example 1: A system for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising a passive electrostatic discharge mitigation system located on the first space launch vehicle, wherein the passive electrostatic discharge mitigation system comprises one or more resistors.

[0040] Example 2: According to the system described in Example 1, the passive electrostatic discharge mitigation system reduces the electrostatic discharge current between the first space launch vehicle and the second space launch vehicle to less than or equal to about 800 mA within a time period of less than or equal to about 90 nanoseconds.

[0041] Example 3: According to the system described in Example 1, the passive electrostatic discharge mitigation system further includes one or more inductive elements selected from the group consisting of inductors, ferrite beads, and chokes.

[0042] Example 4: The system according to Example 1, wherein the system further includes one or more tentacles.

[0043] Example 5: A system and method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising an active electrostatic discharge system located on the first space launch vehicle, wherein the active electrostatic discharge system manipulates a plasma field.

[0044] Example 6: The system according to Example 5, wherein the plasma field is the plasma field of the first space launch vehicle.

[0045] Example 7: The system according to Example 5, wherein the plasma field is the plasma field of the second space launch vehicle.

[0046] Example 8: The system according to Example 5, wherein the active electrostatic discharge system uses an electric propulsion device to manipulate the plasma field.

[0047] Example 9: According to the system described in Example 8, the electric propulsion device generates a plasma field sufficient to reduce the potential between the first space launch vehicle and the second space launch vehicle to less than approximately + / - 200 volts.

[0048] Example 10: The system according to Example 8, wherein the electric propulsion device is one or more Hall effect thrusters.

[0049] Example 11: A system and method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising: a propulsion mechanism for maneuvering the first space launch vehicle in space; a capture mechanism positioned on the first space launch vehicle for at least temporarily engaging the first space launch vehicle and the second space launch vehicle; and a mechanism for passively mitigating electrostatic discharge during engagement of the first space launch vehicle and the second space launch vehicle.

[0050] Example 12: The system according to Example 11, wherein the mechanism for passively mitigating electrostatic discharge includes one or more tentacles.

[0051] Example 13: The system according to Example 12, wherein the tendrils provide a first physical contact between the first space launch vehicle and the second space launch vehicle.

[0052] Example 14: The system according to Example 13, wherein the tentacles are configured to contact the engine of the second space launch vehicle when the first space launch vehicle and the second space launch vehicle are engaged.

[0053] Example 15: The system according to Example 14, wherein when the first space launch vehicle and the second space launch vehicle are engaged, the tendrils are positioned to contact the engine before any other physical structure on the second space launch vehicle.

[0054] Example 16: The system according to Example 15, wherein the engine is a liquid apogee engine.

[0055] Example 17: The system according to Example 12, wherein the capture mechanism includes a probe.

[0056] Example 18: The system according to Example 17, wherein the tendrils are located on the probe.

[0057] Example 19: The system according to Example 13, wherein there are multiple tentacles.

[0058] Example 20: The system according to Example 13, wherein when the first space launch vehicle and the second space launch vehicle are engaged, the tendrils extend from the first space launch vehicle in a direction toward the location of the second space launch vehicle.

[0059] Example 21: The system according to Example 13, wherein the whiskers are configured to be highly compliant.

[0060] Example 22: The system according to Example 21, wherein the compliance is partially provided by a torsion spring.

[0061] Example 23: The system according to Example 13, wherein the tentacles comprise a beryllium copper alloy.

[0062] Example 24: A method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising: providing a passive electrostatic discharge mitigation system on the first space launch vehicle, wherein the passive electrostatic discharge mitigation system includes one or more resistors; and configuring a first contact between the first space launch vehicle and the second space launch vehicle to ensure that any electrostatic discharge between the first space launch vehicle and the second space launch vehicle is guided through the passive electrostatic discharge mitigation system.

[0063] Example 25: The method according to Example 24, wherein the method further includes using the passive electrostatic discharge mitigation system to reduce the electrostatic discharge current between the first space launch vehicle and the second space launch vehicle to less than or equal to about 800 mA for a time period of less than or equal to about 90 nanoseconds.

[0064] Example 26: According to the method of Example 24, the passive electrostatic discharge mitigation system further includes one or more inductive elements selected from the group consisting of inductors, ferrite beads, and chokes.

[0065] Example 27: According to the method of Example 24, the passive electrostatic discharge mitigation system further includes one or more tentacles.

[0066] Example 28: A method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising: providing an active electrostatic discharge system located on the first space launch vehicle; and using the active electrostatic discharge system to manipulate a plasma field.

[0067] Example 29: According to the method of Example 28, the manipulation of the plasma field includes manipulating the plasma field of the first space launch vehicle.

[0068] Example 30: The method according to Example 28, wherein the manipulation of the plasma field includes manipulating the plasma field of the second space launch vehicle.

[0069] Example 31: The method according to Example 28, wherein the manipulation of the plasma field includes manipulating the plasma field using an electric propulsion device.

[0070] Example 32: The method according to Example 31, wherein the method further includes using the electric propulsion device to generate a plasma field sufficient to reduce the potential between the first space launch vehicle and the second space launch vehicle to less than about + / - 200 volts.

[0071] Example 33: According to the method of Example 31, the electric propulsion device is one or more Hall effect thrusters.

[0072] Example 34: A method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising: providing a capture mechanism located on the first space launch vehicle for at least temporarily engaging the first space launch vehicle and the second space launch vehicle; maneuvering the first space launch vehicle close to the second space launch vehicle in space; and providing a mechanism for passively mitigating electrostatic discharge between the first space launch vehicle and the second space launch vehicle.

[0073] Example 35: The method according to Example 34, wherein the mechanism for passively mitigating electrostatic discharge includes a whisker.

[0074] Example 36: The method according to Example 35, wherein the method further includes making the first space launch vehicle and the second space launch vehicle physically contact each other, wherein the tentacles provide the first physical contact between the first space launch vehicle and the second space launch vehicle when they engage.

[0075] Example 37: The method according to Example 36, wherein the tentacles are configured to contact the engine of the second space launch vehicle when the first space launch vehicle and the second space launch vehicle are engaged.

[0076] Example 38: According to the method of Example 37, the tendrils are configured to contact the engine before any other physical structure on the second space launch vehicle when the first space launch vehicle and the second space launch vehicle are engaged.

[0077] Example 39: The method according to Example 38, wherein the engine is a liquid apogee engine.

[0078] Example 40: The method according to Example 35, wherein the capture mechanism includes a probe.

[0079] Example 41: The method according to Example 40, wherein the tendrils are located on the probe.

[0080] Example 42: The method according to Example 36, wherein there are multiple tentacles.

[0081] Example 43: According to the method of Example 36, when the first space launch vehicle and the second space launch vehicle are engaged, the tendrils extend from the first space launch vehicle in a direction toward the location of the second space launch vehicle.

[0082] Example 44: The method according to Example 36, wherein the tendrils are highly compliant.

[0083] Example 45: The method according to Example 44, wherein the compliance is partially provided by a torsion spring.

[0084] Example 46: The method according to Example 36, wherein the tentacles comprise a beryllium copper alloy.

[0085] Example 47: A system for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising: a passive electrostatic discharge mitigation system located on the first space launch vehicle, wherein the passive electrostatic discharge mitigation system includes one or more resistors; and an active electrostatic discharge system located on the first space launch vehicle, wherein the active electrostatic discharge system manipulates a plasma field.

[0086] Example 48: According to the system described in Example 47, the passive electrostatic discharge mitigation system reduces the electrostatic discharge current between the first space launch vehicle and the second space launch vehicle to less than or equal to about 800 mA for a time period of less than or equal to about 90 nanoseconds.

[0087] Example 49: According to the system of Example 47, the passive electrostatic discharge mitigation system further includes one or more inductive elements selected from the group consisting of inductors, ferrite beads, and chokes.

[0088] Example 50: The system according to Example 47, wherein the system further includes one or more tentacles.

[0089] Example 51: The system according to Example 47, wherein the plasma field is the plasma field of the first space launch vehicle.

[0090] Example 52: The system according to Example 47, wherein the plasma field is the plasma field of the second space launch vehicle.

[0091] Example 53: The system according to Example 47, wherein the active electrostatic discharge system uses an electric propulsion device to manipulate the plasma field.

[0092] Example 54: According to the system of Example 53, the plasma field generated by the electric propulsion device is sufficient to reduce the potential between the first space launch vehicle and the second space launch vehicle to less than about + / - 200 volts.

[0093] Example 55: The system according to Example 53, wherein the electric propulsion device is one or more Hall effect thrusters.

[0094] Example 56: The system according to Example 49, wherein the tentacles are configured to provide the first physical contact between the first space launch vehicle and the second space launch vehicle.

[0095] Example 57: The system according to Example 56, wherein the tentacles are configured to contact the engine of the second space launch vehicle.

[0096] Example 58: The system according to Example 57, wherein the tendrils are configured to contact the engine before any other physical structure on the second space launch vehicle.

[0097] Example 59: The system according to Example 49, wherein the capture mechanism includes a probe.

[0098] Example 60: The system according to Example 59, wherein the tendrils are located on the probe.

[0099] Example 61: The system according to Example 60, wherein there are multiple tentacles.

[0100] Example 62: According to the system of Example 60, when the first space launch vehicle and the second space launch vehicle are engaged, the tendrils extend from the first space launch vehicle in a direction toward the position where the second space launch vehicle is located.

[0101] Example 63: The system according to Example 60, wherein the tendrils are highly compliant.

[0102] Example 64: The system according to Example 63, wherein the compliance is partially provided by a torsion spring.

[0103] Example 65: The system according to Example 60, wherein the tentacles comprise a beryllium copper alloy.

[0104] Example 66: A method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, comprising: providing an active electrostatic discharge system located on the first space launch vehicle; manipulating a plasma field using the active electrostatic discharge system; providing a passive electrostatic discharge mitigation system on the first space launch vehicle, wherein the passive electrostatic discharge mitigation system includes one or more resistors; and configuring a first contact between the first space launch vehicle and the second space launch vehicle to ensure that any electrostatic discharge between the first space launch vehicle and the second space launch vehicle is guided through the passive electrostatic discharge mitigation system.

[0105] Example 67: The method according to Example 66, wherein the method further includes using the passive electrostatic discharge mitigation system to reduce the electrostatic discharge current between the first space launch vehicle and the second space launch vehicle to less than or equal to about 800 mA for a time period of less than or equal to about 90 nanoseconds.

[0106] Example 68: According to the method of Example 66, the passive electrostatic discharge mitigation system further includes one or more inductive elements selected from the group consisting of inductors, ferrite beads, and chokes.

[0107] Example 69: According to the method of Example 66, the passive electrostatic discharge mitigation system further includes one or more tentacles.

[0108] Example 70: The method according to Example 66, wherein the manipulation of the plasma field includes manipulating the plasma field of the first space launch vehicle.

[0109] Example 71: The method according to Example 66, wherein the manipulation of the plasma field includes manipulating the plasma field of the second space launch vehicle.

[0110] Example 72: The method according to Example 66, wherein the manipulation of the plasma field includes manipulating the plasma field using an electric propulsion device.

[0111] Example 73: The method according to Example 72, wherein the method further includes using the electric propulsion device to generate a plasma field sufficient to reduce the potential between the first space launch vehicle and the second space launch vehicle to less than about + / - 200 volts.

[0112] Example 74: The method according to Example 72, wherein the electric propulsion device is one or more Hall effect thrusters.

[0113] Example 75: The method according to Example 69, wherein the method further includes making the first space launch vehicle and the second space launch vehicle physically contact each other, wherein the tentacles provide a first physical contact between the first space launch vehicle and the second space launch vehicle when they engage.

[0114] Example 76: The method according to Example 75, wherein the tentacles are configured to contact the engine of the second space launch vehicle.

[0115] Example 77: The method according to Example 76, wherein the tentacles are configured to contact the engine before any other physical structure on the second space launch vehicle.

[0116] Example 78: The method according to Example 76, wherein the engine is a liquid apogee engine.

[0117] Example 79: The method according to Example 68, wherein the capture mechanism includes a probe.

[0118] Example 80: The method according to Example 79, wherein the tendrils are located on the probe.

[0119] Example 81: According to the method of Example 80, wherein when the first space launch vehicle and the second space launch vehicle are engaged, the tendrils extend from the first space launch vehicle in a direction toward the location of the second space launch vehicle.

[0120] Example 82: The method according to Example 69, wherein the tendrils are highly compliant.

[0121] Example 83: The method according to Example 82, wherein the compliance is partially provided by a torsion spring.

[0122] Example 84: The method according to Example 68, wherein the tentacles comprise a beryllium copper alloy.

[0123] The embodiments of this disclosure described above and illustrated in the accompanying drawings do not limit the scope of this disclosure, which is covered by the appended claims and their legal equivalents. Any equivalent embodiments are within the scope of this disclosure. In fact, various modifications to the content of this disclosure (such as alternative useful combinations of described elements) will be apparent to those skilled in the art beyond what is shown and described herein. Such modifications and embodiments also fall within the scope of the appended claims and their equivalents. The terminology used herein is chosen to explain the principles of the embodiments, practical applications of technology found in the market, or technical improvements, or to enable others of ordinary skill in the art to understand the embodiments disclosed herein.

Claims

1. A system for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, the system comprising: An active electrostatic discharge system is configured to be located on the first space launch vehicle, wherein the active electrostatic discharge system is configured to use an electric propulsion device on the first space launch vehicle to reduce the potential between the first space launch vehicle and the second space launch vehicle by engulfing both the first space launch vehicle and the second space launch vehicle in a plasma field.

2. The system of claim 1, wherein, The active electrostatic discharge system is configured to manipulate the plasma field of only one of the first or second space launch vehicles.

3. The system of claim 2, wherein, The electric propulsion device is configured to generate the plasma field, which is configured to reduce the potential between the first and second space launch vehicles to less than approximately + / - 200 volts.

4. The system according to claim 2 or 3, wherein, The electric propulsion device is one or more Hall effect thrusters.

5. The system according to claim 2 or 3, further comprising a passive electrostatic discharge mitigation system located on the first space launch vehicle, wherein, The passive electrostatic discharge mitigation system includes one or more resistors and one or more compliant members electrically connected to the one or more resistors.

6. The system according to claim 5, wherein, The one or more compliant members include tentacles.

7. The system according to claim 5, wherein, The one or more compliant members are configured to provide a first physical contact between the first space launch vehicle and the second space launch vehicle when the first space launch vehicle and the second space launch vehicle are engaged.

8. The system according to claim 5, wherein, The one or more compliant members are positioned on the first space launch vehicle to extend from the first space launch vehicle in a direction toward the position where the first space launch vehicle is positioned when the first space launch vehicle and the second space launch vehicle are engaged.

9. A system for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, the system comprising: A propulsion mechanism for maneuvering the first space launch vehicle in space; A capture mechanism, which is to be positioned on the first space launch vehicle for at least temporary engagement of the first space launch vehicle and the second space launch vehicle; as well as A mechanism for mitigating electrostatic discharge, the mechanism being configured to use the propulsion mechanism on the first space launch vehicle to reduce the potential between the first space launch vehicle and the second space launch vehicle by engulfing the first space launch vehicle, the second space launch vehicle, or both the first space launch vehicle and the second space launch vehicle in a plasma field.

10. The system of claim 9, further comprising a mechanism for passively mitigating electrostatic discharge, wherein, The mechanism for passively mitigating electrostatic discharge includes one or more resistors and one or more compliant members electrically connected to the one or more resistors, wherein the one or more compliant members comprise a conductive material, and wherein the one or more compliant members are configured such that one of the one or more compliant members provides a first physical contact between the first spacecraft and the second spacecraft when the first spacecraft and the second spacecraft are engaged.

11. The system according to claim 10, wherein, The compliance of the one or more compliant members is provided at least in part by a torsion spring.

12. The system according to claim 10, wherein, The one or more compliant members are electrically isolated from the capture mechanism.

13. The system of claim 12, further comprising one or more insulating posts for electrically isolating the one or more compliant members from the capture mechanism.

14. A method for mitigating electrostatic discharge between a first space launch vehicle and a second space launch vehicle, the method comprising: To rendezvous between the first and second space launch vehicles; Electrostatic discharge between the first space launch vehicle and the second space launch vehicle is mitigated by an electrostatic discharge system configured to use an electric propulsion device on the first space launch vehicle to reduce the potential between the first space launch vehicle and the second space launch vehicle by engulfing the first space launch vehicle, the second space launch vehicle, or both the first space launch vehicle and the second space launch vehicle in a plasma field. as well as The first space launch vehicle and the second space launch vehicle are secured by a capture mechanism positioned on at least one of the first space launch vehicle or the second space launch vehicle, so that the first space launch vehicle and the second space launch vehicle are engaged at least temporarily.

15. The method of claim 14, further comprising engulfing both the first space launch vehicle and the second space launch vehicle in a plasma field.

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