DETAILED ACTION
Notice of Pre-AIA or AIA Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Status of Claims
This Office Action is in response to the application filed on 23 December 2025. Claims 1-3, 8-13, 18-19, 22, 24, and 26-32 are presently pending and are presented for examination. Claims 4-7, 14-17, 20-21, 23, and 25 are cancelled.
Response to Amendments
In response to Applicant’s amendments dated 23 December 2025, Examiner withdraws the previous claims objections; withdraws the previous 35 U.S.C. 112(a) rejections; and withdraws the previous prior art rejections.
Response to Arguments
Applicant's arguments, see Remarks, filed 23 December 2025, have been fully considered but they are not persuasive.
Applicant's arguments, see Remarks, pg. 13-15, fail to comply with 37 CFR 1.111(b) because they amount to a general allegation that the claims define a patentable invention without specifically pointing out how the language of the claims patentably distinguishes them from the references. For these reasons, examiner is unpersuaded and maintains the corresponding rejections. For a detailed explanation of how prior art teaches the amended limitations, see the Claim Rejections - 35 USC § 103 section, below.
The remaining arguments are essentially the same as those addressed above and/or below and are unpersuasive for at least the same reasons. Therefore, examiner is unpersuaded and maintains the corresponding rejections.
Claim Objections
Claim(s) 29 is/are objected to because of the following informalities: “second direction until the desired time to achieve the constant phase offset has elapsed” should be “second direction until the desired time to achieve the constant phase offset has elapsed.” (inserted a period). Appropriate correction is required.
Claim Rejections - 35 USC § 103
In the event the determination of the status of the application as subject to AIA 35 U.S.C. 102 and 103 (or as subject to pre-AIA 35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis (i.e., changing from AIA to pre-AIA ) for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of 35 U.S.C. 103 which forms the basis for all obviousness rejections set forth in this Office action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
The factual inquiries for establishing a background for determining obviousness under 35 U.S.C. 103 are summarized as follows:
1. Determining the scope and contents of the prior art.
2. Ascertaining the differences between the prior art and the claims at issue.
3. Resolving the level of ordinary skill in the pertinent art.
4. Considering objective evidence present in the application indicating obviousness or nonobviousness.
This application currently names joint inventors. In considering patentability of the claims the examiner presumes that the subject matter of the various claims was commonly owned as of the effective filing date of the claimed invention(s) absent any evidence to the contrary. Applicant is advised of the obligation under 37 CFR 1.56 to point out the inventor and effective filing dates of each claim that was not commonly owned as of the effective filing date of the later invention in order for the examiner to consider the applicability of 35 U.S.C. 102(b)(2)(C) for any potential 35 U.S.C. 102(a)(2) prior art against the later invention.
Claim(s) 1-2, 8-12, 18, and 26-32 is/are rejected under 35 U.S.C. 103 as being unpatentable over US-20180155065-A1, hereinafter “Silva” (previously of record), in view of CN-113697129-A, hereinafter “Bai” (previously of record) and “Orbital Mechanics for Engineering Students (Fourth Edition)” (2020), hereinafter “Curtis” (previously of record).
Regarding claim 11, and analogous claim 1, Silva discloses a control system operable for controlling a phased transfer of multiple spacecraft from a separation orbit to a target orbit (Silva, para. 0006: “In another example, a system [i.e., control system] for deploying a constellation of spacecraft [i.e., multiple spacecraft] is described. The system comprises a cluster of spacecraft in orbit at a first orbit [i.e., from a separation orbit], and a ground station in communication with spacecraft of the cluster of spacecraft when the spacecraft of the cluster are visible to the ground station. The ground station sends a first command to each spacecraft in the cluster of spacecraft indicating to separate from each other so that relative phasing of each spacecraft [i.e., controlling a phased transfer] to each other minimizes overlapping visibility periods from the ground station, and after the spacecraft have separated, the ground station sends a second command to each of the spacecraft indicating to raise as separated simultaneously in a synchronized ascent to a respective final orbit [i.e., to a target orbit].”),
Regarding claim 1, Silva also discloses a method for controlling a phased transfer of multiple spacecraft from a separation orbit to a target orbit…” (Silva, para. 0004: “In one example, a method of deploying a constellation of spacecraft [i.e., method for controlling…multiple spacecraft] is described.”).
the multiple spacecraft comprising a lead spacecraft and a follower spacecraft, the control system comprising (Silva, para. 0062: “In FIG. 6, a spacecraft [i.e., a lead spacecraft] in the cluster 102 [i.e., follower spacecraft] begins a maneuver to increase or decrease the orbit size (altitude), and therefore induce a drift rate with respect to the spacecraft remaining in the cluster 102 [i.e., follower spacecraft]...A resulting drift orbit 156 of the spacecraft 104 [i.e., a lead spacecraft] causes the spacecraft 104 to be separated from the cluster 102 [i.e., follower spacecraft].”):
a processor (Silva, para. 0051: “Within examples, the processor(s) 116 of the ground station 114 can execute the executable instructions 136 stored in the data storage 120 to perform functions of sending a first command to each spacecraft 104, 106, 108, and 110 in the cluster 102 indicating to separate from each...”); and
a tangible, non-transitory computer readable storage medium on which is recorded an instruction set, wherein execution of the instruction set by the processor causes the control system to (Silva, para. 0046: “The computer-readable storage media can include volatile and/or non-volatile storage components, such as optical, magnetic, organic or other memory or disc storage, which can be integrated in whole or in part with the processor(s) 116. The data storage 120 [i.e., a tangible, non-transitory computer readable storage medium] is considered non-transitory computer readable media.”; para. 0047: “The data storage 120 thus is a non-transitory computer readable storage medium, and executable instructions 136 are stored thereon [i.e., on which is recorded an instruction set]. The executable instructions 136 include computer executable code. When the executable instructions 136 are executed by the processor(s) 116, the processor(s) 116 are caused to perform functions.”):
while maintaining an in-phase relationship of the multiple spacecraft relative to each other within the separation orbit (Silva, para. 0062: “In FIG. 6, a spacecraft in the cluster 102 begins a maneuver to increase or decrease the orbit size (altitude) [i.e., separation orbit], and therefore induce a drift rate with respect to the spacecraft remaining in the cluster 102. Over time, the drift incurs a phasing or separation relative to the spacecraft remaining in the cluster 102. The spacecraft 104 stops the burn maneuver when the desired drift rate from the cluster 102 is obtained. In this example, a desired drift rate may result in about a phase separation angle of 90° [i.e., maintaining an in-phase relationship].”):
compute respective desired trajectories for the lead spacecraft and the follower spacecraft to reach the target orbit (Silva, para. 0072: “Following separation, each of the spacecraft 104, 106, 108, and 110 [i.e., the lead spacecraft and the follower spacecraft] is raised, as separated, simultaneously in a synchronized ascent to their respective final orbits [i.e., compute respective desired trajectories]. Thus, the separation can be achieved by raising all the spacecraft by a small amount (e.g., 1000 km to 1400 km) sequentially to provide phasing (separation) at a lower orbit, and then a timed simultaneous ascent of all spacecraft 104, 106, 108, and 110 results in the spacecraft 104, 106, 108, and 110 arriving at their final orbits [i.e., spacecraft to reach the target orbit] retaining the desired separation.”); and
establish a constant phase offset between the multiple spacecraft in a mean anomaly of the separation orbit (Silva, para. 0073: “For example, as the last spacecraft 110 approaches the drift altitude, all the spacecraft 104, 106, 108, and 110 begin their pre-loaded burn to simultaneously maneuver from about the same starting drift orbit 156 with 90° phase separation from each other [i.e., a constant phase offset between the multiple spacecraft in mean anomaly of the separation orbit]. The ground station 114 may send a timed command to each of the spacecraft 104, 106, 108, and 110, when visible to the ground station 114, indicating when to maneuver to the respective final orbit. The ground station 114 individually commands the spacecraft 104, 106, 108, and 110 to separate and then raise in altitude. The timed command indicates when to begin the maneuver (e.g., a start time) in order to achieve the synchronized ascent and to maintain the relative phasing of each spacecraft to each other.”);
during a series of transfer orbits of the multiple spacecraft from the separation orbit to the target orbit, apply the respective desired trajectories such that the constant phase offset is maintained and the follower spacecraft is simultaneously transferred to the target orbit in-phase with the lead spacecraft (Silva, para. 0073: “For example, as the last spacecraft 110 approaches the drift altitude, all the spacecraft 104, 106, 108, and 110 begin their pre-loaded burn to simultaneously maneuver from about the same starting drift orbit 156 with 90° phase separation from each other [i.e., series of transfer orbits…apply the respective desired trajectories such that the constant phase offset is maintained]. The ground station 114 may send a timed command to each of the spacecraft 104, 106, 108, and 110, when visible to the ground station 114, indicating when to maneuver to the respective final orbit [i.e., the follower spacecraft is simultaneously transferred to the target orbit in-phase with the lead spacecraft]. The ground station 114 individually commands the spacecraft 104, 106, 108, and 110 to separate and then raise in altitude. The timed command indicates when to begin the maneuver (e.g., a start time) in order to achieve the synchronized ascent and to maintain the relative phasing of each spacecraft to each other.”);
calculate a delta phase rate needed to maintain the constant phase offset between a sequential spacecraft pair of the multiple spacecraft in one or more of the series of transfer orbits as a difference between a respective mean motion of the sequential spacecraft pair (Silva, para. 0090: “As shown at block 216, additional functions can include separating the spacecraft in the cluster 102 relative to each other to cause a separation phase from spacecraft to spacecraft [i.e., calculate a delta phase rate needed to maintain the constant phase offset between a sequential spacecraft pair of the multiple spacecraft].”; para. 0091: “As shown at block 218, additional functions can include sending a timed command to each of the spacecraft, when visible to the ground station 114, indicating to maneuver [i.e., one or more of the series of transfer orbits difference between a respective mean motion of the sequential spacecraft pair] to the respective final orbit 142, and the timed command indicates when to begin the maneuver in order to achieve the synchronized ascent [i.e., in one or more the series of transfer orbits] and to maintain the relative phasing of each spacecraft to each other.”),
wherein the delta phase rate is based on the difference between the respective mean motion (Silva, para. 0082: “At block 202, the method 200 includes releasing the cluster 102 of spacecraft from the launch vehicle 150 at the first orbit 140. Within examples, after releasing the cluster 102 of spacecraft from the launch vehicle 150 at the first orbit 140, relative phasing of each spacecraft to each other is about zero degrees [i.e., the delta phase rate is based on the difference between the respective mean motion].”; para. 0090: “As shown at block 216, additional functions can include separating the spacecraft in the cluster 102 relative to each other to cause a separation phase from spacecraft to spacecraft.”); and
Silva does not appear to disclose the following:
cause the multiple spacecraft to maintain the constant phase offset between the sequential spacecraft pair based on applying the delta phase rate to the multiple spacecraft by changing a relative semi-major axis between the sequential spacecraft pair at a defined rate according to a predetermined relationship, wherein phase offset is provided based on a first SMA of a first spacecraft of the multiple spacecraft that is moved away from a second SMA of a second spacecraft of the multiple spacecraft, wherein the first spacecraft and the second spacecraft are a sequential pair, and wherein the predetermined relationship comprises: changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration; and changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed.
However, in the same field of endeavor, Bai teaches:
cause the multiple spacecraft to maintain the constant phase offset between the sequential spacecraft pair based on applying the delta phase rate to the multiple spacecraft by changing a relative semi-major axis (SMA) between the sequential spacecraft pair at a defined rate according to a predetermined relationship (translated document of Bai, para. 5: “The relative control strategy realizes the maintenance of the constellation configuration by controlling the relative position between the satellites”; para. 49: “In the solution provided by the embodiment of the present application, before the orbit prediction is performed in the preset forecast time period according to the preset orbit perturbation equation to obtain the orbit elements, it is also necessary to set the basic parameters in advance, where the basic parameters include…set the nominal phase difference between satellites in the same orbital plane [i.e., a predetermined relationship] and the error tolerance of the phase deviation.”; para. 69: “Step 102: Determine whether there is a phase deviation between any adjacent satellites [i.e., cause the multiple spacecraft to maintain the constant phase offset between the sequential spacecraft pair based on applying the delta phase rate] that exceeds a preset error tolerance.”; para. 70: “Step 103, if it exists, calculate the orbit semi-major axis correction amount corresponding to each satellite in the satellite constellation according to the preset optimization target, and perform orbit control for each satellite according to the orbit semi-major axis correction amount [i.e., changing a relative semi-major axis between the sequential spacecraft pair at a defined rate].”),
wherein phase offset is provided based on a first SMA of a first spacecraft of the multiple spacecraft that is moved away from a second SMA of a second spacecraft of the multiple spacecraft, wherein the first spacecraft and the second spacecraft are a sequential pair (translated document of Bai, para. 5: “The relative control strategy realizes the maintenance of the constellation configuration by controlling the relative position between the satellites”; para. 69: “Step 102: Determine whether there is a phase deviation between any adjacent satellites [i.e., wherein the first spacecraft and the second spacecraft are a sequential pair] that exceeds a preset error tolerance.”; para. 70: “Step 103, if it exists, calculate the orbit semi-major axis correction amount corresponding to each satellite in the satellite constellation according to the preset optimization target, and perform orbit control for each satellite according to the orbit semi-major axis correction amount [i.e., phase offset is provided based on a first SMA of a first spacecraft of the multiple spacecraft that is moved away from a second SMA of a second spacecraft of the multiple spacecraft].”), and
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, with the concept of changing a relative semi-major axis between a sequential spacecraft pair at a defined rate according to a predetermined relationship, taught by Bai, in order to control the spacecraft to their target orbit in the most efficient manner possible while also maintaining the desired constellation configuration (translated document of Bai, para. 8: “The technical problem solved by this application is: in view of the problem that phase control in the prior art is not conducive to extending the on-orbit lifetime of the constellation, this application provides a method and computer equipment for relative phase control of the constellation. In the solution provided in the embodiment of the application, , On the one hand, to minimize the overall fuel consumption of the constellation as the goal, to find the phase correction amount that meets the phase difference constraint between the satellites, to ensure the consistency of the fuel consumption of each satellite to a certain extent, and to ensure the balance of the fuel consumption rate of each satellite It is beneficial to extend the on-orbit life of the constellation from the perspective of fuel consumption; on the other hand, the full information crossover between the various satellites is beneficial to the independent maintenance of the entire constellation configuration and the coordination of mission planning.”).
Silva and Bai not appear to explicitly disclose the following:
wherein the predetermined relationship comprises: changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration; and changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed.
However, in the same field of endeavor, Curtis teaches:
wherein the predetermined relationship comprises: changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration; and changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed (Curtis, Chapter 6, section 6.5 Phasing maneuvers: "A phasing maneuver is a two-impulse Hohmann transfer from [i.e., in a first direction until no more than half of a desired time to achieve the constant phase offset has elapsed] and then back to the same orbit [i.e., thereafter changing the semi-major axis of the designated spacecraft in a second direction opposite the first direction until the desired time to achieve the constant phase offset has elapsed], as illustrated in Fig. 6.10. The Hohmann transfer ellipse is the phasing orbit with a period selected to return the spacecraft to the main orbit within a specified time [i.e., a desired time]. Phasing maneuvers are used to change the position of a spacecraft in its orbit.").
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, as modified by Bai, with the concept of a two-impulse Hohmann transfer, taught by Curtis, in order to complete a phasing maneuver that changes a position of one spacecraft in its orbit, in order to maintain a constant phase offset between two spacecraft (Curtis, Chapter 6, section 6.5: “Phasing maneuvers are used to change the position of a spacecraft in its orbit.”).
Regarding claims 12, and analogous claim 2, Silva, Bai, and Curtis teach the control system of claim 11, and Silva further discloses the following:
wherein the execution of the instruction set by the processor causes the control system to apply the respective desired trajectories by modifying an in-plane change in velocity (Δv) component of each of the multiple spacecraft in relation to an out-of-plane Δv component for each of the multiple spacecraft (Silva, para. 0006: “The ground station sends a first command to each spacecraft in the cluster of spacecraft [i.e., the execution of the instruction set by the processor causes the control system] indicating to separate from each other [i.e., to apply the respective desired trajectories by modifying an in-plane change in velocity (Δv) component of each of the multiple spacecraft] so that relative phasing of each spacecraft to each other minimizes overlapping visibility periods from the ground station, and after the spacecraft have separated, the ground station sends a second command to each of the spacecraft indicating to raise as separated simultaneously in a synchronized ascent to a respective final orbit [i.e., in relation to an out-of-plane Δv component for each of the multiple spacecraft].”).
Regarding claim 8, Silva, Bai, and Curtis teach the method of claim 1, and Silva further discloses the following:
wherein maintaining the constant phase offset includes periodically repeating the method during the series of transfer orbits to account for thrust variations and/or mass variations between the multiple spacecraft (Silva, para. 0037: “Following deployment of the spacecraft cluster from a launch vehicle, a method of relative spatial separation from spacecraft to spacecraft is described to create an orbit phase between the spacecraft at some staging orbit. After phasing of the spacecraft within the staging orbit, the constellation ascends in a simultaneous fashion keeping their relative phasing [i.e., maintaining the constant phase offset].”; para. 0100: “Thus, the ground station 114 can perform these functions by sending commands to each of the spacecraft in the cluster 102 to increase or decrease in altitude in a sequential manner [i.e., repeating the method during the series of transfer orbits], and sending commands to each spacecraft indicating a start time for the synchronized ascent, for example.”; para. 0101: “The ground station 114 can send commands at appropriate times to the spacecraft [i.e., repeating the method during the series of transfer orbits], when the spacecraft are visible to the ground station 114. For example, once a first spacecraft is separated from the cluster 102 a desired amount of distance, the ground station 114 sends a command to a second spacecraft to initiate spacing separation, and so on.”; para. 0086: “The method 200 may take a month, 6 months, a year, etc. to complete the full separation and ascent of the spacecraft to the final orbit 142 depending on an altitude of the first orbit 140, and altitude of the final orbit 142, a mass of the spacecraft [i.e., account for thrust variations and/or mass variations between the multiple spacecraft] and other factors to consider.”).
Regarding claim 9, Silva, Bai, and Curtis teach the method of claim 1, and Silva further discloses the following:
further comprising: deploying the multiple spacecraft from a single launch vehicle (Silva, para. 0001: “The present disclosure relates generally to a system for deploying a constellation of spacecraft [i.e., multiple spacecraft], and more particularly, to a procedure to separate spacecraft deployed from a single launch vehicle from each other in such a way that during ascent maneuvers…”).
Regarding claim 10, Silva, Bai, and Curtis teach the method of claim 9, and Silva further discloses the following:
wherein deploying the multiple spacecraft from the single launch vehicle includes deploying multiple satellites from the single launch vehicle (Silva, para. 0041: “The spacecraft 104, 106, 108, and 110 are shown to be satellites. However, in other examples, the spacecraft 104, 106, 108, and 110 may include other vehicles for orbit per specific missions, and the spacecraft 104, 106, 108, and 110 can include different combinations of spacecraft as well depending on a specific mission.”).
Regarding claim 18, Silva, Bai, and Curtis teach the control system of claim 11, and Silva further discloses the following:
wherein the execution of the instruction set by the processor causes the control system to: maintain the constant phase offset during the series of transfer orbits to account for thrust variations and/or mass variations between the multiple spacecraft (Silva, para. 0037: “Following deployment of the spacecraft cluster from a launch vehicle, a method of relative spatial separation from spacecraft to spacecraft is described to create an orbit phase between the spacecraft at some staging orbit. After phasing of the spacecraft within the staging orbit, the constellation ascends in a simultaneous fashion keeping their relative phasing [i.e., maintain the constant phase offset].”; para. 0086: “The method 200 may take a month, 6 months, a year, etc. to complete the full separation and ascent of the spacecraft to the final orbit 142 depending on an altitude of the first orbit 140, and altitude of the final orbit 142, a mass of the spacecraft [i.e., account for thrust variations and/or mass variations between the multiple spacecraft] and other factors to consider.”).
Regarding claim 27, Silva, Bai, and Curtis teach The method of claim 1, and Silva further discloses the following:
wherein a respective mean motion
n
of the respective mean motion is calculated based on:
n
=
180
°
π
μ
a
3
where
μ
is an Earth gravitational parameter and
a
is a semi-major axis (SMA) value of a spacecraft of the sequential spacecraft pair (Silva, para. 0042: “The spacecraft 104, 106, 108, and 110 may be configured to revolve around the Earth (or other celestial body) in respective orbits. In some examples, the orbits of the spacecraft 104, 106, 108, and 110 may have some inclination angle relative to an orbital plane in which the target orbit lies.”; Note: This claim recites aspects of Kepler’s 3rd Law of orbital motion. It would be obvious to one of ordinary skill in the art, at the time of the application, to know that Kepler’s 3rd Law is used for orbital calculations of spacecraft orbiting Earth (see “Methods of Celestial Mechanics”; published by Academic Press, 1961; pages 20-21).).
Regarding claim 28, and analogous claim 26, Silva, Bai, and Curtis teach The control system of claim 11, and Silva further discloses the following:
wherein a respective mean motion of the respective mean motion is calculated based on an Earth gravitational parameter and a semi-major axis (SMA) value of a spacecraft of the sequential spacecraft pair (Silva, para. 0042: “The spacecraft 104, 106, 108, and 110 may be configured to revolve around the Earth (or other celestial body) in respective orbits. In some examples, the orbits of the spacecraft 104, 106, 108, and 110 may have some inclination angle relative to an orbital plane in which the target orbit lies.”; Note: This claim recites aspects of Kepler’s 3rd Law of orbital motion. It would be obvious to one of ordinary skill in the art, at the time of the application, to know that Kepler’s 3rd Law is used for orbital calculations of spacecraft orbiting Earth (see “Methods of Celestial Mechanics”; published by Academic Press, 1961; pages 20-21).).
Regarding claim 29, Silva, Bai, and Curtis teach The method of claim 1, and Curtis further teaches the following:
wherein predetermined relationship comprises: changing the first SMA in relation to the second SMA in the first predetermined direction until no more than half of a desired time to achieve the constant phase offset has elapsed; and changing the first SMA in the second direction until the desired time to achieve the constant phase offset has elapsed (Curtis, Chapter 6, section 6.5 Phasing maneuvers: "A phasing maneuver is a two-impulse Hohmann transfer from [i.e., in a first direction until no more than half of a desired time to achieve the constant phase offset has elapsed] and then back to the same orbit [i.e., thereafter changing the semi-major axis of the designated spacecraft in a second direction opposite the first direction until the desired time to achieve the constant phase offset has elapsed], as illustrated in Fig. 6.10. The Hohmann transfer ellipse is the phasing orbit with a period selected to return the spacecraft to the main orbit within a specified time [i.e., a desired time]. Phasing maneuvers are used to change the position of a spacecraft in its orbit.").
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, as modified by Bai and Curtis, with the concept of a two-impulse Hohmann transfer, taught by Curtis, in order to complete a phasing maneuver that changes a position of one spacecraft in its orbit, in order to maintain a constant phase offset between two spacecraft (Curtis, Chapter 6, section 6.5: “Phasing maneuvers are used to change the position of a spacecraft in its orbit.”).
Regarding claim 31, Silva, Bai, and Curtis teach The method of claim 30, and Silva further discloses the following:
wherein the single launch vehicle is associated with a high-thrust multi-stage rocket (Silva, para. 0057: “A single launch vehicle 150 is used, and the launch vehicle 150 may take many forms include a spacecraft or rocket ship, for example.”).
Regarding claim 32, and analogous claim 30, Silva, Bai, and Curtis teach The control system of claim 11, and Silva further discloses the following:
wherein the multiple spacecraft are deployed from a single launch vehicle (Silva, para. 0001: “The present disclosure relates generally to a system for deploying a constellation of spacecraft [i.e., multiple spacecraft], and more particularly, to a procedure to separate spacecraft deployed from a single launch vehicle from each other in such a way that during ascent maneuvers…”).
Claim(s) 3, 13, 19, 22, and 24 is/are rejected under 35 U.S.C. 103 as being unpatentable over Silva, in view of Bai and Curtis, as applied to claims 2, 12, 1, and 11, respectively, above, and further in view of US-8930048-B1, hereinafter “Batla” (previously of record).
Regarding claim 13, and analogous claim 3, Silva, Bai, and Curtis teach the control system of claim 12, but do not appear to explicitly teach the following:
wherein the execution of the instruction set by the processor causes the control system to: implement a compound steering control law having a plurality of weight factors; and change the in-plane Δv component in relation to the out-of-plane Δv component by manipulating a predetermined one of the plurality of weight factors over a duration of the series of transfer orbits to thereby control a mean motion of each of the multiple spacecraft.
However, in the same field of endeavor, Batla teaches:
wherein the execution of the instruction set by the processor causes the control system to: implement a compound steering control law having a plurality of weight factors (Batla, col. 3, line 4: “In one or more embodiments, a method for application of a compound steering law for efficient low thrust transfer orbit trajectory for a spacecraft involves calculating, with at least one processor [i.e., execution of the instruction set by the processor causes the control system to], a desired orbit for the spacecraft.”; col. 11, line 37: “As previously mentioned, enhanced compound steering law [i.e., compound steering control law] offers a shorter mission duration as compared to existing solutions. It also offers more a flexible mission design, i.e. it allows for achieving non-time sensitive orbital targets (e.g., the semi-major axis, the eccentricity, and the inclination) either simultaneously, or at different times, using strategy based weighting factors [i.e., a plurality of weight factors].”); and
change the in-plane Δv component in relation to the out-of-plane Δv component by manipulating a predetermined one of the plurality of weight factors over a duration of the series of transfer orbits to thereby control a mean motion of each of the multiple spacecraft (Batla, col. 3, line 63: “In addition, the method involves calculating, with at least one processor, the direction of the vector sum of the velocity change [i.e., change the in-plane Δv component in relation to the out-of-plane Δv component] required to achieve the orbit eccentricity [i.e., in-plane Δv component], the velocity change required to achieve the semi-major axis [i.e., in-plane Δv component], and the velocity change required to achieve the orbit inclination [i.e., out-of-plane Δv component].”; col. 4, line 64: “In one or more embodiments, at least one processor uses at least one weighting factor for the calculating of the direction of the vector sum [i.e., mean motion of each of the multiple spacecraft].”; col. 11, line 37: “As previously mentioned, enhanced compound steering law offers a shorter mission duration as compared to existing solutions. It also offers more a flexible mission design, i.e. it allows for achieving non-time sensitive orbital targets (e.g., the semi-major axis, the eccentricity, and the inclination) either simultaneously, or at different times, using strategy based weighting factors [i.e., manipulating a predetermined one of the plurality of weight factors].”; col. 11, line 7: “In particular, these existing solutions for low thrust transfer strategies [i.e., over a duration of the series of transfer orbits] (also referred to as baseline solutions in this document) involve dividing the transfer into two phases.”).
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, as modified by Bai and Curtis, with the concept of implementing a compound steering law and using weighting factors to adjust in-plane change-in-velocity components in relation to out-of-plane change-in-velocity components of a spacecraft, taught by Batla, in order to efficiently (i.e., save time and/or energy, etc.) cause the spacecraft to transfer from one orbit to another (Batla, col. 7, line 7: “One advantage is that, while the existing prior art steering strategies impart the required changes to the orbit semi-major axis and eccentricity in a piece-meal fashion, the disclosed compound steering law can accomplish both objectives (i.e. the orbital eccentricity and the semi-major axis targets) efficiently at the same time. Instead of discontinuous segments and abrupt changes in the spacecraft's steering and configuration, the disclosed compound steering law enables continuous firing of the spacecraft's thrusters and a smooth transition of the spacecraft's orientation. Another advantage is that, except for extreme cases when one of the orbital objectives is predominating, the disclosed compound steering law reduces the total transfer orbit mission duration (TOD) as compared to the existing prior art solutions.”).
Regarding claim 19, Silva discloses a method for controlling a phased transfer of multiple spacecraft from a separation orbit to a target orbit, the multiple spacecraft comprising a lead spacecraft and a follower spacecraft (Silva, para. 0004: “In one example, a method of deploying a constellation of spacecraft is described. The method comprises releasing a cluster of spacecraft from a launch vehicle at a first orbit, separating spacecraft in the cluster of spacecraft from each other to minimize overlapping visibility periods from a ground station, and raising each of the spacecraft as separated simultaneously in a synchronized ascent to a respective final orbit.”), the method comprising:
deploying the multiple spacecraft into the separation orbit from a single launch vehicle (Silva, para. 0001: “The present disclosure relates generally to a system for deploying a constellation of spacecraft [i.e., multiple spacecraft], and more particularly, to a procedure to separate spacecraft deployed from a single launch vehicle from each other in such a way that during ascent maneuvers…”);
while maintaining an in-phase relationship of the multiple spacecraft relative to each other within the separation orbit (Silva, para. 0062: “In FIG. 6, a spacecraft in the cluster 102 begins a maneuver to increase or decrease the orbit size (altitude) [i.e., separation orbit], and therefore induce a drift rate with respect to the spacecraft remaining in the cluster 102. Over time, the drift incurs a phasing or separation relative to the spacecraft remaining in the cluster 102. The spacecraft 104 stops the burn maneuver when the desired drift rate from the cluster 102 is obtained. In this example, a desired drift rate may result in about a phase separation angle of 90° [i.e., maintaining an in-phase relationship].”)…
…computing, via the control system, respective desired trajectories for the lead spacecraft and the follower spacecraft to reach the target orbit (Silva, para. 0072: “Following separation, each of the spacecraft 104, 106, 108, and 110 [i.e., the lead spacecraft and the follower spacecraft] is raised, as separated, simultaneously in a synchronized ascent to their respective final orbits [i.e., computing, via the control system, respective desired trajectories]. Thus, the separation can be achieved by raising all the spacecraft by a small amount (e.g., 1000 km to 1400 km) sequentially to provide phasing (separation) at a lower orbit, and then a timed simultaneous ascent of all spacecraft 104, 106, 108, and 110 results in the spacecraft 104, 106, 108, and 110 arriving at their final orbits [i.e., spacecraft to reach the target orbit] retaining the desired separation.”); and
establishing a constant phase offset between the multiple spacecraft in mean anomaly of the separation orbit (Silva, para. 0073: “For example, as the last spacecraft 110 approaches the drift altitude, all the spacecraft 104, 106, 108, and 110 begin their pre-loaded burn to simultaneously maneuver from about the same starting drift orbit 156 with 90° phase separation from each other [i.e., a constant phase offset between the multiple spacecraft in mean anomaly of the separation orbit]. The ground station 114 may send a timed command to each of the spacecraft 104, 106, 108, and 110, when visible to the ground station 114, indicating when to maneuver to the respective final orbit. The ground station 114 individually commands the spacecraft 104, 106, 108, and 110 to separate and then raise in altitude. The timed command indicates when to begin the maneuver (e.g., a start time) in order to achieve the synchronized ascent and to maintain the relative phasing of each spacecraft to each other.”); and
during a series of transfer orbits of the multiple spacecraft from the separation orbit to the target orbit, applying the respective desired trajectories computed via the control system such that the constant phase offset is maintained and the follower spacecraft is simultaneously transferred to the target orbit in-phase with the lead spacecraft (Silva, para. 0073: “For example, as the last spacecraft 110 approaches the drift altitude, all the spacecraft 104, 106, 108, and 110 begin their pre-loaded burn to simultaneously maneuver from about the same starting drift orbit 156 with 90° phase separation from each other [i.e., series of transfer orbits…applying the respective desired trajectories computed via the control system such that the constant phase offset is maintained]. The ground station 114 may send a timed command to each of the spacecraft 104, 106, 108, and 110, when visible to the ground station 114, indicating when to maneuver to the respective final orbit [i.e., the follower spacecraft is simultaneously transferred to the target orbit in-phase with the lead spacecraft]. The ground station 114 individually commands the spacecraft 104, 106, 108, and 110 to separate and then raise in altitude. The timed command indicates when to begin the maneuver (e.g., a start time) in order to achieve the synchronized [i.e., simultaneously] ascent and to maintain the relative phasing of each spacecraft to each other [i.e., such that the constant phase offset is maintained].”)…
…calculating a delta phase rate needed to maintain the constant phase offset between a sequential spacecraft pair of the multiple spacecraft in one or more of the series of transfer orbits as a difference between a respective mean motion of the sequential spacecraft pair (Silva, para. 0090: “As shown at block 216, additional functions can include separating the spacecraft in the cluster 102 relative to each other to cause a separation phase from spacecraft to spacecraft [i.e., calculating a delta phase rate needed to maintain the constant phase offset between a sequential spacecraft pair of the multiple spacecraft].”; para. 0091: “As shown at block 218, additional functions can include sending a timed command to each of the spacecraft, when visible to the ground station 114, indicating to maneuver [i.e., one or more of the series of transfer orbits difference between a respective mean motion of the sequential spacecraft pair] to the respective final orbit 142, and the timed command indicates when to begin the maneuver in order to achieve the synchronized ascent [i.e., in one or more the series of transfer orbits] and to maintain the relative phasing of each spacecraft to each other.”),
wherein the delta phase rate is based on the difference between the respective mean motion (Silva, para. 0082: “At block 202, the method 200 includes releasing the cluster 102 of spacecraft from the launch vehicle 150 at the first orbit 140. Within examples, after releasing the cluster 102 of spacecraft from the launch vehicle 150 at the first orbit 140, relative phasing of each spacecraft to each other is about zero degrees.”); and…
Silva does not appear to disclose the following:
…implementing a compound steering control via a control system, the compound steering control having a plurality of weight factors…wherein applying the respective desired trajectories includes modifying an in-plane change in velocity (Δv) component of each of the multiple spacecraft in relation to an out-of-plane Δv component for each of the multiple spacecraft by manipulating a predetermined one of the plurality of weight factors over a duration of the series of transfer orbits to thereby control a mean motion of each of the multiple spacecraft…causing the multiple spacecraft to maintain the constant phase offset between the sequential spacecraft pair based on applying the delta phase rate to the multiple spacecraft by changing a relative semi-major axis (SMA between the sequential spacecraft pair at a defined rate according to a predetermined relationship, wherein phase offset is provided based on a first SMA of a first spacecraft of the multiple spacecraft that is moved away from a second SMA of a second spacecraft of the multiple spacecraft, wherein the first spacecraft and the second spacecraft are a sequential pair, and wherein the predetermined relationship comprises: changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration; and changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed.
However, in the same field of endeavor, Batla teaches:
…implementing a compound steering control via a control system, the compound steering control having a plurality of weight factors (Batla, col. 11, line 37: “As previously mentioned, enhanced compound steering law [i.e., compound steering control law] offers a shorter mission duration as compared to existing solutions. It also offers more a flexible mission design, i.e. it allows for achieving non-time sensitive orbital targets (e.g., the semi-major axis, the eccentricity, and the inclination) either simultaneously, or at different times, using strategy based weighting factors [i.e., a plurality of weight factors].”)…
…wherein applying the respective desired trajectories includes modifying an in-plane change in velocity (Δv) component of each of the multiple spacecraft in relation to an out-of-plane Δv component for each of the multiple spacecraft by manipulating a predetermined one of the plurality of weight factors over a duration of the series of transfer orbits to thereby control a mean motion of each of the multiple spacecraft (Batla, col. 3, line 63: “In addition, the method involves calculating, with at least one processor, the direction of the vector sum of the velocity change [i.e., modifying an in-plane change in velocity (Δv) component…in relation to an out-of-plane Δv component] required to achieve the orbit eccentricity [i.e., in-plane Δv component], the velocity change required to achieve the semi-major axis [i.e., in-plane Δv component], and the velocity change required to achieve the orbit inclination [i.e., out-of-plane Δv component].”; col. 4, line 64: “In one or more embodiments, at least one processor uses at least one weighting factor for the calculating of the direction of the vector sum [i.e., mean motion of each of the multiple spacecraft].”; col. 11, line 37: “As previously mentioned, enhanced compound steering law offers a shorter mission duration as compared to existing solutions. It also offers more a flexible mission design, i.e. it allows for achieving non-time sensitive orbital targets (e.g., the semi-major axis, the eccentricity, and the inclination) either simultaneously, or at different times, using strategy based weighting factors [i.e., manipulating a predetermined one of the weight factors].”; col. 11, line 7: “In particular, these existing solutions for low thrust transfer strategies [i.e., over a duration of the series of transfer orbits] (also referred to as baseline solutions in this document) involve dividing the transfer into two phases.”)…
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, with the concept of implementing a compound steering law having a plurality of weight factors, for spacecraft changing orbits, and using the weighting factors to modify the changes-in-velocity (or velocity vector) of the in-plane and out-of-plane components, taught by Batla, in order to efficiently (i.e., save time and/or energy, etc.) cause the spacecraft to transfer from one orbit to another (Batla, col. 7, line 7: “One advantage is that, while the existing prior art steering strategies impart the required changes to the orbit semi-major axis and eccentricity in a piece-meal fashion, the disclosed compound steering law can accomplish both objectives (i.e. the orbital eccentricity and the semi-major axis targets) efficiently at the same time. Instead of discontinuous segments and abrupt changes in the spacecraft's steering and configuration, the disclosed compound steering law enables continuous firing of the spacecraft's thrusters and a smooth transition of the spacecraft's orientation. Another advantage is that, except for extreme cases when one of the orbital objectives is predominating, the disclosed compound steering law reduces the total transfer orbit mission duration (TOD) as compared to the existing prior art solutions.”).
Silva and Batla do not appear to teach the following:
…causing the multiple spacecraft to maintain the constant phase offset between the sequential spacecraft pair based on applying the delta phase rate to the multiple spacecraft by changing a relative semi-major axis (SM) between the sequential spacecraft pair at a defined rate according to a predetermined relationship, wherein phase offset is provided based on a first SMA of a first spacecraft of the multiple spacecraft that is moved away from a second SMA of a second spacecraft of the multiple spacecraft, wherein the first spacecraft and the second spacecraft are a sequential pair, and wherein the predetermined relationship comprises: changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration; and changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed.
However, in the same field of endeavor, Bai teaches:
causing the multiple spacecraft to maintain the constant phase offset between the sequential spacecraft pair based on applying the delta phase rate to the multiple spacecraft by changing a relative semi-major axis (SMA) between the sequential spacecraft pair at a defined rate according to a predetermined relationship (translated document of Bai, para. 5: “The relative control strategy realizes the maintenance of the constellation configuration by controlling the relative position between the satellites”; para. 49: “In the solution provided by the embodiment of the present application, before the orbit prediction is performed in the preset forecast time period according to the preset orbit perturbation equation to obtain the orbit elements, it is also necessary to set the basic parameters in advance, where the basic parameters include…set the nominal phase difference between satellites in the same orbital plane [i.e., a predetermined relationship] and the error tolerance of the phase deviation.”; para. 69: “Step 102: Determine whether there is a phase deviation between any adjacent satellites [i.e., causing the multiple spacecraft to maintain the constant phase offset between the sequential spacecraft pair] that exceeds a preset error tolerance.”; para. 70: “Step 103, if it exists, calculate the orbit semi-major axis correction amount corresponding to each satellite in the satellite constellation according to the preset optimization target, and perform orbit control for each satellite according to the orbit semi-major axis correction amount [i.e., changing a relative semi-major axis between the sequential spacecraft pair at a defined rate].”).
wherein phase offset is provided based on a first SMA of a first spacecraft of the multiple spacecraft that is moved away from a second SMA of a second spacecraft of the multiple spacecraft, wherein the first spacecraft and the second spacecraft are a sequential pair (translated document of Bai, para. 5: “The relative control strategy realizes the maintenance of the constellation configuration by controlling the relative position between the satellites”; para. 69: “Step 102: Determine whether there is a phase deviation between any adjacent satellites [i.e., wherein the first spacecraft and the second spacecraft are a sequential pair] that exceeds a preset error tolerance.”; para. 70: “Step 103, if it exists, calculate the orbit semi-major axis correction amount corresponding to each satellite in the satellite constellation according to the preset optimization target, and perform orbit control for each satellite according to the orbit semi-major axis correction amount [i.e., phase offset is provided based on a first SMA of a first spacecraft of the multiple spacecraft that is moved away from a second SMA of a second spacecraft of the multiple spacecraft].”), and…
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, as modified by Batla, with the concept of changing a relative semi-major axis between a sequential spacecraft pair at a defined rate according to a predetermined relationship, taught by Bai, in order to control the spacecraft to their target orbit in the most efficient manner possible while also maintaining the desired constellation configuration (translated document of Bai, para. 8: “The technical problem solved by this application is: in view of the problem that phase control in the prior art is not conducive to extending the on-orbit lifetime of the constellation, this application provides a method and computer equipment for relative phase control of the constellation. In the solution provided in the embodiment of the application, , On the one hand, to minimize the overall fuel consumption of the constellation as the goal, to find the phase correction amount that meets the phase difference constraint between the satellites, to ensure the consistency of the fuel consumption of each satellite to a certain extent, and to ensure the balance of the fuel consumption rate of each satellite It is beneficial to extend the on-orbit life of the constellation from the perspective of fuel consumption; on the other hand, the full information crossover between the various satellites is beneficial to the independent maintenance of the entire constellation configuration and the coordination of mission planning.”).
Silva, Batla, and Bai do not appear to explicitly teach the following:
…wherein the predetermined relationship comprises: changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration; and changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed.
However, in the same field of endeavor, Curtis teaches:
…wherein the predetermined relationship comprises: changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration; and changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed (Curtis, Chapter 6, section 6.5 Phasing maneuvers: "A phasing maneuver is a two-impulse Hohmann transfer from [i.e., changing the first SMA in relation to the second SMA in a first direction and holding the first SMA at a fixed SMA fora calibrated duration] and then back to the same orbit [i.e., changing the first SMA in a second direction, which is opposite to the first direction, until a desired time to achieve the constant phase offset has elapsed], as illustrated in Fig. 6.10. The Hohmann transfer ellipse is the phasing orbit with a period selected to return the spacecraft to the main orbit within a specified time [i.e., a desired time]. Phasing maneuvers are used to change the position of a spacecraft in its orbit.").
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, as modified by Batla and Bai, with the concept of a two-impulse Hohmann transfer, taught by Curtis, in order to complete a phasing maneuver that changes a position of one spacecraft in its orbit, in order to maintain a constant phase offset between two spacecraft (Curtis, Chapter 6, section 6.5: “Phasing maneuvers are used to change the position of a spacecraft in its orbit.”).
Regarding claim 24, and analogous claim 22, Silva, Bai, and Curtis teach the control system of claim 11, but do not appear to teach the following:
wherein applying the respective desired trajectories includes modifying an in-plane change in velocity (Δv) component of each of the multiple spacecraft in relation to an out-of-plane Δv component for each of the multiple spacecraft, and wherein increasing the in-plane Δv component provides a first rate of rising semi-major axis (SMA), while increasing the out-of-plane Δv component provides a second rate of rising SMA, wherein the first rate is faster than the second rate.
However, in the same field of endeavor, Batla teaches:
wherein applying the respective desired trajectories includes modifying an in-plane change in velocity (Δv) component of each of the multiple spacecraft in relation to an out-of-plane Δv component for each of the multiple spacecraft (Batla, col. 3, line 63: “In addition, the method involves calculating, with at least one processor, the direction of the vector sum of the velocity change [i.e., change the in-plane Δv component in relation to the out-of-plane Δv component] required to achieve the orbit eccentricity [i.e., in-plane Δv component], the velocity change required to achieve the semi-major axis [i.e., in-plane Δv component], and the velocity change required to achieve the orbit inclination [i.e., out-of-plane Δv component].”; col. 4, line 64: “In one or more embodiments, at least one processor uses at least one weighting factor for the calculating of the direction of the vector sum [i.e., mean motion of each of the multiple spacecraft].”; col. 11, line 37: “As previously mentioned, enhanced compound steering law offers a shorter mission duration as compared to existing solutions. It also offers more a flexible mission design, i.e. it allows for achieving non-time sensitive orbital targets (e.g., the semi-major axis, the eccentricity, and the inclination) either simultaneously, or at different times, using strategy based weighting factors [i.e., manipulating a predetermined one of the plurality of weight factors].”; col. 11, line 7: “In particular, these existing solutions for low thrust transfer strategies [i.e., over a duration of the series of transfer orbits] (also referred to as baseline solutions in this document) involve dividing the transfer into two phases.”), and
wherein increasing the in-plane Δv component provides a first rate of rising semi-major axis (SMA), while increasing the out-of-plane Δv component provides a second rate of rising SMA, wherein the first rate is faster than the second rate. (Batla, col. 11, line 7: “In particular, these existing solutions for low thrust transfer strategies (also referred to as baseline solutions in this document) involve dividing the transfer into two phases. The first phase involves orbit raising to target the semi-major axis. For this phase, the thrust direction is perpendicular to the radial (i.e. nadir) direction [i.e., increasing the in-plane Δv component provides a first rate of rising semi-major axis (SMA)], while the out-of-plane angle is optimized to remove the majority of the inclination [i.e., while increasing the out-of-plane Δv component provides a second rate of rising SMA]. The second phase involves orbit circularization, or eccentricity reduction. In this phase, the in-plane thrust direction is fixed in the inertial direction, while the out-of-plane angle is optimized to remove the remainder of inclination [i.e., wherein the first rate is faster than the second rate].”; col. 13, line 3: “For these formulas,
F
→
i
n
-
p
l
a
n
e
is a unit vector in a direction of the in-plane thrust direction that will change the current orbit of the spacecraft to the desired orbit for the spacecraft, thereby changing the orbit eccentricity and the semi-major axis simultaneously to achieve the specific orbit eccentricity and the semi-major axis that correspond to the desired orbit for the spacecraft [i.e., in-plane Δv component provides a first rate of rising semi-major axis (SMA)].”).
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention and with a reasonable likelihood of success to modify the invention disclosed by Silva, as modified by Bai and Curtis, with the concept of implementing a compound steering law and using weighting factors to adjust in-plane change-in-velocity components in relation to out-of-plane change-in-velocity components of a spacecraft, taught by Batla, in order to efficiently (i.e., save time and/or energy, etc.) cause the spacecraft to transfer from one orbit to another (Batla, col. 7, line 7: “One advantage is that, while the existing prior art steering strategies impart the required changes to the orbit semi-major axis and eccentricity in a piece-meal fashion, the disclosed compound steering law can accomplish both objectives (i.e. the orbital eccentricity and the semi-major axis targets) efficiently at the same time. Instead of discontinuous segments and abrupt changes in the spacecraft's steering and configuration, the disclosed compound steering law enables continuous firing of the spacecraft's thrusters and a smooth transition of the spacecraft's orientation. Another advantage is that, except for extreme cases when one of the orbital objectives is predominating, the disclosed compound steering law reduces the total transfer orbit mission duration (TOD) as compared to the existing prior art solutions.”).
Conclusion
Applicant's amendment necessitated the new ground(s) of rejection presented in this Office action. Accordingly, THIS ACTION IS MADE FINAL. See MPEP § 706.07(a). Applicant is reminded of the extension of time policy as set forth in 37 CFR 1.136(a).
A shortened statutory period for reply to this final action is set to expire THREE MONTHS from the mailing date of this action. In the event a first reply is filed within TWO MONTHS of the mailing date of this final action and the advisory action is not mailed until after the end of the THREE-MONTH shortened statutory period, then the shortened statutory period will expire on the date the advisory action is mailed, and any nonprovisional extension fee (37 CFR 1.17(a)) pursuant to 37 CFR 1.136(a) will be calculated from the mailing date of the advisory action. In no event, however, will the statutory period for reply expire later than SIX MONTHS from the mailing date of this final action.
Any inquiry concerning this communication or earlier communications from the examiner should be directed to Leah N Miller whose telephone number is (703)756-1933. The examiner can normally be reached M-Th 8:30am - 5:30pm ET.
Examiner interviews are available via telephone, in-person, and video conferencing using a USPTO supplied web-based collaboration tool. To schedule an interview, applicant is encouraged to use the USPTO Automated Interview Request (AIR) at http://www.uspto.gov/interviewpractice.
If attempts to reach the examiner by telephone are unsuccessful, the examiner’s supervisor, Abby Flynn can be reached on (571) 272-9855. The fax phone number for the organization where this application or proceeding is assigned is 571-273-8300.
Information regarding the status of published or unpublished applications may be obtained from Patent Center. Unpublished application information in Patent Center is available to registered users. To file and manage patent submissions in Patent Center, visit: https://patentcenter.uspto.gov. Visit https://www.uspto.gov/patents/apply/patent-center for more information about Patent Center and https://www.uspto.gov/patents/docx for information about filing in DOCX format. For additional questions, contact the Electronic Business Center (EBC) at 866-217-9197 (toll-free). If you would like assistance from a USPTO Customer Service Representative, call 800-786-9199 (IN USA OR CANADA) or 571-272-1000.
/L.N.M./Examiner, Art Unit 3663 /ABBY J FLYNN/Supervisory Patent Examiner, Art Unit 3663