Office Action 1: Notice of Allowance (2025-04-17)

CLAIMS:
CLAIMS

What is claimed is:

1\. A system for atmospheric reentry, comprising:a spacecraft configured to transit an atmosphere;
a plurality of tethers, each tether having:
a proximal end attached to the spacecraft at an attachment point offset from a center of mass of the spacecraft; and a distal end configured to trail behind the spacecraft during atmospheric transit;
wherein the tethers are configured to engage with the atmosphere to passively generate drag and impart a differential tether force on the spacecraft, thereby creating a moment about the center of mass of the spacecraft to passively alter the spacecraft's angle of attack.

2\. The system of claim 1, wherein the distal ends of the tethers are free to interact with atmospheric forces without active control during atmospheric transit.

3\. The system of claim 1, wherein the tethers are configured such that the differential tether force they generate, the resulting moment, and the change in angle of attack induces a skip entry trajectory.

4\. The system of claim 1, wherein at least one of the plurality of tethers is uncontrolled.

5\. The system of claim 1, wherein the attachment point of at least one of the tethers to the spacecraft is offset along at least one axis relative to the spacecraft's center of mass.

6\. The system of claim 1, wherein the attachment points of the tethers to the spacecraft are asymmetrically positioned to create non-uniform aerodynamic forces for modulating a trajectory of the spacecraft during reentry.

7\. The system of claim 1, wherein at least one of the plurality of tethers comprises an aftbody attached at its distal end, the aftbody having aerodynamic properties configured to passively stabilize the tether during atmospheric transit.

8\. The system of claim 7, wherein the aerodynamic properties of the aftbody are configured to increase the drag acting on the tether.

9\. The system of claim 1, wherein the plurality of tethers are configured to passively align the spacecraft in a dynamically preferred orientation during atmospheric reentry without the need for active adjustment mechanisms.

10\. The system of claim 9, wherein the dynamically preferred orientation automatically adjusts to changes in the velocity and trajectory of the spacecraft during atmospheric reentry.

11\. The system of claim 1, wherein the tethers are configured with one or more of predefined lengths, aerodynamic properties, and attachment point locations, to enable the spacecraft to follow a desired trajectory without requiring active control.

12\. The system of claim 1, wherein the differential tether force and the resulting moment and change of the spacecraft's angle of attack are directly related to the spacecraft's velocity such that as the spacecraft's velocity increases or decreases, so too does the corresponding corrective force generated by the tethers and the change in the spacecraft's angle of attack, resulting in cyclic variations in lift and drag forces acting on the system that alternately reduce its forward velocity of the spacecraft, and limit its downward velocity.

13\. A method for atmospheric reentry of a spacecraft, the method comprising:deploying a plurality of tethers from a spacecraft during atmospheric entry;
allowing atmospheric forces to act on the deployed tethers, thereby generating tension forces in each tether and causing a differential tether force, which imparts a moment about the spacecraft's center of mass;
adjusting the spacecraft's angle of attack via said moment to increase an amount of lift acting on the spacecraft during atmospheric reentry.

14\. The method of claim 13, wherein adjusting the spacecraft's angle of attack to increase lift reduces the velocity of the spacecraft and reduces heating and deceleration loads acting on the spacecraft during atmospheric reentry.

15\. The method of claim 13, wherein at least one of the plurality of tethers is uncontrolled.

16\. The method of claim 13, wherein one or more of the lengths and aerodynamic properties of the tethers are configured to optimize the differential tether force, thereby passively controlling the spacecraft's trajectory in a manner configured to correspond to one or more of a desired flight path, a desired acceleration load, and a desired heat load.

17\. The method of claim 13, wherein the lift generated by adjusting the spacecraft's angle of attack induces the spacecraft to engage in a skip entry trajectory.

18\. The method of claim 17, wherein the skip entry trajectory comprises multiple successive skips through the atmosphere.

19\. The method of claim 18, wherein each of the skips reduces a velocity of the spacecraft.

20\. A system for reducing thermal and acceleration loads experienced by a spacecraft during atmospheric entry, comprising:a spacecraft;
a tether-based reaction control system affixed to the spacecraft and comprising:
a plurality of tethers attached to the spacecraft and extending therefrom so as to create a differential tether force via atmospheric drag acting on the tethers during the spacecraft's atmospheric transit;
wherein the differential tether force alters the spacecraft's trajectory by changing its lift-to-drag ratio, thereby reducing acceleration and heat loads on the spacecraft.

21\. The system of claim 20, wherein the tethers are configured to passively interact with the atmosphere during the system's atmospheric transit, enabling continuous adjustment of the spacecraft's lift-to-drag ratio.

22\. The system of claim 20, wherein the tether-based reaction control system is configured to automatically induce a skip-entry trajectory without requiring active control or feedback mechanisms.

23\. The system of claim 20, wherein the tethers are arranged to create an asymmetric distribution of aerodynamic forces to passively adjust the spacecraft's angle of attack during atmospheric entry.

24\. The system of claim 20, wherein the passive interaction between the tethers and the atmosphere eliminates the need for onboard propulsion or powered actuation to achieve trajectory adjustment.

25\. The system of claim 20, wherein the tethers are configured with fixed, predefined aerodynamic characteristics such that trajectory adjustments are entirely driven by natural atmospheric interactions.

26\. The system of claim 20, wherein the reaction control system is configured to generate the differential tether force in response to the atmospheric drag naturally encountered during reentry, without requiring real-time sensing or monitoring systems.

OFFICE ACTION:
ALLOWANCE
Notice of Pre-AIA  or AIA  Status
This action is in response to the application filed 6 March 2025 is a Continuation of application 18/304,812 filed 21 April 2023 which is a Continuation in Part of application 14/469,941, now US Patent 11,724,836 filed 27 August 2014 which is a Continuation in Part of now abandoned application 14/025,822 filed 13 September 2013 which claims priority to PRO 61/801,092 filed 15 March 2013 and is a Continuation in Part of PCT/US2012/028841 filed 12 March 2012. The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .  The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
EXAMINER’S AMENDMENT
The application has been amended as follows:
Claim 13. A method for atmospheric reentry of a spacecraft, the method comprising:
deploying a plurality of tethers from a spacecraft during atmospheric entry;
allowing atmospheric forces to act on the deployed tethers, thereby passively generating tension forces in each tether and passively causing a differential tether force, which imparts a moment about the spacecraft's center of mass;
adjusting the spacecraft's angle of attack via said moment to increase an amount of lift acting on the spacecraft during atmospheric reentry.
Claim 14. The method of claim 13, wherein adjusting the spacecraft’s angle of attack to increase lift reduces the velocity of the spacecraft and reduces one or more of heating and deceleration loads acting on the spacecraft during atmospheric reentry.
Claim 16. The method of claim 13, wherein one or more of the lengths and the aerodynamic properties of the tethers are configured to optimize the differential tether force, thereby nherently controlling the spacecraft's trajectory in a manner configured to correspond to one or more of a desired flight path, a desired acceleration load, and a desired heat load.
Claim 20. A system for reducing thermal and acceleration loads experienced by a spacecraft during atmospheric entry, comprising:
a spacecraft;
a tether-based reaction control system affixed to the spacecraft and comprising:
a plurality of tethers, wherein at least one of the tethers is uncontrolled after deployment, attached to the spacecraft and extending therefrom so as to create a differential tether force via atmospheric drag acting on the tethers during the spacecraft's atmospheric transit;
wherein the differential tether force alters the spacecraft's state vector by changing its angle of attack, thereby reducing acceleration and heat loads on the spacecraft.
Claim 21. The system of claim 20, wherein the tethers are configured to  state vector.
Claim 22. The system of claim 20, wherein the tether-based reaction control system is configured to
Claim 23.  The system of claim 20, wherein the tethers are arranged to create an asymmetric distribution of aerodynamic forces to
Claim 24. The system of claim 20, wherein the passive interaction between the tethers and the atmosphere alters a state vector of the spacecraft, wherein the spacecraft does not have onboard propulsion or powered actuation of the tethers
Claim 25. The system of claim 20, wherein the tethers are configured with fixed, predefined aerodynamic characteristics such that trajectory adjustments are entirely driven by natural atmospheric interactions of a celestial body.
Claim 26. The system of claim 20, wherein the tether-based reaction control system is configured to generate the differential tether force in response to the atmospheric drag
Allowable Subject Matter
Claims 1-26 are allowed.
The following is an examiner’s statement of reasons for allowance:
Regarding Claim 1, the prior art of record fails to disclose or teach “a spacecraft configured to transit an atmosphere; a plurality of tethers, each tether having: a proximal end attached to the spacecraft at an attachment point offset from a center of mass of the spacecraft; and a distal end configured to trail behind the spacecraft during atmospheric transit; wherein the tethers are configured to engage with the atmosphere to passively generate drag and impart a differential tether force on the spacecraft, thereby creating a moment about the center of mass of the spacecraft to passively alter the spacecraft's angle of attack” in combination with the rest of the limitations in the claim.  This limitation, in the apparatus as claimed in claim 1 is neither anticipated nor made obvious by the prior art of record.  Claims 2-12 depend from claim 1 and are therefore also found allowable.
Regarding Claim 13, the prior art of record fails to disclose or teach “deploying a plurality of tethers from a spacecraft during atmospheric entry; allowing atmospheric forces to act on the deployed tethers, thereby passively generating tension forces in each tether and passively causing a differential tether force, which imparts a moment about the spacecraft's center of mass; adjusting the spacecraft's angle of attack via said moment to increase an amount of lift acting on the spacecraft during atmospheric reentry” in combination with the rest of the limitations in the claim.  This limitation, in the method as claimed in claim 13 is neither anticipated nor made obvious by the prior art of record.  Claims 14-19 depend from claim 13 and are therefore also found allowable.
Regarding Claim 20, the prior art of record fails to disclose or teach “a spacecraft; a tether-based reaction control system affixed to the spacecraft and comprising: a plurality of tethers, wherein at least one of the tethers is uncontrolled after deployment, attached to the spacecraft and extending therefrom so as to create a differential tether force via atmospheric drag acting on the tethers during the spacecraft's atmospheric transit; wherein the differential tether force alters the spacecraft's state vector by changing its angle of attack, thereby reducing acceleration and heat loads on the spacecraft” in combination with the rest of the limitations in the claim.  This limitation, in the apparatus as claimed in claim 20 is neither anticipated nor made obvious by the prior art of record.  Claims 21-26 depend from claim 20 and are therefore also found allowable.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Conclusion



CITED REFERENCE SUMMARIES: Ref US3582016A: Claims:
In combination: a. a satellite main body adapted for projection into an orbit around the earth, b. an auxiliary body, c. means for effecting deployment of said auxiliary body in space substantially vertically at a distance from said main body which is more than 10 times the length of any rigid part of said main body, d. a flexible tether adapted for maintaining interconnection of said bodies so deployed and for maintaining said distance by tension therein, e. facilities for storing said tether compactly for launching into orbit, said tether being withdrawable from said facilities during said deployment, f. means for measuring the state of oscillations tilting the tetherline from the local vertical. g. a thruster in said main body positioned to provide thrust directed substantially horizontally and through the center of mass of said main body, said thrust being controllable in response to the output of said means for measuring so as to reduce the amplitude of said oscillations of the tetherline, and h. means for damping rotational oscillations of said main body.
The combination as defined by claim 1 wherein said tether is substantially without bending stiffness, light in weight, and more than 30 times the length of any rigid part of said main body.
The combination as defined by claim 1 wherein one of said bodies comprises a first portion, a second portion, and a damping hinge interconnecting said first and second portions.
The combination as defined by claim 1 wherein said main body includes earth-pointing instrumentation.
The combination as defined by claim 4 wherein said instrumentation comprises a directional antenna.
The combination as defined by claim 1 wherein said last-named means comprises a single-degree-of-freedom gyroscope.
The combination as defined by claim 1 wherein said main body comprises a structure attached to said tether and extensible in orbit, thereby to provide a point of attachment of said tether to said main body relatively distant from the center-of-mass of said main body.
The combination as defined by claim 1 wherein said auxiliary body is releasably attachable to said main body for transport into orbit and said deployment means comprises means for thrusting said auxiliary body away from said main body consequent upon its release therefrom.
The combination as defined by claim 8 wherein said auxiliary body is the last stage of the launch vehicle which places said main body into orbit.
The combination as defined by claim 1 wherein said means for measuring is an earth-horizon sensor.
The combination as defined by claim 12 wherein said last-named means comprises a single-degree-of-freedom gyroscope.
The combination as defined by claim 1 wherein said means for measuring comprises ground-based means for measuring the strength of radio signals received from a directional radio antenna mounted on said main body.
The combination as defined by claim 1 wherein the thrust of said thruster is controllable by radio command from the ground consequent upon grOund-based analysis of the output of said means for measuring.
The combination as defined by claim 2 wherein the length of said tether is more than 100 times the greatest dimension of any rigid part of said main body.
The combination as defined by claim 2 wherein the length of said tether is more then 500 times the greatest dimension of any rigid part of said main body.
The combination as defined by claim 2 wherein one of said bodies comprises a first portion, a second portion, and a damping hinge interconnecting said first and second portions.
The combination as defined by claim 7 wherein said last-named means comprises a single-degree-of-freedom gyroscope.
The combination as defined by claim 8 wherein said last-named damping means comprises a single-degree-of-freedom gyroscope.
The combination as defined by claim 10 wherein said means for damping comprises a single-degree-of-freedom gyroscope.
The combination as defined by claim 11 wherein said last-named damping means comprises a single-degree-of-freedom gyroscope.
The combination as defined by claim 8 wherein said main body comprises a structure attached to said tether and extensible in orbit, thereby to provide a point of attachment of said tether to said main body relatively distant from the center-of-mass of said main body.
The combination as defined by claim 10 wherein said main body comprises a structure attached to said tether and extensible in orbit, thereby to provide a point of attachment of said tether to said main body relatively distant from the center-of-mass of said main body.
The combination as defined by claim 11 wherein said main body comprises a structure attached to said tether and extensible in orbit, thereby to provide a point of attachment of said tether to said main body relatively distant from the center-of-mass of said main body.
The combination as defined by claim 12 wherein said main body comprises a structure attached to said tether and extensible in orbit, thereby to provide a point of attachment of said tether to said main body relatively distant from the center-of-mass of said main body. 25. The combination as defined by claim 12 wherein said auxiliary body is releasably attached to said main body for transport into orbit and said deployment means comprises means for thrusting said auxiliary body away from said main body consequent upon its release therefrom.
The combination as defined by claim 3 wherein said auxiliary body is the last stage of the launch vehicle which places said main body into orbit.
The combination as defined by claim 6 wherein said auxiliary body is the last stage of the launch vehicle which places said main body into orbit.
The combination as defined by claim 7 wherein said auxiliary body is the last stage of the launch vehicle which places said main body into orbit.
The combination as defined by claim 10 wherein said auxiliary body is the last stage of the launch vehicle which places said main body into orbit.
The combination as defined by claim 11 wherein said auxiliary body is the last stage of the launch vehicle which places said main body into orbit.
The combination as defined by claim 12 wherein said auxiliary body is the last stage of the launch vehicle which places said main body into orbit.
The combination as defined by claim 12 wherein said means for measuring is an earth-horizon sensor.
The combination as defined by claim 12 wherein said means for measuring comprises ground-based means for measuring the strength of radio signals received from a directional radio antenna mounted on said main body.
In combination: a. a satellite main body adapted for projection into an orbit around the earth, b. an auxiliary body, c. a flexible tether, d. facilities fOr containing said tether compactly for launching, e. means for effecting deployment of said auxiliary body in space substantially vertically at a distance from said main body which is more than 10 times the length of any rigid part of said main body, for withdrawing said tether from said facilities during said deployment, and for maintaining said distance by tension in said tether, f. means for measuring the state of oscillations tilting the tetherline from the local vertical, g. a thruster in said main body positioned to provide thrust directed substantially horizontally and through the center of mass of said main body, said thrust being controllable in response to the output of said means for measuring so as to reduce the amplitude of said oscillations of the tetherline, and h. means for damping rotational oscillations of said main body.
In combination: a. a satellite main body adapted for projection into an orbit around the earth, b. an auxiliary body, c. a flexible tether adapted for maintaining taut interconnection of said bodies substantially along the local vertical, the length of said tether being at least 10 times the greatest dimension of any rigid part of said main body, d. means for effecting deployment of said auxiliary body to the required position relative to said main body after orbital injection, e. facilities for storing said tether completely for launching into orbit, said tether being withdrawable from said facilities during said deployment, f. means for measuring the state of oscillations tilting the tetherline from the local vertical, g. a thruster in said main body positioned to provide thrust directed substantially horizontally and through the center of mass of said main body, said thrust being controllable in response to the output of said means for measuring so as to reduce the amplitude of said oscillations of the tetherline, and h. means for damping rotational oscillations of said main body.
For use with a satellite main body adapted for projection into an orbit around the earth, a system for maintaining said body in a preferred earth orientation comprising; a. an auxiliary body, b. a tether, c. facilities for containing said tether for launching, d. means for deploying said auxiliary body in space at a distance from said main body which is more than 20 times the length of any rigid part of said main body, connected to said main body by said tether, and for maintaining said distance by tension in said tether, e. reaction means, part of said main body, positioned to provide thrust directed substantially horizontally and through the center of mass of said main body, f. means for measuring the state of dumbbell oscillations of said tetherline, g. means for controlling said thrust responsive to the state of said dumbbell oscillations, and h. means for damping rotational oscillations of said main body.
The combination as defined by claim 36 wherein said main body contains for yaw stabilization, a rotor constrained with its spin axis oriented so as to produce a substantial quantity of angular momentum directed substantially in that plane of the main body that is to be stabilized vertically and perpendicular to the orbital plane.
The combination as defined by claim 37 wherein said rotor is the rotor element of a gyroscope and said gyroscope comprises viscous damping means applied to its gimbal axis motion.
Ref US5310144A: Claims:
A satellite for orbiting the earth in an orbit plane about the earth with a sun based inertial reference system centered at said satellite having mutually perpendicular overturning, windmill and pitch axes, with the windmill axis being along the projection of the sun line in the orbit plane, said overturning axis being in the orbit plane and normal to the windmill axis, and said pitch axis extending perpendicular to said orbit plane, said satellite comprising: a main body; at least one solar panel extending from said main body out of the orbit plane in a deployed position with said panel facing the sun; said panel having proximal and distal ends relative to said main body and said panel in said deployed position being oriented at an angle relative to a line normal to said orbit plane; and tilting means for rotating said panel about a tilt axis parallel to the overturning axis and adjacent said proximal end for variably orienting said panel relative to aid normal line in said deployed position whereby said panel is tilted to move said distal end toward or away from the sun and solar pressure and gravity gradient forces on said panel can be varied in order to minimize overturning torque acting on said satellite about said overturning axis.
at least one solar panel extending from said main body out of the orbit plane in a deployed position with said panel facing the sun; said panel having proximal and distal ends relative to said main body and said panel in said deployed position being oriented at an angle relative to a line normal to said orbit plane; and
tilting means for rotating said panel about a tilt axis parallel to the overturning axis and adjacent said proximal end for variably orienting said panel relative to aid normal line in said deployed position whereby said panel is tilted to move said distal end toward or away from the sun and solar pressure and gravity gradient forces on said panel can be varied in order to minimize overturning torque acting on said satellite about said overturning axis.
The satellite of claim 1 wherein said tilting means includes: a panel mounting bracket coupled to said body at one end for rotation about said tilt axis and means for mounting said panel at the opposite end of said bracket; a support link having two ends and coupled at one end to said body for rotation about a second axis parallel to said tilt axis and spaced therefrom, coupling means for attaching the opposite end of said link to said bracket at an attachment point spaced from both said tilt axis and said second axis thereby forming a three bar linkage; and actuator means for varying the distance between said attachment point and said tilt axis whereby said bracket is rotated about said tilt axis and the orientation of said panel relative to said normal line is varied.
a panel mounting bracket coupled to said body at one end for rotation about said tilt axis and means for mounting said panel at the opposite end of said bracket;
a support link having two ends and coupled at one end to said body for rotation about a second axis parallel to said tilt axis and spaced therefrom, coupling means for attaching the opposite end of said link to said bracket at an attachment point spaced from both said tilt axis and said second axis thereby forming a three bar linkage; and
actuator means for varying the distance between said attachment point and said tilt axis whereby said bracket is rotated about said tilt axis and the orientation of said panel relative to said normal line is varied.
The satellite of claim 2 wherein said actuator means includes linear drive means for moving said attachment point relative to said tilt axis.
The satellite of claim 3 wherein said linear drive means includes a jack screw carried by said bracket for rotation about the longitudinal axis of said screw, said screw axis being oriented normal to said first and second axes, said link opposite end being threadably engaged with said jack screw whereby said link opposite end is moved longitudinally along said jack screw upon rotation of said jack screw, and motor means for rotating said jack screw.
The satellite of claim 1 further comprising: sensing means for measuring the attitude change of said satellite in response to environmental disturbance torques acting on said satellite; processing means, coupled to said sensing means, for determining the desired angle of said panel relative to said normal line to minimize the overturning torque acting on said satellite; and said processing means being coupled to said tilting means for activating said tilting means to rotate said panel to vary the orientation of said panel relative to said normal line to said desired angle.
sensing means for measuring the attitude change of said satellite in response to environmental disturbance torques acting on said satellite;
processing means, coupled to said sensing means, for determining the desired angle of said panel relative to said normal line to minimize the overturning torque acting on said satellite; and
said processing means being coupled to said tilting means for activating said tilting means to rotate said panel to vary the orientation of said panel relative to said normal line to said desired angle.
A satellite for orbiting the earth in an orbit plane about the earth with a sun based inertial reference system centered at said satellite having mutually perpendicular overturning, windmill and pitch axes, with the windmill axis being along the projection of the sun line in the orbit plane, said overturning axis being in the orbit plane and normal to the windmill axis, and said pitch axis extending perpendicular to said orbit plane, said satellite comprising: a main body; a solar panel extending from said main body out of the orbit plane in a deployed position with said panel generally facing the sun; said panel having proximal and distal ends relative to said main body and said panel in said deployed position being oriented at an angle relative to a line normal to said orbit plane; and means for mounting said panel to said main body including sun tracking means for rotating said panel relative to said body about a rotation axis parallel to said pitch axis to track the sun and tilting means for rotating said panel about a tilt axis parallel to the overturning axis and adjacent said panel proximal end to variably orient said panel relative to said normal line whereby said panel is tilted to move said distal end toward or away from the sun and the overturning torque acting on said satellite can be minimized by varying the solar pressure and gravity gradient forces acting on said panel.
a solar panel extending from said main body out of the orbit plane in a deployed position with said panel generally facing the sun; said panel having proximal and distal ends relative to said main body and said panel in said deployed position being oriented at an angle relative to a line normal to said orbit plane; and
means for mounting said panel to said main body including sun tracking means for rotating said panel relative to said body about a rotation axis parallel to said pitch axis to track the sun and tilting means for rotating said panel about a tilt axis parallel to the overturning axis and adjacent said panel proximal end to variably orient said panel relative to said normal line whereby said panel is tilted to move said distal end toward or away from the sun and the overturning torque acting on said satellite can be minimized by varying the solar pressure and gravity gradient forces acting on said panel.
The satellite of claim 6 further comprising: sensing means for measuring the attitude change of said satellite in response to environmental disturbance torques acting on said satellite; processing means, coupled to said sensing means, for determining the desired angle of said panel relative to said normal line to minimize the overturning torque acting on said satellite; and said processing means being coupled to said tilting means for activating said tilting means to rotate said panel to vary the orientation of said panel relative to said normal line to said desired angles.
said processing means being coupled to said tilting means for activating said tilting means to rotate said panel to vary the orientation of said panel relative to said normal line to said desired angles.
The satellite of claim 6 wherein said mounting means includes a mounting bracket carried by said sun tracking means for rotation with said panel for tracking the sun and attached to said sun tracking means at one end of said bracket with said solar panel being attached to the opposite end of said bracket; said bracket having a use position extending from said main body at an angle relative to said normal line; and said tilting means including a pivotal attachment of said bracket one end to said sun tracking means for rotation about said tilt axis to variably orient said bracket and said panel relative to said normal line when said bracket is in said use position.
said bracket having a use position extending from said main body at an angle relative to said normal line; and
said tilting means including a pivotal attachment of said bracket one end to said sun tracking means for rotation about said tilt axis to variably orient said bracket and said panel relative to said normal line when said bracket is in said use position.
The satellite of claim 8 further comprising: deployment means for rotating said bracket after satellite launch about said tilt axis to said use position; and latch means for holding said bracket in said use position and fixed from rotation by said deployment means.
deployment means for rotating said bracket after satellite launch about said tilt axis to said use position; and
latch means for ho
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