Patent Application
20170225805
This disclosure relates to the field of spacecraft, and more particularly, to transfer orbits of a spacecraft.
Satellites or other spacecraft are configured to orbit around the Earth for a variety of purposes, such as communications, exploration, etc. For example, a geosynchronous satellite orbits the Earth and follows the direction of the Earth's rotation. Geosynchronous satellites orbit at a radius of about 42,164 kilometers from the center of the Earth. One revolution of a geosynchronous satellite around the Earth takes about 24 hours, which is the same amount of time it takes for the Earth to rotate once about its axis. These types of satellites are considered geosynchronous because they appear stationary when viewed from a particular location on the Earth, and are commonly used as communication satellites.
To put a geosynchronous satellite into a geosynchronous orbit, the satellite is loaded into a payload of a launch vehicle, and the launch vehicle carries the satellite into space. The launch vehicle may not carry the satellite all the way to the geosynchronous orbit (e.g., 42,164 kilometers), but instead releases the satellite at a lower orbit. The lower orbit may be a few hundred kilometers from Earth. The satellite then performs maneuvers with onboard thrusters to enter a transfer orbit that takes the satellite to the geosynchronous altitude.
Three-axis attitude control may be performed in the transfer orbit. For example, solar power may support the maneuvers of the satellite after separation from the launch vehicle, so the solar panels on the satellite are deployed after separation. High disturbance torques on the satellite makes it difficult to maintain attitude control when passing through low perigees. Therefore, it is desirable to identify new and improved ways for raising a satellite to a higher orbit, before reaching a geosynchronous orbit.
Spinning a spacecraft (e.g., a satellite) in a transfer orbit mitigates issues of attitude control by providing gyroscopic rigidity. The systems and methods described herein stabilize the spinning spacecraft in the transfer orbit, and increase efficiency of thrusters used in the orbital maneuvers. In a powered ascent, low-thrust electric thrusters produce a velocity change (ΔV) in the spacecraft to raise the orbit of the spacecraft. The momentum vector of the spinning spacecraft is parallel to the desired ΔV direction. This attitude of the spacecraft momentum is achieved either at separation from the launch vehicle or through reorientation maneuvers using thrusters. Furthermore, the momentum vector attitude may be maintained or adjusted by the thrusters during orbit maneuvers. A spin controller stabilizes the spacecraft to spin about a target axis in the spacecraft frame. At steady-state, the spin axis aligns with the spacecraft momentum vector.
The electric thrusters are gimbaled, and thrust forces are pointed parallel to the spin axis instead of through the center of mass of the spacecraft. Because the electric thrusters are pointed along the spin axis, the cosine loss for the electric thrusters is minimized for the maneuvers and the electric thrusters are used efficiently. However, when the thrusters are pointed along the spin axis, firing of the thrusters results in a body fixed torque acting on the spacecraft. To compensate for the torque created by firing one or more thrusters, a subsystem (e.g., a reaction wheel subsystem, a control momentum gyroscope (CMG), etc.) produces an angular momentum that couples with the angular velocity of the satellite to create an offsetting torque that cancels the torque from the thruster(s). Therefore, the spin axis of the spacecraft maintains fixed in the inertial frame and parallel to the desired direction of ΔV.
One embodiment comprises a spacecraft comprising a bus having a nadir side and a zenith side opposite the nadir side. The spacecraft further includes a momentum subsystem configured to store angular momentum relative to a center of mass of the spacecraft, and a propulsion subsystem that includes a plurality of electric thrusters installed on the zenith side of the bus to produce a change in velocity (delta-V) on the spacecraft. Each of the electric thrusters is coupled to the bus by a two-axis gimbal assembly. The spacecraft further includes a controller configured to identify a target spin axis for the spacecraft, to determine gimbal angles for at least one of the electric thrusters that so that thrust forces from the at least one electric thruster are parallel to the target spin axis, and to initiate a burn of the at least one electric thruster at the gimbal angles. The controller is configured to control the momentum subsystem to compensate for a thruster torque produced by the burn of the at least one electric thruster at the gimbal angles.
In another embodiment, the momentum subsystem is configured to produce a target angular momentum, where a coupling between the target angular momentum and an angular velocity of the spacecraft creates an offset torque to counteract the thruster torque created from the burn of the at least one electric thruster.
In another embodiment, the controller is configured to estimate an angular velocity of the spacecraft, to estimate the thruster torque created from the burn of the at least one electric thruster, to determine the target angular momentum based on the angular velocity and the thruster torque, and to generate a command to instruct the momentum subsystem to produce the target angular momentum.
In another embodiment, the target spin axis is aligned with a target delta-V direction.
In another embodiment, the momentum subsystem comprises a reaction wheel subsystem having a plurality of reaction wheels.
In another embodiment, the momentum subsystem comprises a control momentum gyroscope (CMG).
In another embodiment, a sensor subsystem that includes an attitude sensor configured to provide measurement data of an attitude of the spacecraft.
In another embodiment, a sensor subsystem that includes a rate sensor configured to provide measurement data of an angular velocity of the spacecraft.
In another embodiment, the plurality of electric thrusters includes a northwest thruster, a northeast thruster, a southwest thruster, and a southeast thruster installed on the zenith side of the bus.
In another embodiment, the electric thrusters use xenon as a propellant.
Another embodiment comprises a method for controlling a spacecraft in a transfer orbit, where the spacecraft comprises a bus having a nadir side and a zenith side opposite the nadir side, a momentum subsystem configured to store angular momentum relative to a center of mass of the spacecraft, and a propulsion subsystem that includes a plurality of electric thrusters installed on the zenith side of the bus by a two-axis gimbal assembly. The method includes identifying a target spin axis for the spacecraft in the transfer orbit, determining gimbal angles for at least one of the electric thrusters that so that thrust forces from the at least one electric thruster are parallel to the target spin axis, initiating a burn of the at least one electric thruster at the gimbal angles, and controlling the momentum subsystem to compensate for a thruster torque produced by the burn of the at least one electric thruster at the gimbal angles.
Another embodiment comprises a controller configured to control a spacecraft in a transfer orbit, where the spacecraft comprises a bus having a nadir side and a zenith side opposite the nadir side, a momentum subsystem that stores angular momentum relative to a center of mass of the spacecraft, and a propulsion subsystem that includes a plurality of electric thrusters installed on the zenith side of the bus. Each of the electric thrusters is coupled to the bus by a two-axis gimbal assembly. The controller is configured to identify a target spin axis for the spacecraft, to determine gimbal angles for at least one of the electric thrusters so that thrust forces from the at least one electric thruster are parallel to the target spin axis, to initiate a burn of the at least one electric thruster at the gimbal angles, and to control the momentum subsystem to compensate for a thruster torque produced by the burn of the at least one electric thruster at the gimbal angles.
The features, functions, and advantages that have been discussed can be achieved independently in various embodiments or may be combined in yet other embodiments, further details of which can be seen with reference to the following description and drawings.
Some embodiments of the present invention are now described, by way of example only, with reference to the accompanying drawings. The same reference number represents the same element or the same type of element on all drawings.
The figures and the following description illustrate specific exemplary embodiments. It will be appreciated that those skilled in the art will be able to devise various arrangements that, although not explicitly described or shown herein, embody the principles described herein and are included within the contemplated scope of the claims that follow this description. Furthermore, any examples described herein are intended to aid in understanding the principles of the disclosure, and are to be construed as being without limitation. As a result, this disclosure is not limited to the specific embodiments or examples described below, but by the claims and their equivalents.
Satellite 100 may comprise a geosynchronous satellite that orbits the Earth and follows the direction of the Earth's rotation. Satellite 100 is initially released into space by a launch vehicle and begins to orbit the Earth at an altitude that is much lower than a geosynchronous altitude. Satellite 100 then performs orbital maneuvers to rise from the initial orbit to the geosynchronous orbit on a “transfer orbit”. A transfer orbit is defined herein as a trajectory by which satellite 100 moves from one orbit to another.
Some satellites may use chemical thrusters in maneuvers for a transfer orbit. A chemical thruster is a type of thruster that burns liquid propellant to produce thrust. One type of chemical thruster is referred to as a bipropellant (or biprop) thruster that burns a liquid fuel and a liquid oxidizer in a combustion chamber. Satellite 100 uses a propulsion subsystem for maneuvers in a transfer orbit that includes only electric thrusters. An electric thruster is a type of “low-thrust” thruster (e.g., produces thrust that is less than 0.1 Newton) that produces electric thrust by accelerating ions. In a typical electric thruster, a propellant (e.g., xenon) is injected into an ionization chamber and ionized by electron bombardment. The ions are then accelerated by an electromagnetic field, and emitted from the thruster as exhaust that produces thrust. One example of an electric thruster is a Xenon Ion Propulsion System (XIPS©) manufactured by L-3 Communications of Aurora, Colo.
Thrusters 314-317 are respectively positioned in northwest, northeast, southwest, and southeast regions of zenith side 105 in this embodiment. Each thruster 314-317 is coupled to bus 102 by a gimbal assembly. For example, northwest thruster 314 is coupled to bus 102 by gimbal assembly 324, northeast thruster 315 is coupled to bus 102 by gimbal assembly 325, southwest thruster 316 is coupled to bus 102 by gimbal assembly 326, and southeast thruster 317 is coupled to bus 102 by gimbal assembly 327. Each gimbal assembly 324-327 is configured to pivot along two axes to alter the thrust vector of its corresponding thruster 314-317. The gimbal angles for each gimbal assembly are referred to herein as p and y. Although not specifically shown in
The number or positions of thrusters 314-317 as shown in
In the embodiment described below, satellite 100 is allowed to spin in at least a portion of the transport orbit (e.g., through low perigees). In some traditional transfer orbits, three-axis attitude control was maintained for the satellite as it moved from one orbit to another. Unfortunately, the satellite is subjected to different forces that make three-axis attitude control difficult when passing through low perigees. In the embodiments described herein, satellite 100 is allowed to spin about a fixed axis of the spacecraft instead of maintaining three-axis attitude control, and is stabilized by active nutation control.
Controller 522 is coupled to propulsion subsystem 310, momentum subsystem 502, and sensor subsystem 510. Controller 522 comprises devices, components, or modules (including hardware, software, or a combination of hardware and software) that process data from sensor subsystem 510, and computes control signals for propulsion subsystem 310 and momentum subsystem 502. Controller 522 may be located on Earth, and able to communicate with subsystems 310, 502, and 510 via wireless signals. Controller 522 may alternatively be located on satellite 100. Controller 522 may alternatively be modularized with a portion of controller 522 located on Earth, and a portion located locally on satellite 100.
It is assumed that satellite 100 has been transported into space with a launch vehicle, has separated from the launch vehicle, and solar panels 108-109 have been deployed (see
Controller 522 identifies a target spin axis for satellite 100 (step 602) in an inertial frame.
For a spinning spacecraft, ideally the thrust vector is co-aligned with the spin axis of the spacecraft to maximize thrust vector efficiency. For example, controller 522 may receive data from attitude sensor 512 indicating the present attitude of satellite 100, and may orient the spacecraft such that the thrust vector in the spacecraft body frame is aligned with the desired target ΔV direction. The target ΔV direction is computed in order to achieve the proper change in orbital elements required to place the spacecraft into the final orbit.
In
Controller 522 initiates a burn of the electric thruster(s) 314-317 at the gimbal angles (step 606 in
Because thrust forces are not through the center of mass 440 of satellite 100 and are pointed substantially along the target ΔV direction, the cosine loss for electric thrusters 314-317 is minimized for the maneuvers. Cosine loss is the portion of thrust forces that do not contribute to ΔV due to thrust forces not pointing straight along the ΔV direction. Because cosine loss is minimized, electric thrusters 314-317 are used efficiently in the maneuvers.
However, when thrust forces are not through the center of mass 440 of satellite 100, the thrust forces will be spaced by a moment arm from the center of mass 440 of satellite 100 (see
Controller 522 controls momentum subsystem 502 to compensate for the thruster torque produced by the burn of the electric thruster(s) 314-317 at the gimbal angles (step 608). For example, controller 522 may include an estimator module 524 (see
Method 600 is performed continuously while satellite 100 is in the transfer orbit to correct the rotation axis of satellite 100 with momentum subsystem 502. Momentum subsystem 502 may also be used to assist in stabilizing the rotation of satellite 100 along the target spin axis 720. Stabilization of a spinning satellite is described in U.S. Patent 6,062,512, which is incorporated by reference.
The following provides an example of spin stabilization for satellite 100.
The following is a dynamic model of satellite 100:
I
+{right arrow over (u)}+{right arrow over (ω)}×(I{right arrow over (ω)}+{right arrow over (h)})={right arrow over (T)}thruster (1)
={right arrow over (u)} (2)
where I is the inertia matrix of satellite 100, {right arrow over (ω)} is the angular velocity (rate) vector of satellite 100 represented in the body coordinate system, {right arrow over (h)} is the momentum from the momentum subsystem 510 in the body coordinate system, {right arrow over (T)}thruster is the body fixed thruster torque, and {right arrow over (u)} is the stabilizing control law in the form of:
{right arrow over (u)}=K(s){right arrow over (ω)} (3)
The target spin axis 720 is the basis of the null space of K(s). At the steady state,
{right arrow over (ω)}×(l{right arrow over (ω)}×{right arrow over (h)})={right arrow over (T)}thruster (4)
The steady state momentum from momentum subsystem 510 includes of two parts:
{right arrow over (h)}={right arrow over (h)}
spinbalancing
+{right arrow over (h′)}
wheel (5)
The angular momentum {right arrow over (h)}spinbalancing balances satellite 100, making the wobble angle zero, as follows:
{right arrow over (ω)}×(l{right arrow over (ω)}+{right arrow over (h)}sb)=0. (6)
The angular momentum {right arrow over (h)}′wheel balances the body fixed thruster torque as follows:
{right arrow over (ω)}×{right arrow over (h)}′wheel={right arrow over (T)}thruster (7)
At steady state, −{right arrow over (ω)}×{right arrow over (h)}′wheel cancels out {right arrow over (T)}thruster.
The spin speed of satellite 100 becomes constant, and its spin axis is fixed in the inertial frame. At the steady state, the body fixed wheel momentum {right arrow over (h)}′wheel rotates about the z-axis in the inertia frame. Therefore, the momentum vector H of satellite 100 forms a cone about the z-axis (target spin axis 720) with a small angle.
The above system and methods provide advantages for orbit maneuvers. At steady-state, the nutation and wobble angles are near zero. The angular momentum from momentum subsystem 502 automatically balances satellite 100 about the target spin axis 720 and also balances the thruster torque in the transverse plane. There is no momentum accumulation in momentum subsystem 502 and satellite 100. At steady state, the spin axis of satellite 100 is fixed in both body frame and inertial frame. If the spin axis is aligned with target ΔV direction 710 in the inertial frame, then the thrust forces are all applied to the target ΔV direction 710. Therefore, the cosine loss of thrust forces is minimized. Minimizing cosine loss can shorten the duration of the transfer orbit, save fuel usage, and prolong the operation life of satellite 100. Because the thrust forces are parallel to the target spin axis 720, it produces near zero torque about the spin axis. Therefore, thruster firing for orbit maneuvers will not cause satellite 100 to spin up or spin down. The total momentum of satellite 100 is near constant, which minimizes the need to adjust the satellite's momentum.
Any of the various elements shown in the figures or described herein may be implemented as hardware, software, firmware, or some combination of these. For example, an element may be implemented as dedicated hardware. Dedicated hardware elements may be referred to as “processors”, “controllers”, or some similar terminology. When provided by a processor, the functions may be provided by a single dedicated processor, by a single shared processor, or by a plurality of individual processors, some of which may be shared. Moreover, explicit use of the term “processor” or “controller” should not be construed to refer exclusively to hardware capable of executing software, and may implicitly include, without limitation, digital signal processor (DSP) hardware, a network processor, application specific integrated circuit (ASIC) or other circuitry, field programmable gate array (FPGA), read only memory (ROM) for storing software, random access memory (RAM), non-volatile storage, logic, or some other physical hardware component or module.
Also, an element may be implemented as instructions executable by a processor or a computer to perform the functions of the element. Some examples of instructions are software, program code, and firmware. The instructions are operational when executed by the processor to direct the processor to perform the functions of the element. The instructions may be stored on storage devices that are readable by the processor. Some examples of the storage devices are digital or solid-state memories, magnetic storage media such as a magnetic disks and magnetic tapes, hard drives, or optically readable digital data storage media.
Although specific embodiments were described herein, the scope is not limited to those specific embodiments. Rather, the scope is defined by the following claims and any equivalents thereof.
