Patent Description:
A spacecraft typically utilizes momentum control actuators such as thrusters and magnetic torque rods to maintain an orbit around a celestial body. When the spacecraft enters safing mode all non-essential systems are shut down. However, essential functions such as thermal management and attitude control are still active. When the spacecraft is in safing mode, momentum control actuators may be used to reduce any external disturbance torques and keep the spacecraft momentum under control.

Some significant disturbance torques that the spacecraft may experience include aerodynamic torque and gravity gradient torque. Aerodynamic torque is created as a spacecraft orbits around a celestial body having an atmosphere. For example, aerodynamic torque is created as a spacecraft moves through the Earth's atmosphere. Furthermore, the aerodynamic torque is stronger at lower altitudes, where the atmosphere on Earth is relatively dense. Gravity gradient torque is created when a difference in gravity exists between some parts of a spacecraft. The difference in gravity may be created when some components or portions of the spacecraft are located closer to a celestial body when compared to some other portions of the spacecraft.

Momentum control actuators may require significant resources. For example, thrusters require fuel to operate. Fuel not only adds mass to the spacecraft, but also increases the overall cost required to operate the spacecraft. Furthermore, magnetic torque rods usually require accurate ephemeris knowledge as well as a magnetic field model or magnetometer. Accordingly, the spacecraft may include a global positioning system (GPS) which potentially adds mass.

<CIT> relates to a sun-referenced safe-hold control system for momentum-biased satellites which employs a safe-hold processor responsive to a set of sun sensors to detect attitude errors of the solar wings.

<CIT> relates to a method and apparatus for maneuvering a satellite in orbit to alternately optimize the collection of solar energy and to take sensor data of terrestrial objects, wherein the longitudinal axis of a large payload package is oriented perpendicular to the orbital plane to minimize the disturbance torque due to gravity gradient, and to allow simple rotation about the axis for attitude change between optimal Sun and optimal ground coverage.

According to an aspect, a spacecraft according to claim <NUM> is disclosed.

In yet another aspect, a method according to claim <NUM> is disclosed.

The features, functions, and advantages that have been discussed may be achieved independently in various examples or may be combined in other examples, further details of which can be seen with reference to the following description and drawings.

The present disclosure is directed towards a control system for a spacecraft that orbits a celestial body. The control system executes an attitude control strategy that reduces disturbance torques experienced by the spacecraft during safing mode. Specifically, the spacecraft reduces both gravity gradient torque and aerodynamic torque experienced by the spacecraft during safing mode, while only requiring partial ephemeris knowledge and an inertial attitude of the spacecraft. The spacecraft may be able to utilize momentum control actuators less frequently when operating in safing mode, which in turn reduces fuel and power consumption. In some instances, the spacecraft may also include smaller, lighter momentum control actuators as well, which also improves fuel and power consumption.

Referring to <FIG>, a schematic diagram of an exemplary spacecraft <NUM> is illustrated. The spacecraft <NUM> includes a control system <NUM> including one or more flight computers <NUM> that are in electronic communication with a plurality of sensing devices <NUM>, one or more actuators <NUM>, one or more power subsystems <NUM>, and one or more antennas <NUM>. A detailed diagram of an exemplary flight computer <NUM> is shown in <FIG> and is described below. The one or more actuators <NUM> include a plurality of internal actuators 28A and a plurality of external actuators 28B. Referring to <FIG>, the spacecraft <NUM> is configured to revolve around a celestial body <NUM> that is surrounded by an atmosphere. In the example as shown, the celestial body <NUM> is the Earth. However, it is to be appreciated that the spacecraft <NUM> may orbit around any celestial body <NUM> having an atmosphere. In another example, the celestial body <NUM> is the planet Mars. When the spacecraft <NUM> orbits the celestial body <NUM>, aerodynamic torque is created. As explained below, the control system <NUM> executes an attitude control strategy that reduces or substantially eliminates the aerodynamic torque, as well as other disturbance torques, exerted upon the spacecraft <NUM>.

Referring to <FIG>, the spacecraft <NUM> revolves in an orbit <NUM> around the celestial body <NUM>. Referring to both <FIG> and <FIG>, the sensing devices <NUM> detect the attitude of the spacecraft <NUM>. The sensing devices <NUM> include, but not limited to, a star tracker, a gyroscope, a magnetometer, a sun sensor, an earth sensor, an accelerometer, and a global positioning system (GPS). The internal actuators 28A are momentum storage devices such as, but not limited to, reaction wheels and control moment gyroscopes. The external actuators 28B may also be referred to as momentum control actuators. The external actuators 28B exert a force upon the spacecraft <NUM> and include actuators such as, but not limited to, thrusters and magnetic torque rods. The thrusters include any type of thruster such as, but not limited to, chemical thrusters, ion thrusters and Hall thrusters. A chemical thruster generates thrust based on a chemical reaction such as, for example, oxidizing a fuel. The power subsystems <NUM> store and provide electrical power to the various components of the spacecraft <NUM> and include devices such as, but not limited to, solar panels, radioisotope thermoelectric generators, batteries, capacitor banks, and heat engines.

<FIG> is an enlarged view of the spacecraft <NUM> shown in <FIG>. The spacecraft <NUM> includes a main body <NUM> that defines an axis that is aligned with a minor principal moment of inertia, which is referred to as a principal axis A-A. The principal axis A-A is positioned substantially perpendicular with respect to a roll axis R and a yaw axis Y of the main body <NUM> of the spacecraft <NUM>. The principal axis A-A is also substantially aligned with a pitch axis P of the main body <NUM> of the spacecraft <NUM>. The spacecraft <NUM> may include two or more solar wings 42A, 42B that project outward from the main body <NUM> of the spacecraft <NUM> that are attached to the main body <NUM> of the spacecraft <NUM>. The solar wings 42A and 42B are substantially aligned with the principal axis A-A of the spacecraft <NUM>, where the upper or north solar wing is designated as solar wing 42A and a lower or south solar wing is designated as solar wing 42B. In the example as shown, a solar axis S-S of the spacecraft <NUM> is substantially aligned with the principal axis A-A. Although <FIG> illustrates solar wings 42A, 42B that derive electrical power from sunlight, it is to be appreciated that other electrical devices for generating power may be used as well.

Referring specifically to <FIG>, the orbit <NUM> around the celestial body <NUM> is shown as an elliptical orbit having a relatively high eccentricity (e.g., where the eccentricity e is about <NUM>). However, it is to be appreciated that this illustration is merely exemplary in nature and other eccentricities may be used as well. For example, the orbit <NUM> may have a relatively low eccentricity of about zero. A vector <NUM> is defined, where the vector <NUM> is substantially normal with respect to the orbit <NUM> that the spacecraft <NUM> follows around the celestial body <NUM>. In the example as shown in <FIG>, the orbit <NUM> around the celestial body <NUM> is an equatorial orbit. That is, the orbit <NUM> is substantially aligned with an equator E of the celestial body <NUM>. However, it is to be appreciated that the orbit <NUM> and the vector <NUM> may be positioned into orientations other than the illustration shown in <FIG>. For example, the celestial body <NUM> may include an inclined orbit instead.

Referring to both <FIG> and <FIG>, the flight computers <NUM> of the spacecraft <NUM> are in wireless communication with a ground control system <NUM> by the antennas <NUM>. The ground control system <NUM> may be located upon the celestial body <NUM> that the spacecraft <NUM> orbits around. For example, the ground control system <NUM> may be located upon the Earth. Alternatively, the ground control system <NUM> may be located on Earth, but the spacecraft <NUM> may be orbiting another celestial body <NUM> that has an atmosphere. The ground control system <NUM> includes one or more computers that send and receive data from the flight computers <NUM> of the spacecraft <NUM>. The ground control system <NUM> may send instructions to the flight computer <NUM>.

The spacecraft <NUM> may enter the safing mode in response to the flight computers <NUM> determining one or more pre-defined spacecraft safing criteria are met. The pre-defined spacecraft safing criteria includes data collected by the sensing devices <NUM> and other on-board data such as, but not limited to, solar wing current, temperature readings of the various components of the spacecraft <NUM>, and stored momentum in a momentum storage device. The spacecraft <NUM> may enter the safing mode when the solar wing current is below a pre-defined current limit and the current conditions indicate the solar wings 42A, 42B should be generating a substantial amount of current. Additionally, some other examples of when the spacecraft <NUM> enters the safing mode include when the temperature of a specific component (or multiple components) of the spacecraft <NUM> (e.g., a payload module) exceeds a pre-defined temperature limit, or when the stored momentum of one or the momentum storage devices exceed a pre-defined momentum limit. Alternatively, the ground control system <NUM> may transmit a signal to the spacecraft <NUM> indicating the spacecraft <NUM> is to enter the safing mode.

Upon entering a safing mode, the control system <NUM> employs an attitude control strategy that is now described. The attitude control strategy includes reducing disturbances torques such as a gravity gradient torque and the aerodynamic torque. Although the disclosure describes reducing the gravity gradient torque and the aerodynamic torque separately, the control system <NUM> employs a single attitude control strategy that addresses both the gravity gradient torque and aerodynamic torque simultaneously. The spacecraft <NUM> is launched into space and revolves around the celestial body <NUM> while following the orbit <NUM>.

In response to entering the safing mode, the flight computer <NUM> executes attitude control. Specifically, during attitude control, the flight computer <NUM> instructs the one or more actuators <NUM> to substantially align the principal axis A-A of the spacecraft <NUM> with the vector <NUM> that is normal to the orbit <NUM> around the celestial body <NUM>. Aligning the principal axis A-A of the spacecraft <NUM> reduces or substantially eliminates gravity gradient torque. Gravity gradient torque is exerted upon the spacecraft <NUM> when a difference in gravitational forces between various sections or portions of the spacecraft <NUM> exists. In the example as shown in <FIG>, the solar wings 42A, 42B are each about equidistant from the celestial body <NUM>. Accordingly, the gravitational force on each solar wing 42A, 42B of the spacecraft <NUM> is about equal, and therefore the gravity gradient torque is reduced or substantially eliminated.

Referring to <FIG> and <FIG>, the flight computers <NUM> instruct the one or more actuators <NUM> to rotate the spacecraft <NUM> about the principal axis A-A at a constant rate, where a rotational orientation of the spacecraft <NUM> relative to the celestial body <NUM> is shifted by about one-half a rotation about the principal axis A-A each time the spacecraft <NUM> completes the orbit <NUM> around the celestial body <NUM>. For example, referring to <FIG> and <FIG>, a first face <NUM> of the main body <NUM> of the spacecraft <NUM> would face the celestial body <NUM> as the spacecraft <NUM> completes a first orbit around the celestial body <NUM>. However, a second face <NUM> of the main body <NUM> of the spacecraft <NUM> that is opposite to the first face <NUM> would face the celestial body <NUM> as the spacecraft completes a second orbit around the celestial body <NUM>, where the second orbit is performed immediately after the first orbit.

It is to be appreciated that shifting the rotational orientation of the spacecraft <NUM> by about one-half a rotation about the principal axis A-A reduces or substantially eliminates disturbance torques exerted upon the spacecraft <NUM>. Specifically, the aerodynamic torque exerted upon the spacecraft <NUM> is reduced or substantially eliminated by shifting the rotational orientation of the spacecraft <NUM>. In addition to the aerodynamic torque, other attitude-dependent torques such as magnetic disturbance torques, are also partially canceled as well.

<FIG> is an exemplary illustration of the spacecraft <NUM>, where the spacecraft <NUM> has just completed a first orbit around the celestial body <NUM> (<FIG>). <FIG> is an illustration of the spacecraft <NUM>, where the spacecraft <NUM> has just completed a second orbit around the celestial body <NUM>. The first orbit and the second orbit are performed consecutively with respect to one another. A center of pressure <NUM> of the spacecraft <NUM> and a center of mass <NUM> of the spacecraft <NUM> are shown in both <FIG>. The center of pressure <NUM> of the spacecraft <NUM> is located in a different position than the center of mass <NUM> of the spacecraft <NUM>.

A net torque T is exerted upon the spacecraft <NUM> because the center of pressure is <NUM> is offset from the center of mass <NUM>. In the example as shown in <FIG>, a first net torque T<NUM> exerted upon the spacecraft <NUM> while completing the first orbit is oriented around a first axis A<NUM>. When the spacecraft <NUM> is completing the second orbit, a second net torque T<NUM> exerted upon the spacecraft <NUM> is oriented around a second axis A<NUM>. A sum of the first net torque T<NUM> and the second net torque T<NUM> is about zero. In other words, the first net torque T<NUM> and the second net torque T<NUM> partially cancel one another, and therefore the total net torque between the first orbit and the second orbit is reduced when compared to a spacecraft that does not rotate through the first and second orbit as described.

Referring back to <FIG>, <FIG>, and <FIG>, the flight computer <NUM> instructs the one or more actuators <NUM> to rotate the spacecraft <NUM> a predetermined number of rotations about the principal axis A-A at a constant rate as the spacecraft <NUM> completes a single orbit around the celestial body <NUM>. The predetermined number of rotations per orbit <NUM> is determined by Equation <NUM>: <MAT> where a value N represents any positive integer including zero. Accordingly, the predetermined number of rotations is at least one-half if the value N is set to zero. Alternatively, if the value N is any whole number, then the spacecraft <NUM> would always rotate an extra one-half rotation about the principal axis A-A while completing a single orbit around the celestial body <NUM>. For example, if the value N is one, then the spacecraft <NUM> would rotate one and a half time about the principal axis A-A while completing a single orbit around the celestial body <NUM>.

The value N is determined based on one or more characteristics of the spacecraft <NUM>. The value N may be determined based on one or more characteristics of the spacecraft that include: thermal characteristics of the spacecraft <NUM> based on proximity to a source of heat, a solar wing angle, a rate limit of the spacecraft <NUM>, a momentum limit of the spacecraft <NUM>, and a structural rate limit of the spacecraft <NUM>. The thermal characteristics of the spacecraft <NUM> refer to heat generated by the sun. For example, sometimes it may not be ideal to heat a particular area of the spacecraft <NUM> above a specific temperature by the sun, and so the spacecraft <NUM> may need to rotate more rapidly around the principal axis A-A. Therefore, the value N may need to increase in order to accommodate the increased rotational speed. The solar wing angle is related to an amount of electrical power that is required by the spacecraft <NUM>. Specifically, the rotational speed that the spacecraft <NUM> rotates at about the principal axis A-A may either increase or decrease depending upon how much sunlight is required to produce the required electrical power. The rate limit of the spacecraft <NUM> represents a maximum speed at which the spacecraft <NUM> may rotate about a given axis. The momentum limit and structural rate limit of the spacecraft <NUM> are based on a momentum management limit of the spacecraft <NUM> and the structural limits of the spacecraft <NUM> respectively. The momentum limit represents a limit on the amount of momentum that the momentum storage devices may store. The structural rate limit represents the maximum rate at which the spacecraft <NUM> may rotate about any given axis without adversely affecting the spacecraft's structure.

Referring to <FIG> and <FIG>, it is to be appreciated that the one or more actuators <NUM> that align the principal axis A-A of the spacecraft <NUM> with the vector <NUM> may be selected from any of the internal actuators 28A and the external actuators 28B. The one or more actuators <NUM> that rotate the spacecraft <NUM> about the principal axis A-A may also be selected from any of the internal actuators 28A and the external actuators 28B. The actuators <NUM> used to align the principal axis A-A of the spacecraft <NUM> with the vector <NUM> may be different than the actuators <NUM> used to rotate the spacecraft <NUM> about the principal axis A-A. In other words, the one or more actuators <NUM> include at least one of the following: a control moment gyroscope, a reaction wheel, thrusters, and magnetic torque rods.

<FIG> is a process flow diagram illustrating an exemplary method <NUM> for executing an attitude control strategy that reduces or substantially eliminates disturbance torques experienced by the spacecraft <NUM>. Referring generally to <FIG>, <FIG>, <FIG>, and <FIG>, the method <NUM> begins at block <NUM>. In block <NUM>, the flight computers <NUM> instruct the spacecraft <NUM> to enter the safing mode. As mentioned above, the flight computers <NUM> may determine that one or more pre-defined spacecraft safing criteria are met. Alternatively, the ground control system <NUM> may transmit a signal to the spacecraft <NUM> directly which causes the spacecraft to enter the safing mode. The method <NUM> may then proceed to block <NUM>.

In block <NUM>, in response to the spacecraft <NUM> entering the safing mode, the flight computers <NUM> instruct the one or more actuators <NUM> to align the principal axis A-A of the spacecraft <NUM> with the vector <NUM> that is normal to the orbit <NUM> around the celestial body <NUM>. The method <NUM> may then proceed to block <NUM>.

In block <NUM>, the flight computers <NUM> instruct the one or more actuators <NUM> to rotate the spacecraft <NUM> about the principal axis A-A, where the rotational orientation of the spacecraft <NUM> relative to the celestial body <NUM> is shifted by about one-half a rotation about the principal axis A-A each time the spacecraft <NUM> completes the orbit <NUM> around the celestial body <NUM>.

Referring generally to the figures, the disclosed attitude control strategy provides various technical effect and benefits by reducing or substantially eliminating disturbance torques upon the spacecraft while also reducing the need for momentum control actuators, such as thrusters and magnetic torque rods, during safing mode. In some instances, the disclosed spacecraft may require smaller momentum control actuators. Accordingly, the mass of the spacecraft is reduced, which results in fuel savings. Many conventional approaches for controlling a spacecraft in safing mode may extensively utilize momentum control devices. It is also to be appreciated some conventional approaches for controlling movement of the spacecraft during safing mode may also require complete ephemeris knowledge of the spacecraft. In contrast, the disclosed attitude control strategy only requires partial ephemeris knowledge (i.e., the orbit normal vector in inertial space) and the inertial attitude to operate.

Referring now to FIG. <NUM>, the flight computer <NUM> and the ground control system <NUM> are implemented on one or more computer devices or systems, such as exemplary computer system <NUM>. The computer system <NUM> includes a processor <NUM>, a memory <NUM>, a mass storage memory device <NUM>, an input/output (I/O) interface <NUM>, and a Human Machine Interface (HMI) <NUM>. The computer system <NUM> is operatively coupled to one or more external resources <NUM> via the network <NUM> or I/O interface <NUM>. External resources may include, but are not limited to, servers, databases, mass storage devices, peripheral devices, cloud-based network services, or any other suitable computer resource that may be used by the computer system <NUM>.

The processor <NUM> includes one or more devices selected from microprocessors, microcontrollers, digital signal processors, microcomputers, central processing units, field programmable gate arrays, programmable logic devices, state machines, logic circuits, analog circuits, digital circuits, or any other devices that manipulate signals (analog or digital) based on operational instructions that are stored in the memory <NUM>. Memory <NUM> includes a single memory device or a plurality of memory devices including, but not limited to, read-only memory (ROM), random access memory (RAM), volatile memory, non-volatile memory, static random-access memory (SRAM), dynamic random-access memory (DRAM), flash memory, cache memory, or any other device capable of storing information. The mass storage memory device <NUM> includes data storage devices such as a hard drive, optical drive, tape drive, volatile or non-volatile solid-state device, or any other device capable of storing information.

The processor <NUM> operates under the control of an operating system <NUM> that resides in memory <NUM>. The operating system <NUM> manages computer resources so that computer program code embodied as one or more computer software applications, such as an application <NUM> residing in memory <NUM>, may have instructions executed by the processor <NUM>. In an alternative example, the processor <NUM> may execute the application <NUM> directly, in which case the operating system <NUM> may be omitted. One or more data structures <NUM> also reside in memory <NUM>, and may be used by the processor <NUM>, operating system <NUM>, or application <NUM> to store or manipulate data.

The I/O interface <NUM> provides a machine interface that operatively couples the processor <NUM> to other devices and systems, such as the network <NUM> or external resource <NUM>. The application <NUM> thereby works cooperatively with the network <NUM> or external resource <NUM> by communicating via the I/O interface <NUM> to provide the various features, functions, applications, processes, or modules comprising examples of the invention. The application <NUM> also includes program code that is executed by one or more external resources <NUM>, or otherwise rely on functions or signals provided by other system or network components external to the computer system <NUM>. Indeed, given the nearly endless hardware and software configurations possible, persons having ordinary skill in the art will understand that examples of the invention may include applications that are located externally to the computer system <NUM>, distributed among multiple computers or other external resources <NUM>, or provided by computing resources (hardware and software) that are provided as a service over the network <NUM>, such as a cloud computing service.

The HMI <NUM> is operatively coupled to the processor <NUM> of computer system <NUM> in a known manner to allow a user to interact directly with the computer system <NUM>. The HMI <NUM> may include video or alphanumeric displays, a touch screen, a speaker, and any other suitable audio and visual indicators capable of providing data to the user. The HMI <NUM> also includes input devices and controls such as an alphanumeric keyboard, a pointing device, keypads, pushbuttons, control knobs, microphones, etc., capable of accepting commands or input from the user and transmitting the entered input to the processor <NUM>.

Claim 1:
A spacecraft (<NUM>) comprising:
a main body (<NUM>) defining a principal axis (A-A);
two or more solar wings (42A, 42B), wherein the two or more solar wings (42A, 42B) are substantially aligned with the principal axis (A-A) of the main body (<NUM>) of the spacecraft (<NUM>); and
a control system (<NUM>) comprising:
one or more actuators (<NUM>);
one or more processors (<NUM>) in electronic communication with the one or more actuators (<NUM>); and
a memory (<NUM>) coupled to the one or more processors (<NUM>), the memory (<NUM>) storing data into a database (<NUM>) and program code that, when executed by the one or more processors (<NUM>), causes the control system (<NUM>) to:
instruct the spacecraft (<NUM>) to enter a safing mode, wherein the spacecraft (<NUM>) revolves in an orbit (<NUM>) around a celestial body (<NUM>) having an atmosphere;
in response to entering the safing mode, instruct the one or more actuators (<NUM>) to substantially align a principal axis (A-A) of the spacecraft (<NUM>) with a vector (<NUM>) that is normal to the orbit (<NUM>) around the celestial body (<NUM>); and
instruct the one or more actuators (<NUM>) to rotate the spacecraft (<NUM>) about the principal axis (A-A), characterised in that a rotational orientation of the spacecraft (<NUM>) relative to the celestial body (<NUM>) is shifted by about one-half a rotation about the principal axis (A-A) each time the spacecraft (<NUM>) completes the orbit (<NUM>) around the celestial body (<NUM>).