1. Field of the Invention
The present invention relates generally to spinning satellites and more particularly to satellite spin and attitude control.
2. Description of the Related Art
During their operational lifetime, satellites are involved in a variety of maneuvers which require careful control of the satellite's attitude. An exemplary situation is shown in the orbit maneuvers 18 of FIG. 1 that place a satellite 20 into a geostationary orbit (GEO) 24 about the earth 26. In these maneuvers, the satellite 20 is initially launched into a low earth orbit (LEO) 30 by a launcher.
A first velocity increment is then applied (e.g., by a last stage of the launcher) to the satellite 20 to inject it into an elliptical geosynchronous transfer orbit (GTO) 32 which has a perigee 34 tangent to the LEO 30 and an apogee 36 tangent to the GEO 24. At the apogee 36, a second velocity increment injects the satellite 20 into the GEO 24. This final velocity increment is typically applied with a liquid apogee motor (LAM) 38 that is carried by the satellite.
In general, the GTO 32 and the GEO 24 are not coplanar but are inclined from each other by an inclination angle. The inclination angle is essentially a function of the latitude of the earth site from which the satellite 20 was launched (e.g., the GTO 32 and the GEO 24 are typically coplanar only if the launch site was on the earth's equator).
Accordingly, the schematized illustration 40 of FIG. 2 illustrates that the GTO 32 of FIG. 1 is inclined from an equatorial plane 42 by an inclination angle 44. The satellite (20 in FIG. 1) has a velocity at the GTO's apogee (36 in FIG. 1) that is represented by the apogee velocity vector 46 which is in the plane of the GTO 32 and tangent to the GTO. The apogee velocity vector has a magnitude representative of a typical apogee velocity (e.g., .about.1.6 km/sec). In contrast, a GEO velocity vector 48 is in the equatorial plane 42, is tangent to the GEO (24 in FIG. 1) and has a magnitude (.about.3.07 km/sec) that represents the velocity (at the perigee 36) of a satellite in the GEO. For an exemplary GTO which resulted from a launch at an earth latitude of 28 degrees, the inclination angle 44 is .about.28 degrees.
For clarity, an enlarged view 45 of the apogee velocity vector 46 and the GEO velocity vector 48 is shown in FIG. 3. If no plane change were required, a velocity increment equal to the difference between these vectors (.about.1.5 km/sec) would be sufficient for circularization of the satellite's LEO and ejection of the satellite into the GEO (24 in FIG. 1). However, the vector addition of FIG. 3 forms an injection velocity vector 50 with a magnitude of .about.1.8 km/sec and an orientation dictated by the velocity vectors 46 and 48. Accordingly, to inject the satellite 20 into the GEO 24 its LAM 38 must be aligned with the velocity vector 50 and the LAM must generate a force sufficient to achieve a velocity increase of .about.1.8 km/sec in the satellite 20.
This simplified injection process implies that the LAM can deliver an impulse of sufficient magnitude to achieve orbit injection. In actual practice, a lengthy LAM burn time is required and it is generally more fuel efficient to divide the burn time into a plurality of LAM burns. For example, a first LAM burn at the apogee 36 might reduce the ellipticity of the GTO 32 as indicated by the modified elliptical orbit 32A of FIG. 1. This LAM burn could also reduce the inclination angle 44 so that the apogee velocity vector 46 is changed to a modified apogee velocity vector 46A in FIG. 3. Additional LAM burns would then generate elliptical orbits 32B, 32C and so on with the velocity vector changed to apogee velocity vectors 46B, 46C and so on. A final LAM burn would circularize the orbit and reduce the inclination angle 44 to zero.
A summary of conventional GEO injection methods will be facilitated by preceding it with a description of the satellite 20 as envisioned in the present invention and as detailed in FIGS. 4A-4C.
The spatial relationships of elements of the satellite 20 can be defined by reference to an orthogonal body frame 62 that has a longitudinal z axis 63 and a pair of transverse axes in the form of a y axis 64 and an x axis 65. A satellite body 66 carries the LAM 38 on a lower body face 67 with the LAM having a LAM axis 68 that is nominally aligned with the z axis 63. In addition, the body 62 carries a thruster system 70 and a wheel system 72.
Thruster systems of the invention may be arranged to effect a 3 axis control, i.e., they can generate any torque vector that can be defined by the body frame 62. Because the thrusters oppose primarily transverse disturbance torques, the teachings of the invention can be practiced, however, with thruster systems having two degrees of freedom (e.g., two thrusters whose torques are noncolinear with each other and transverse to the LAM axis 68) or having a single degree of freedom (e.g., a single thruster whose torque is transverse to the LAM axis).
An exemplary thruster system 70 is shown to have four axially-oriented thrusters 74 on the lower body face 67. Each is spaced from the z axis 63 and positioned in a different body corner. These thrusters can generate torques about the x and y body axes 64 and 65 (i.e., torque vectors aligned with the axes 64 and 65). In addition, tangentially-oriented thruster pairs 78 and 79 are carried on opposite body faces 80 and 81. These thrusters can generate torques about the z body axis 63 (i.e., torque vectors aligned with the axis 63). Accordingly, the thruster system 70 can effect satellite rotation about any axis of the body frame 62.
Wheel systems of the invention may also be arranged to effect a 3 axis control. Because the wheels also oppose primarily transverse disturbance torques, the teachings of the invention can be practiced with wheel systems having two degrees of freedom (e.g., two noncolinear wheels which are noncolinear with the LAM axis 68) or having a single degree of freedom (e.g., a single wheel which is noncolinear with the LAM axis).
An exemplary wheel system 72 is shown with four wheels 84 which are spaced from the z axis 63, equally spaced circumferentially and angled inward. When the satellite is spinning about the z axis 63, they preferably generate torque about axes 64 and 65 (i.e., torque vectors in x-y plane) by virtue of their angular momentum and a body rate spin. The wheels of the invention can be either momentum wheels or reaction wheels (i.e., they can form either a zero-momentum or a momentum-bias system).
The satellite 20 also includes other conventional satellite elements such as fuel tanks 88 and an electrical power supply system which has batteries 89 and solar panels 90 that carry solar cell arrays 92. The fuel tanks are coupled to supply fuel to the LAM 38 and the thruster system 70. After injection of the satellite 20 into the GEO (24 in FIG. 1), the solar panels 90 would be extended to the deployed position that is indicated in broken lines (for clarity of illustration, the solar panels are shown only in FIGS. 4A and 4C). Although the satellite 20 has been shown to have planar exterior faces, it can have, in general, a variety of configurations (e.g., a cylindrical shape as indicated by the broken circle 96 in the bottom view of FIG. 4C.
Transferring the satellite 20 from the LEO 30 to the GEO 24 of FIG. 1 requires the use of considerable thruster fuel. Accordingly, a substantial portion of the satellite's original mass is lost in this maneuver. In an exemplary situation, a 2700 kg satellite in the LEO 30 may have a mass of only 1800 kg after placement into the GEO 24. Maneuvers which can reduce this mass loss (i.e., reduce fuel use) are of significant importance because the reduction facilitates an increase of satellite life and satellite revenue.
Satellite stabilization during orbit injection has typically been achieved by simply spinning the satellite about the axis of its LAM (the z axis 63 of FIGS. 4A and 4B). If the resultant momentum is sufficiently large, it inhibits the attitude disturbances that are urged by various disturbance torques that are imposed during the orbit injection process. The largest disturbance torque is typically due to spatial offset between the satellite's center of gravity (86 in FIGS. 4A and 4B) and the LAM axis (68 in FIGS. 4A and 4B).
To carry out a useful orbit injection process, therefore, a satellite's spin and attitude control system preferably, (a) spins the satellite about its LAM axis (even if the LAM axis is not an equilibrium spin axis, i.e., a spin axis that is colinear with a spin-induced momentum vector) and (b) aligns the LAM axis with a desired inertial direction (e.g., the velocity vector 50 of FIG. 2) during the LAM's burn time. When the stabilization provided by spin momentum is not sufficiently stiff, these process objectives have often been performed (e.g., see Salvatore, J., et al., Galaxy IIIR Transfer Orbit Attitude Control, G&C Conference, Noordwijk, Netherlands, Fall, 1996) by an all-thruster control system (e.g., exclusive use of the thruster system 70 of FIGS. 4A-4C). Although all-thruster control systems generally have the torque authority to manage processes (a) and (b), such systems are not fuel-efficient.
In another prior art control method (e.g., see Loebel, Michael, Guidance and Control - II, Academic Press, 1964, New York, pp. 313-337), process objective (a) has been performed by an all-wheel control system (e.g., exclusive use of the wheel system 72 of FIGS. 4A-4C). Although the wheels of a wheel system can be appropriately sized to handle LAM-induced torques, this requires them to be larger than is typically needed for efficient management of operational attitude control after orbit injection (i.e., during the service phase of the satellite's lifetime). For example, Loebel employs control moment gyroscopes which are powerful but are larger, more expensive and more complex than fixed wheel systems.
A simple combination of an all-thruster system and an independent all-wheel system is discouraged because corrective action of the thruster system suppresses rotational errors. Thus, the error signals that are typically required to generate corrective action in an all-wheel system are reduced. In addition, the thruster action adds noise to these error signals so that torque and power are wasted as the all-wheel system responds to degraded signals. These problems are also unresolved in a system discussed by Loebel which used wheels and thrusters for above-mentioned process (a) and (b) respectively.
Conventional spin and attitude control systems, therefore, have typically been either fuel-inefficient, have required wheel systems with grossly oversized torque authority (for conventional satellite missions) or have presented system operational problems.
The present invention is directed to accurate, fuel-efficient control structures and methods which have sufficient torque authority to effect satellite maneuvers (e.g., spin a satellite about a desired spin axis and align that spin axis with a desired inertial direction) in the presence of disturbance torques but which can also realize enhanced fuel efficiency.
These objectives are achieved by cooperatively combining a thruster control loop with a wheel control loop in which the wheels respond to an estimate of the disturbance torques (e.g., an angular acceleration estimate, a thruster torque command or a filtered thruster torque command).
In process steps of the thruster control loop, at least one of a desired satellite rotation rate, a desired satellite attitude and a desired angular acceleration is commanded, the commanded parameter is compared with sensed satellite parameters to generate at least one error signal and thruster correction torques are applied to the satellite to reduce the error signal.
In process steps of the wheel control loop, an estimate of the disturbance torques is generated, and, in response to this estimate, wheel correction torques are used to change the internal momentum of the satellite to reduce the magnitude of the thruster correction torques and thereby reduce the use of thruster fuel. The methods of the invention can also be practiced by using received or stored predetermined estimates for the disturbance torque estimate.
Simulations of the invention's methods have shown they have the authority of all-thruster control systems and fuel efficiency that rivals that that of all-wheel control systems.
The novel features of the invention are set forth with particularity in the appended claims. The invention will be best understood from the following description when read in conjunction with the accompanying drawings.