Space satellite with agile payload orientation system

An agile payload orientation system for a space satellite comprises a three-axis gimbal (30) for coupling an appended body (20) containing the payload (24, 25, 26, 27) to a main body (10) of the satellite. The appended body (20) also contains moment control gyroscopes (29), which are responsive to external command for changing the orientation of the appended body (20) relative to an inertial reference. The gimbal (30) comprises three separate torquer-resolver units (31, 32, 33), which are connected to each other in series so that their axes are mutually orthogonal to each other and coincide at a single point (34) located at the center of mass of the appended body (20). Since the center of mass of the appended body (20) does not move when the appended body (20) is rotated about the axis of any of the torquer-resolver units (31, 32, 33), a change in orientation of the appended body (20) can be made without disturbing the orientation of the main body (10).

BACKGROUND OF THE INVENTION 
1. Field of the Invention 
This invention pertains generally to space satellites, and particularly to 
payload orientation systems for space satellites. 
2. Description of Prior Art 
Proposed space satellite missions require a payload orientation system for 
maneuvering the payload of the satellite rapidly on command without 
disturbing the orientation of the main body of the satellite. Agility of 
the satellite payload is especially important in tracking fast-moving 
targets or for shifting surveillance rapidly from one target to another. 
It has been proposed to support the payload of a satellite on a separate 
body appended to the main body of the satellite by an articulating joint, 
so that the orientation of the payload can be changed without having to 
maneuver the main body. However, until the present invention, no technique 
had been devised for changing the orientation of an appended body 
containing the satellite payload without concomitantly transferring 
momentum to the main body of the satellite. 
SUMMARY OF THE INVENTION 
It is an object of this invention to provide a space satellite having a 
payload contained in a separate body appended to the main body of the 
satellite, where the orientation of the appended body can be changed 
without causing significant transfer of momentum to the main body. 
It is likewise an object of this invention to provide an agile payload 
orientation system for a space satellite, whereby the orientation of the 
satellite payload can be changed without causing a change in orientation 
of the main body of the satellite. 
It is a particular object of this invention to provide a space satellite 
having an agile payload contained in a separate body appended to the main 
body of the satellite by an articulating joint, where the orientation of 
the payload can be changed on command by transferring momentum internally 
within the appended body without causing any significant transfer of 
momentum to the main body of the satellite. 
In a space satellite according to the present invention, the payload is 
contained in a separate body that is appended to the main body of the 
satellite by a gimbal, which allows the orientation of the appended body 
to be changed independently of the orientation of the main body. The 
gimbal of the present invention has bearings that permit rotation of the 
appended body relative to the main body about three orthogonal axes that 
intersect substantially at the center of mass of the appended body. A 
momentum exchange means comprising three or more gyroscopic devices 
mounted on the appended body permits the appended body to undergo selected 
rotational motions on command about the three orthogonal gimbal axes, 
thereby changing the orientation of the appended body without affecting 
the orientation of the main body. 
In the preferred embodiment of the present invention, a means is provided 
to nullify the effect of any torques applied to the main body of the 
satellite by changes in orientation of the appended body due to friction 
at the bearings of the gimbal connecting the appended body to the main 
body. In particular, an attitude sensor is mounted on the main body to 
detect incipient motions of the main body relative to an inertial 
reference that are caused by torques due to frictional forces at the 
gimbal bearings. The attitude sensor generates electronic signals 
indicative of the frictional forces at the gimbal bearings; and torquing 
motors responsive to these signals apply equal and opposite countertorques 
to the main body to nullify the effect of the frictional forces. In this 
way, the main body of the satellite is effectively isolated from motions 
of the appended body. The orientation of the satellite payload can 
therefore be changed on command without disturbing the orientation of the 
satellite main body.

DESCRIPTION OF THE PREFERRED EMBODIMENT 
As illustrated in FIGS. 1 and 2, a space satellite according to the present 
invention comprises a main body 10 and an appended body 20, which are 
coupled together by a three-axis gimbal 30. The main body 10 typically 
houses a propulsion motor, and also houses electrical systems for 
regulation and distribution of power to components of both the main body 
10 and the appended body 20. In the usual case, the main body 10 is of 
generally cylindrical configuration and is more massive than the appended 
body 20. 
The main body 10 preferrably includes a star tracker 11 for monitoring the 
attitude and position of the satellite relative to an inertial reference 
such as the star Polaris. The main body 10 might also include a 
communications antenna 12 for transmitting information to and receiving 
commands from earth, either directly or via a communications satellite 13. 
The main body 10 might further include foldable solar arrays 14 and 15 
mounted on extendible mast structures that can be deployed on command when 
the satellite achieves a stable orbit. One end 16 of the main body 10 is 
coupled by conventional means, such as a mating collar 17, to one end of 
the gimbal 30. The other end of the gimbal 30 is secured to the appended 
body 20. 
The appended body 20 comprises appropriate structural members for 
supporting the satellite payload. As depicted in FIGS. 1 and 2, the 
appended body 20 could comprise a symmetrically configured frame having 
four side members 21 joined together to form a square structure, with four 
leg members 22 of substantially equal length projecting one from each 
corner of the square structure. The leg members 22 extend to a generally 
circular rim member 23 to which they are attached. The plane of the rim 
member 23 is spaced apart from and generally parallel to the plane of the 
square structure formed by the side members 21. 
Attached to the rim member 23 could be several instruments, as indicated in 
the drawing by reference numbers 24, 25, 26 and 27, which constitute the 
principal components of the satellite payload. The nature of the payload 
instruments depends on the nature of the particular satellite mission. 
Attached to the square structure formed by the side members 21 are two 
support members 28 on which conventional moment control gyroscopes 29 are 
mounted. The gyroscope support members 28 are symmetrically postioned with 
respect to each other on the square structure. The gyroscopes 29 are 
responsive to external command by conventional electronic means for 
changing the orientation of the appended body 20 relative to the inertial 
reference. 
The gimbal 30 is a three-axis gimbal comprising three separate 
torquer-resolver units 31, 32 and 33, which are connected to each other in 
series so that their axes are mutually orthogonal to each other and 
coincide at substantially a single point 34, as shown in detail in FIG. 3. 
The torquer-resolver units 31, 32 and 33 may be conventional, each 
comprising a motor and an angular measuring device such as an optical 
encoder disposed within a housing structure. 
Motor shaft 35 of the torquer-resolver unit 31 coincides with the axis of 
the torquer-resolver unit 31, which coincides with the longitudinal (i.e., 
cylindrical) axis of the main body 10. The motor shaft 35 is attached to a 
mating pad 36, which is secured to or integral with the housing structure 
of the torquer-resolver unit 32. Thus, rotation of the motor shaft 35 of 
the torquer-resolver unit 31 causes yaw of the appended body 20 about the 
longitudinal axis of the main body 10 of the satellite. 
Motor shaft 37 of the torquer-resolver unit 32 coincides with the axis of 
the torquer-resolver unit 32, and is orthogonal to the axis of the 
torquer-resolver unit 31. Two straps 38 (only one of which is seen in the 
drawing) are attached one at each end of the motor shaft 37, and extend to 
connections with the housing structure of the torquer-resolver unit 33. 
Thus, rotation of the motor shaft 37 of the torquer-resolver unit 32 
causes pitch of the appended body 20 relative to the main body 10 about 
the axis of the torquer-resolver unit 32. 
The torquer-resolver unit 33 is supported by the straps 38 so that its axis 
is orthogonal to the axes of the torquer-resolver units 31 and 32, and so 
that the axes of the torquer-resolver units 31, 32 and 33 all intersect at 
the common point 34. Motor shaft 39 of the torquer-resolver unit 33 
coincides with the axis of the torquer-resolver unit 33, and is connected 
by a journalled attachment member 40 to one of the sides of the square 
structure formed by the side members 21 of the appended body 20. Thus, 
rotation of the motor shaft 39 of the torquer-resolver unit 33 causes roll 
of the appended body 20 relative to the main body 10 about the axis of the 
torquer-resolver unit 33. 
The positioning of the control moment gyroscopes 29 on their respective 
support structures 28, and of the payload instruments 24, 25, 26 and 27 on 
the rim member 23, is such as to locate the center of mass of the appended 
body 20 substantially at the common point of intersection 34 of the three 
axes of the gimbal 30. Ordinarily, the various payload instruments 24, 25, 
26 and 27 would be configured and positioned symmetrically with respect to 
each other around the rim member 23. Each moment control gyroscope 29 
comprises a motor, a gimbal mounted for rotation by the gyroscope motor 
about a gimbal axis, and a momentum wheel attached to the gimbal for 
rotation about a spin axis orthogonal to the gimbal axis. The gimbal of 
each gyroscope 29 is responsive to external command for rotating the 
attached momentum wheel about the gimbal axis. A set of appropriate 
commands to one or more of the gyroscopes 29 thereby permits the appended 
body 20 to be rotated about a selected one or more of the axes of the 
gimbal 30. 
Since the center of mass of the appended body 20 is located at the point of 
intersection 34 of the axes of the gimbal 30, the center of mass of the 
appended body 20 does not move when the appended body 20 is rotated about 
the gimbal axes. Neglecting the effect of frictional forces at the 
bearings of the gimbal 30, a change in the orientation of the appended 
body 20 does not cause any torque to be applied to the main body 10. Thus, 
the orientation of the payload supported on the appended body 20 can be 
changed without transferring momentum to the main body 10 of the 
satellite. The gimbal 30 effectively decouples the momentum of the main 
body 10 from the momentum of the appended body 20. 
Frictional forces at the bearings of the gimbal 30 cause torques to be 
applied to the main body 10 when the appended body 20 changes orientation. 
If the main body 10 is sufficiently more massive than the appended body 
20, the effect of friction at the bearings of the gimbal 30 may be 
negligible. However, when the masses of the main body 10 and the appended 
body 20 do not differ too greatly, the effect of friction at the gimbal 
bearings may become significant. Usually, also, one or more electrical 
cables (not shown in the drawing) would extend from the main body 10 to 
the appended body 20 to carry electrical power and signals to the appended 
body 20. Depending upon the relative masses of the main body 10 and the 
appended body 20, stiffness in the electrical cable or cables might cause 
non-negligible torques to be applied to the main body 10 when the appended 
body 20 changes orientation. 
An attitude sensor system 50, as shown schematically in FIG. 4 can be used 
to apply countertorques to the main body 10 in order to nullify the effect 
of any torques applied to the main body 10 due to frictional forces at the 
bearings of the gimbal 30 and/or stiffness of electrical cables extending 
from the main body 10 to the appended body 20. 
The attitude sensor system 50 comprises a computer 51 for receiving a 
continuous input from the star tracker 11, which tracks two or more stars 
to enable the position and attitude of the main body 10 to be continuously 
monitored. An instrument gyro 52 capable of measuring three-axis inertial 
rotation rates, is also a component of the attitude sensor system 50. The 
gyro 52 senses deviations of the attitude of the main body 10 from a 
desired attitude, and provides input signals to the computer 51 indicative 
of such deviations. The computer 51 processes these input signals to 
provide output signals to the motors of the torquer-resolver units 31, 32 
and 33 of the gimbal 30, which cause countertorques to be applied to the 
main body 10 to nullify the effect of the sensed deviations. The 
countertorques exactly balance the torques due to frictional forces at the 
gimbal bearings and/or stiffness of the electrical cables. Thus, the 
appended body 20 of a space satellite according to the present invention 
can be moved without imparting motion to the main body 10. 
The present invention has been described above in terms of a preferred 
embodiment. However, variations in particular design features would be 
apparent to workers skilled in the art upon perusal of the above 
description and the accompaning drawing. Therefore, the foregoing 
description is to be understood as illustrative of the invention. However, 
the scope of the invention is not limited to any particular embodiment, 
but rather is defined by the following claims and their equivalents.