Spin-stabilized spacecraft and methods

Spin-stabilized spacecraft are provided which facilitate simple attitude maneuvers and the use of high-gain antennas and efficient solar panels without the need for complex control systems and an excessive number of thrusters. They include a rotor and first and second spacecraft platforms wherein the rotor and the first and second spacecraft platforms are rotatably coupled together and a spin thruster system is arranged to urge and maintain the rotor in a spacecraft-stabilizing spin about the rotor axis. A first rotary-drive mechanism couples the first spacecraft platform to the rotor and a second rotary-drive mechanism couples the second spacecraft platform to the first spacecraft platform. An attitude thruster is spaced from the rotor axis and oriented to generate an axial thrust component in the rotor for synchronous control of the inertial attitude of the rotor axis. A spacecraft of the invention can orbit the earth in an orbital plane and at an orbital rate and be controlled by the steps of (a) spinning the rotor about a rotor axis with a thruster to establish a satellite-stabilizing angular rotor momentum, (b) orienting the rotor axis in an orthogonal relationship with the orbital plane, (c) rotating the first satellite platform substantially at the orbital rate to constantly direct a portion of the first satellite platform at the earth, and (d) counter-rotating the second satellite platform substantially at the orbital rate to constantly direct a portion of the second satellite platform at the sun. The orienting step (b) is simply realized by synchronously generating an axial thrust component in the rotor that is spaced from the rotor axis.

BACKGROUND OF THE INVENTION 
1. Field of the Invention 
The present invention relates generally to spacecraft and more particularly 
to spacecraft configurations. 
2. Description of the Related Art 
FIGS. 1 and 2 illustrate conventional spacecraft configurations which are 
described in various spacecraft references (e.g., see Morgan, Walter L., 
et al., Communications Satellite Handbook, John Wiley and Sons, New York, 
1989, pp. 547-550 and Moral, G., et al., Satellite Communications Systems, 
John Wiley and Sons, New York, 1996, pp. 511-520). In particular, FIG. 1 
illustrates an exemplary spin-stabilized spacecraft 20 and FIG. 2 
illustrates an exemplary body-stabilized spacecraft 40. 
In spin-stabilized spacecraft, all or a significant portion of the 
spacecraft spins about a spin axis. When an entire spacecraft spins, it is 
generally referred to as a "spinner". In contrast, spacecraft which have a 
nonspinning shelf portion are referred to as "dual spinners" or 
"gyrostats". For example, FIG. 1 shows a gyrostat spacecraft 20 which has 
a spinning drum 22 and a despun shelf 24. The drum spins about a spin axis 
25 and is typically cylindrical with its outer surface carrying a solar 
cell array 26. In addition to this array, the drum 22 usually contains 
batteries, fuel tanks and thrusters. 
The spacecraft's payload is generally mounted on the despun shelf 24. In 
FIG. 1, this payload is symbolized by broken lines 28 and a payload 
portion is exemplified by an antenna 30 which has a feed assembly 31 that 
transmits and receives signals 32 that are reflected off of a reflector 
33. 
The reflector 33 is carried at the end of a boom 34. 
In order for spacecraft to meet their design objectives, their attitudes 
must typically remain within predetermined attitude envelopes. However, 
spacecraft are typically subjected to disturbance torques from a variety 
of external sources (e.g., solar pressures and gravity gradients). Because 
these disturbance torques will eventually rotate a spacecraft out of its 
attitude envelope, spacecraft attitude is generally controlled with 
control systems that include onboard torque generators (e.g., thrusters). 
The efficiency and lifetime of these control systems are enhanced if the 
spacecraft has a degree of natural stabilization. 
Spin-stabilized spacecraft obtain their attitude stabilization via the 
gyroscopic principal. A disturbance torque T that is applied to a 
nonspinning spacecraft along an axis about which the spacecraft has a 
moment of inertia I will urge the spacecraft to rotate about that axis 
with a constant angular acceleration T/I. This angular acceleration will 
quickly rotate the spacecraft to the limit of its attitude envelope and, 
therefore, quickly require corrective control action. 
In contrast, if the gyrostat 20 of FIG. 1 has a angular momentum H about 
its spin axis 25 and is subjected to the same disturbance torque T about 
an axis that is orthogonal to the spin axis, the spacecraft will rotate 
about a third orthogonal axis with an angular velocity T/H. Because an 
acceleration has been converted to a velocity and because the angular 
momentum H can be made quite large, the time before a corrective action is 
required has been greatly extended. 
The body-stabilized spacecraft 40 of FIG. 2 has a pair of solar panels 41 
and 42 which extend away from a nonrotating body 44. One side of each of 
the solar panels is covered with a solar cell array 46. The satellite's 
payload is carried within or on the body 44. In the case of a 
communications spacecraft, the payload typically includes antennas 48. 
The attitude of a body-stabilized spacecraft which moves in an orbital 
plane about a celestial body is often referenced to an orthogonal 
coordinate system having pitch, yaw and roll axes. The pitch axis is 
orthogonal to the orbital plane, the yaw axis points to the celestial body 
and the roll axis is aligned with the spacecraft's velocity vector. In a 
geostationary orbit, the solar wings 41 and 42 are typically oriented 
along the pitch axis. 
The body-stabilized spacecraft 40 lacks the attitude stabilization that the 
spin-stabilized spacecraft 20 of FIG. 1 gains from its spinning body. This 
stabilization is replaced with an attitude control system that typically 
includes flywheels whose rotation provides an on-board angular momentum. 
FIGS. 3A-3C show schematized views of the spacecraft 40 of FIG. 2 with 
FIGS. 2 and 3A-3C having like elements indicated by like reference 
numbers. These figures illustrate exemplary stabilization elements for 
body-stabilized spacecraft. 
In FIG. 3A, an attitude control system includes a single spinning flywheel 
50 which is typically referred to as a momentum wheel. The angular 
momentum H of the momentum wheel 50 provides a measure of gyroscopic 
rigidity (similar to that provided in the spin-stabilized spacecraft 20 of 
FIG. 1 by the spinning spacecraft body 22). In particular, the gyroscopic 
rigidity of the momentum wheel 50 limits movement about the roll and yaw 
axes and pitch attitude control is realized by modulating the wheel's 
velocity to thereby exchange momentum between the wheel and the 
spacecraft's body 44. 
Roll attitude control is obtained with an actuator (e.g., a thruster, 
magnetic coil, or use of the solar wings as solar pressure sails) which 
can generate attitude-correction torque about this axis. In the course of 
the orbit, there is an interchange every 6 hours between the roll and yaw 
axes so that yaw control is facilitated through roll control. 
In the control system of FIG. 3A, the amplitude of the angular momentum H 
can be altered but not its orientation. This single degree of freedom is 
expanded to two degrees of freedom in FIG. 3B by adding a second momentum 
wheel 51 and inclining the wheel's axes relative to each other. The wheels 
50 and 51 now have momentums H.sub.1 and H.sub.2 respectively. This system 
permits the orientation of the combined angular momentum to be changed so 
that the attitude of a spacecraft element (e.g., an antenna) can be 
rotated to a desired inertial attitude (i.e., the control system of FIG. 
3B is more agile than that of FIG. 3A). 
Because the momentum wheels of FIGS. 3A and 3B maintain a significant 
angular momentum at all times, attitude control systems of this type are 
referred to as "biased-momentum systems". However, spacecraft agility is 
enhanced if stored momentum is kept low because this facilitates rapid 
attitude corrections. FIG. 3C shows a control system which includes three 
reaction wheels 60, 61 and 62 that are each aligned with one of three 
orthogonal spacecraft axes. The velocity of reaction wheels can be varied 
in either direction from a rest state in which they have no velocity. 
Accordingly, the attitude control systems of such body-stabilized 
spacecraft are referred to as zero-momentum systems. 
With this system, disturbance torques are corrected by continually varying 
the wheel velocities. Because the mean wheel velocity is near zero, 
gyroscopic rigidity of the wheels is minimal which facilitates quick 
attitude changes. However, realizing this agility not only requires a 
reaction-wheel system but typically also requires addition of control 
systems with complex, costly elements (e.g., gyroscopic inertial reference 
units, static earth sensors, magnetometers, star sensors and control 
processors). In addition, a large number of thrusters are needed to unload 
the wheels' momentum when it becomes excessive. 
Although conventional spacecraft configurations (e.g., those exemplified in 
FIGS. 1, 2 and 3A-3C) have successfully stabilized a large number of 
existing spacecraft, they have several disadvantages. Because the antennas 
of spinner spacecraft must also spin, they are inherently low gain 
structures (e.g., omni or toroidal antennas). Although gyrostat spacecraft 
can carry high gain antennas, both spinner and gyrostat spacecraft have 
inefficient solar arrays. At any given time, one half of the solar cells 
of their arrays are in shadow and the effectiveness of most cells of the 
remaining one half is reduced because they are angled away from the sun. A 
small portion of a gyrostat's solar cells would meet its energy needs if 
fully exposed to solar radiation. The weight of the additional solar cells 
must be subtracted from the spacecraft's payload which significantly 
reduces the spacecraft's revenues. 
Body-stabilized spacecraft can carry high-gain antennas and efficient 
planar solar panels. However, momentum-biased systems lack attitude 
agility and zero-momentum systems require complex control systems and a 
large number of momentum-unloading thrusters. 
SUMMARY OF THE INVENTION 
The present invention is directed to spacecraft configurations which 
facilitate simple attitude maneuvers and the use of high-gain antennas and 
efficient solar panels without the need for complex control systems and an 
excessive number of thrusters. 
These goals are achieved with a spin-stabilized spacecraft that has a rotor 
and first and second spacecraft platforms wherein the rotor and the first 
and second spacecraft platforms are rotatably coupled together and a spin 
thruster system is arranged to urge and maintain the rotor in a 
spacecraft-stabilizing spin about the rotor axis. 
Preferably, the spacecraft includes a first rotary-drive mechanism which 
couples the first spacecraft platform to the rotor and a second 
rotary-drive mechanism which couples the second spacecraft platform to the 
first spacecraft platform. Thus, the inertial attitude of the first and 
second spacecraft platforms are responsive to the first and second 
rotary-drive mechanisms. Preferably, the spacecraft also includes an 
attitude thruster that is spaced from the rotor axis and oriented to 
generate an axial thrust component in the rotor for synchronous control of 
the inertial attitude of the rotor axis. 
Spacecraft of the invention are suited for orbits of all altitudes and 
orbital rates and their structures facilitate simple spacecraft control 
methods. For example, a satellite orbiting the earth in an orbital plane 
and at an orbital rate can be controlled by: (a) spinning the rotor about 
a rotor axis with a thruster to establish a satellite-stabilizing angular 
rotor momentum, (b) orienting the rotor axis in an orthogonal relationship 
with the orbital plane, (c) rotating the first satellite platform 
substantially at the orbital rate to constantly direct a portion of the 
first satellite platform at the earth, and (d) counter-rotating the second 
satellite platform substantially at the orbital rate to constantly direct 
a portion of the second satellite platform at the sun. The orienting step 
(b) is simply realized by synchronously generating an axial thrust 
component. 
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.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
FIG. 4 illustrates a spacecraft 80 which has a rotor 82, first and second 
spacecraft platforms and a spin thruster system 84. The rotor and the 
first and second platforms are rotatably coupled together and the spin 
thruster system is arranged to urge and maintain the rotor 82 in a 
spacecraft-stabilizing spin about a rotor axis 86. 
In the spacecraft embodiment 80, the first spacecraft platform is a payload 
platform 88 and the second platform is a solar panel 90 which carries a 
solar cell array 91. A first rotary-drive mechanism 92 couples the payload 
platform 88 to the rotor 82 and a second rotary-drive mechanism 94 couples 
the solar panel 90 to the payload platform 88. In addition, a tilt 
mechanism 96 tiltably couples the solar panel 90 and the second 
rotary-drive mechanism 94. 
The spin thruster system 84 includes a first spin thruster 98 which is 
oriented to generate a first tangential thrust component in the rotor 82 
and a second spin thruster 99 which is oriented to generate a second 
tangential thrust component that rotatably opposes the first tangential 
thrust component. The spin thruster system 84 also preferably includes a 
pair of backup spin thrusters of which one thruster 100 is visible in FIG. 
4. Preferably, the thrusters of the spin thruster system 84 are also 
tilted to each generate an axial thrust component in the rotor 82. 
The rotor 82 also carries an attitude thruster 102 which is spaced from the 
rotor axis 86 and oriented to generate an axial thrust component in the 
rotor 82. The attitude thruster 102 is backed up by a backup attitude 
thruster 104. The backup spin and attitude thrusters insure control of the 
spacecraft 80 in the event of failure of one or more of the primary 
thrusters. The spacecraft 80 may also include other thrusters such as the 
axially positioned liquid apogee motor (LAM) 105 which is particularly 
suited for effecting orbital maneuvers (e.g., transition from a transfer 
orbit to a geostationary orbit). 
A plurality of fuel tanks 106 and a plurality of batteries 108 are carried 
by the rotor 82. Spacecraft energy is stored in the batteries and the fuel 
tanks are coupled to deliver fuel to the spin and attitude thrusters. 
Celestial-body sensors in the form of an earth sensor 110 and a sun sensor 
112 are also carried by the rotor 82. 
In addition to an electric motor, the first rotary-drive mechanism 92 
preferably includes rotation bearings and a power-transfer assembly (e.g., 
one including slip rings) to couple electrical power and electric signals 
between the rotor 82 and the payload platform 88. Similar mechanisms have 
been used in various conventional spacecraft and are sometimes referred to 
as a bearing and power transfer assembly (BAPTA) motor. Exemplary versions 
of the second rotary-drive mechanism 94 and the tilt mechanism 96 have 
also been used in conventional spacecraft and are sometimes referred to as 
a solar wing drive (SWD) and an adjustable solar wing assembly (ASWA) 
respectively. 
An operational use of the spacecraft 80 is shown in FIG. 5 where it travels 
along an orbital path 114 as it orbits the earth 116 in an orbital plane 
117 and at an orbital rate. In this orbit, the thrusters of the spin 
thruster system (84 in FIG. 4) are fired as required to speed up and slow 
down the spin rate of the rotor 82 and thereby maintain its angular 
velocity at a predetermined rate (e.g., in the range of 5 to 100 
revolutions per minute). 
In addition, the attitude thruster (102 of FIG. 4) is fired in synchronism 
with the rotor spin rate to change the attitude of the rotor axis 86 as 
required to keep it in an orthogonal relationship with the orbital plane 
117 (i.e., along an axis generally referred to in body-stabilized 
spacecraft as a pitch axis). 
The attitude thruster can be considered to lie on a first transverse axis 
of the spacecraft 80. If the rotor were not spinning, the axially-oriented 
thrust of the attitude thruster would induce a torque in the spacecraft 
about a second transverse spacecraft axis that is orthogonal to the first 
transverse axis. However, the angular momentum of the rotor gyroscopically 
converts this torque into one about the first transverse axis. 
Because the attitude thruster is mounted on the spinning rotor, it can be 
synchronously fired at any angle about the rotor axis 86 to thereby move 
the axis to any desired inertial attitude. Pulsed input signals from 
celestial-body sensors such as the spinning earth sensor (110 in FIG. 4) 
and sun sensor (112 in FIG. 4) provide the inertial information required 
to time the synchronous firing of the attitude thruster. 
The spinning rotor has a large moment of inertia which causes its angular 
momentum to be considerably larger than any other motion-induced momentum 
in the spacecraft 80. The rotor's moment of inertia is enhanced by the 
weight of elements that are carried by it (e.g., the fuel tanks (106 in 
FIG. 4) and the batteries (108 in FIG. 4)). Accordingly, the effects of 
disturbance torques upon the attitude of the spacecraft 80 are greatly 
reduced and the attitude of the rotor axis requires only an occasional 
realignment. 
The large angular momentum of the spinning rotor and its orthogonal 
orientation to the orbital plane 117 also facilitates a simple rotation of 
the payload platform 88 at the orbital rate to keep a portion of it (e.g., 
communication antennas) directed at the earth 116. The momentum and its 
orientation further facilitates a simple counter-rotation of the solar 
panel 90 at the orbital rate to keep its solar cell array (91 in FIG. 4) 
directed along a sun line 118 to the sun 119. These rotations are effected 
respectively with the first and second rotary-drive mechanisms (92 and 94 
in FIG. 4). When there is an appreciable sun elevation with respect to the 
orbital plane 117, the exposure of the solar array to the sun can be 
enhanced by tilting it with the tilt mechanism (96 in FIG. 4). 
Because of disturbances from various celestial sources, errors in the 
inclination and eccentricity of the spacecraft's orbital path 114 
generally develop over time. The same disturbances cause the spacecraft 80 
to drift from its desired station relative to the rotating earth 116. The 
first and second spin thrusters (98 and 99 in FIG. 4) can be synchronously 
fired together to correct eccentricity and drift errors with their 
tangential thrust components. Undesired inclination disturbances due to 
their axial thrust components can be avoided by firing at first and second 
opposed sides of the orbital path 114. 
The first and second spin thrusters (98 and 99 in FIG. 4) can also be 
synchronously fired together to correct inclination errors with their 
axial thrust components. Undesired eccentricity disturbances due to their 
tangential thrust components can be avoided by firing at first and second 
opposed sides of the orbital path 114. 
Because of its configuration, the spacecraft 80 of FIG. 4 displays 
significant operational and control advantages when compared to 
conventional spacecraft. As opposed to the low-gain antennas of spinner 
spacecraft, it can carry high-gain antennas on at least one of its 
spacecraft platforms. In contrast to the inefficient cylindrical solar 
cell arrays of spinner and gyrostat spacecraft, the planar solar panel 90 
can be steered to be in an efficient orthogonal relationship with a sun 
line (118 in FIG. 5). 
Control of its spin rate, spin axis attitude and stationkeeping (i.e., 
control of inclination, eccentricity and drift) is effected with three 
thrusters (six with backup thrusters) in the spacecraft 80. In contrast, 
control of body-stabilized spacecraft typically requires six thrusters 
(two spaced from each of the spacecraft's roll, pitch and yaw axes) which 
become twelve with backup thrusters. 
Because the attitude of the platforms of the invention can be easily and 
efficiently altered, they can carry other thrusters (e.g., ion engines) 
for effecting various orbit maneuvers (e.g., orbit raising, deorbiting and 
stationkeeping). 
Body-stabilized spacecraft typically require a propulsion management system 
to urge thruster fuel from fuel tanks to thrusters. This system includes 
various fuel-directing structures (e.g., valves and traps). Because of the 
centrifugal forces generated by the spinning rotor 82, fuel can be 
automatically urged from the fuel tanks 106 to the spin thrusters 98 and 
99 and the attitude thruster 102. 
Conventional spacecraft typically require heat pipes to conduct heat away 
from their batteries. In contrast, the batteries 108 are carried on the 
rotor 82 and easily positioned for direct heat radiation into space. 
As previously mentioned, attitude control of body-stabilized spacecraft 
generally requires the use of complex attitude sensors, actuators and 
controllers (e.g., gyroscopic inertial reference units, static earth 
sensors, magnetometers star sensors, reaction or momentum wheels and 
control processors). These complex attitude control systems are not 
required in the spacecraft 80. Instead, accurate attitude control (e.g., 
better than 0.1 degree) is obtained by controlling rotation and 
counter-rotation of the payload platform 88 and solar panel 90. This can 
be effected with a simple attitude control electronics coupled between 
simple spinning sensors (e.g., the earth and sun sensors 110 and 112) and 
the first and second rotary-drive mechanisms 92 and 94. 
Various disturbance forces can cause the rotor axis 86 to nutate or "cone". 
In the spacecraft 80, this nutation can be damped out by coupling the 
signal from a rotor-mounted accelerometer to the first rotary-drive 
mechanism 92. In operation of this damping system, torque variations are 
applied to the first and second spacecraft platforms 88 and 90 and timed 
so that they decrease the nutation. 
FIG. 6 illustrates another spacecraft embodiment 120 whose configuration is 
similar to the spacecraft 80 of FIG. 4. In particular, the spacecraft 120 
has a rotor 122 which spins about a rotor axis 123, a payload platform 124 
and a solar panel 121. The rotor 122 carries spin thrusters (not shown) 
which maintain a spacecraft-stabilizing spin of the rotor 122. The rotor 
also carries an attitude thruster (not shown) which adjusts the inertial 
attitude of the rotor axis 123. 
The payload platform 124 has a communications module 125 and a 
communications antenna 126 which includes antenna feeds 127 that receive 
and transmit communication signals 128 through reflections from an antenna 
reflector 129. The reflector 129 is formed of a tension web 130 that is 
stretched into a reflector shape by a hoop-truss arrangement 132. This 
structure can be stowed in a small volume during launch and subsequently 
deployed at the end of an antenna boom 133. 
The payload platform also carries deployed radiators 136 which enhance heat 
radiation from the communication module into space. The solar panel 121 
incudes tilted reflectors 138 that reflect solar radiation onto a solar 
cell array 139 for increased energy generation. 
In contrast to the spacecraft 80, the rotation axes of the first and second 
spacecraft platforms 124 and 121 of the spacecraft 120 are arranged in a 
coaxial relationship (i.e., they both rotate about the axis 123). Although 
not shown, the second rotary-drive mechanism of the solar panel 121 
extends through the first rotary-drive mechanism of the payload platform 
124 to facilitate rotation between the solar panel and the payload 
platform. 
FIGS. 7A and 7B illustrate another spacecraft embodiment 140 that is 
similar to the spacecraft 80 of FIG. 4. The spacecraft 140 has a rotor 142 
which spins about a rotor axis 143, a payload platform 144 and a second 
spacecraft platform in the form of a solar panel 146. The payload platform 
144 is rotatably coupled to the rotor 142 by a first rotary-drive 
mechanism 148. The payload platform is laterally offset from its 
rotary-drive mechanism 148 with a shelf 168 and a brace 169. The solar 
panel 146 carries a solar cell array 154 and is coupled to a second 
rotary-drive mechanism 150 by a for rotation about a parallel and offset 
axis 155. 
The rotor 142 carries spin thrusters (not shown) which maintain a 
spacecraft-stabilizing spin of the rotor. The rotor also carries an 
attitude thruster (not shown) which adjusts the inertial attitude of the 
rotor axis 143. A plurality of fuel tanks 156 and a plurality of batteries 
158 are carried by the rotor 142. 
The payload platform 144 carries a plurality of communications antennas 160 
which each includes an antenna feed 162 that receives and transmits signal 
radiation 164 through reflections from a corresponding antenna reflector 
166. 
Spacecraft of the invention are suited for orbits of all altitudes and 
orbital rates. As exemplified in embodiments of the invention, the first 
and second spacecraft platforms can house a variety of spacecraft 
structures and separately direct them at celestial bodies. 
As a first example, the first spacecraft platform can include structures 
for communicating with an observing the earth (e.g., antennas, radiation 
sensors and weather sensors) and direct these structures at the earth by a 
simple rotation at the spacecraft's orbital rate. As a second example, the 
second spacecraft platform can include structures for receiving and 
observing solar radiation (e.g., solar cell arrays and solar-observation 
instruments) and direct these structures at the sun by a simple 
counter-rotation at the spacecraft's orbital rate. Although these examples 
envision the earth and the sun as the celestial bodies, the invention can 
be operated in association with other celestial bodies. 
These processes of rotation and counter-rotation are facilitated by the 
spacecraft's inertial stability and that stability is established by the 
angular momentum of the spinning rotor to which the first and second 
platforms are coupled. 
As also shown in embodiments of the invention, the first and second 
rotary-drive mechanisms may be coaxially arranged or can be laterally 
offset from one another. One platform may be coupled to another platform 
which is, in turn, rotatably coupled to the rotor. Alternatively, both 
platforms may be rotatably coupled directly to the rotor although this 
arrangement is more complex because both rotary-drive mechanisms are then 
associated with the high spin rate of the rotor. 
Although spacecraft embodiments have been shown with single solar panels, 
the teachings of the invention also facilitate the use of multiple solar 
panels (e.g., panels which extend above and below the rotor). The 
teachings of the invention generally reduce the need for complex 
stabilization and attitude-control structures. However, embodiments of the 
invention can be supplemented by such structures in specific applications. 
For example, attitude control of the rotor axis may be further enhanced by 
adding reaction wheels to one of the spacecraft platforms. These wheels 
would preferably be arranged with their spin axes orthogonal to the rotor 
axis. 
Other structures can be advantageously added to the illustrated 
embodiments. For example, flywheels have been used as energy sources by 
decreasing their rotational velocity and converting this loss of momentum 
to electrical energy. Conventionally, these "flywheel batteries" are used 
in counter-rotating pairs so that their momentum loss does not induce 
attitude changes in body-stabilized spacecraft. In contrast, any number of 
these devices can be carried on the spinning rotors of the invention 
because their large momentum effectively isolates the spacecraft platforms 
from the smaller momentum change of the flywheels. 
While several illustrative embodiments of the invention have been shown and 
described, numerous variations and alternate embodiments will occur to 
those skilled in the art. Such variations and alternate embodiments are 
contemplated, and can be made without departing from the spirit and scope 
of the invention as defined in the appended claims.