Abstract:
A self-contained momentum control system (MCS) for a spacecraft is provided for small satellites. The MCS features a miniaturized gyroscopic rotor with a rotational speed in excess of 20,000 RPM. The MCS includes at least three control moment gyroscopic mechanical assemblies (CMAs) rigidly mounted within a single enclosure, where each CMA mounted in an orientation whereby the longitudinal axis of each CMA is either orthogonal to every other CMA or is parallel to another CMA but in the opposite orientation. In order to further reduce the size of the MCS, an electronics package that is configured to interface command and control signals with and to provide power to the CMAs is included within the MCS enclosure. A plurality of shock isolation devices are used to secure each of the CMAs to the enclosure in order to reduce the launch load upon the CMAs thereby allowing the use of smaller rotor spin bearings. The MCS enclosure surrounding the CMAs and support structure is hermetically sealed.

Description:
PRIORITY CLAIMS 
     This application claims the benefit of U.S. Provisional Application No. 61/218,291, filed Jun. 18, 2009, and U.S. Provisional Application and is herein incorporated by reference in its entirety. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This invention was made with Government support under Contract FA9453-08-C-0247 awarded by the United States Air Force. The Government has certain rights in this invention. 
    
    
     BACKGROUND 
     Improved attitude agility in small satellites increases their value by improving their data collection rate and increasing the time available for transmission of that data to the ground. This is so because a time lag exists as the satellite progresses from one attitude to another for collection or transmission of the data. The longer the progression time, the shorter the time available for data collection and transmission. 
     The current state-of-the art in agile small satellites allows for slew rates of less than 1°/sec. More agility in small satellites allows for new missions that were previously unachievable such as synthesized large aperture imaging, moving ground force tracking, missile tracking, tactical imaging, space superiority and space situational awareness. These new missions may need slew rates of 2°/min and faster. 
     In the art, reaction wheel assemblies (RWA) have been used to control attitude in smaller satellites and produce slew rates in a vehicle of 2°/min or faster. However, RWA&#39;s have an inherently low torque producing capability and may take over 60 seconds to accelerate a small satellite to this slew rate, which is unacceptably long. 
     Control Moment Gyroscopes (CMG) are presently the only non-expendable actuators capable of supplying high torque (i.e. equal to or greater that 1 N-m) to achieve an acceptably high slew rate. However, because of their size (approx. 16″ disk diameter), their relatively large mass and their power consumption, these devices have historically been impractical for use with small satellites (i.e. &lt;400 kg). As a minimum, three gyros are used to control the attitude of a satellite. Therefore, CMG size is an issue. As such, there is a need for a CMG of smaller dimensions with low power consumption, while at the same time producing sufficient torque to provide sufficient attitude agility. 
     BRIEF SUMMARY 
     It should be appreciated that this Summary is provided to introduce a selection of non-limiting concepts. The embodiments disclosed herein are exemplary as variations in the novel various features of the subject matter disclosed herein may be numerous. The discussion herein is limited to a specific exemplary system for the sake of clarity and brevity. 
     An apparatus to control the attitude of a spacecraft is provided. The apparatus includes at least three control moment assemblies (CMAs) rigidly mounted within an enclosure, each CMA mounted in an orientation whereby the longitudinal axis of each CMA is one of orthogonal and inversely parallel in relation to each other. Each CMA comprises a momentum rotor configured to rotate about a first axis and a spin motor comprising a motor rotor and a non-ferric motor stator. The non-ferric motor rotor rotates about the first axis and has a first end and a second end, the first end being coaxially affixed to the momentum rotor. The Apparatus also includes a bearing concentric with the first axis in which the second end is enjournalled. 
     A self-contained momentum control system for a spacecraft is also provided. The self-contained momentum control system includes a hermetically sealed enclosure, a compartment adjoined to the hermitic enclosure, and at least three control moment assemblies (CMAs) rigidly mounted within the enclosure. Each CMA is mounted in an orientation whereby the longitudinal axis of each CMA is one of orthogonal and inversely parallel in relation to each other. An electronics package is mounted in a separate compartment adjoining the enclosure, the electronics package is configured to interface command and control signals with and to provide power to the at least three CMAs. A plurality of shock isolation devices securing each of the at least three CMAs to the enclosure is also included. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein 
         FIG. 1  is a functional block diagram of an exemplary embodiment of a satellite momentum control system; 
         FIG. 2  is an exploded view of an exemplary embodiment of a satellite momentum control system; 
         FIG. 3  is a depiction of an exemplary mounting structure for a plurality of control moment gyro mechanical assemblies (CMA&#39;s); 
         FIG. 4  is a functional block diagram of an exemplary MCS Control Electronics (MCE); 
         FIG. 5  is a cross sectional view of an exemplary CMA; 
         FIG. 6  is an exploded view of an exemplary CMA; 
         FIG. 7  is an exploded view of an exemplary Torque Motor Assembly (TMA); 
         FIG. 8  is an exploded view of an exemplary Signal Module Assembly (SMA); 
         FIG. 9  is a simplified schematic of a gimbal angle position potentiometer; and 
         FIG. 10  is an exemplary plot of the output voltage vs. gimbal angle for the gimbal angle position potentiometer. 
     
    
    
     DETAILED DESCRIPTION 
     The following detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any theory presented in the preceding background or the following detailed description. 
     The following disclosure describes a momentum control system (MCS) for small spacecraft. It is well known in the art that a minimum of three gyroscopes are required to fully control the attitude of an object in space such as a satellite. A gyroscope operates by spinning a momentum rotor mounted within in a gimbal thereby creating an angular momentum within the rotor. By Newton&#39;s first law of motion, the rotor maintains a constant velocity until an outside force imparts energy that changes the direction or magnitude of the velocity. Conversely, by applying a force against a rigidly mounted spinning gyro, the gyroscope creates an equal and opposite reactive force which changes the attitude of the satellite to which the gyroscope is secured. By using at least three gyros simultaneously, a reactive force may be created in three dimensions to change the attitude of a satellite to any desired attitude. Historically this has been done by communicating three distinct gimbal rate commands from a flight computer, one to each gyroscope. 
     The subject matter disclosed herein provides for a high torque momentum control gyroscope system featuring reduced power consumption relative to legacy systems. By replacing the traditional ferric motor stator of the gyro spin motor of a CMA with a non-ferric motor stator interacting with permanent magnets integrated with the rotor at the motor stator inner and outer diameters, the momentum rotor may be driven to rotational speeds at least four times faster than current speeds of legacy systems. Rotational speeds can range upwards from 23,000 rpm. The increase in speed allows the size of the rotor to be reduced by a similar factor of about four, thereby allowing a smaller size rotor to produce the same amount of inertia. The smaller rotor mass also allows the use of a smaller support structure. A non-limiting example of a non-ferric motor stator may be a ceramic motor stator. 
     Further, size reduction in a CMA may be achieved by utilizing a plurality of shock absorbers by which to suspend and isolate the CMAs from a vehicle frame in order to reduce the launch load on the CMA. A reduced launch load on a CMA rotor permits the use of smaller rotor spin bearings which in turn reduce drag torque caused by those spin bearing within the CMA and further reduces power consumption. Reduced launch loads also allow for the use of an even smaller and lighter CMG support structure. 
       FIG. 1  is a functional block diagram of an exemplary embodiment of a momentum control system (MCS)  100 . The MCS  100  may be mounted to a spacecraft structure  10  via a momentum exchange subsystem (MES) base  52  and an MCS control electronics (MCE) chassis  54 . Together the MES base  52  and the MCE chassis  54  are referred to as the MES base/MCE Chassis (MBEC)  50 . 
     In the particular embodiment illustrated in  FIG. 1 , the MBEC  50  preferably incorporates a hermetically sealed enclosure in which a MES  110  is mounted. The MES  110  may comprise four substantially identical control momentum assemblies (CMA)  115 , a unitary support structure  105  securing the CMAs, and a plurality of shock absorbers  60  and their attachment points  61 . Each CMA  115  within the MES  110  includes a CMG. The MES  110  may be secured to the MBEC  50  via the unitary support structure  105  by the plurality of shock isolators  60 . Additional information concerning an exemplary, non-limiting shock isolator  60  that is suitable for this purpose may be found in Honeywell patents U.S. Pat. Nos. 5,918,865, 5,332,070, and 7,182,188, which are herein incorporated by reference in their entirety. 
     In other embodiments, there may be as few as three CMAs  115  installed within an MES  110  or there may be more than four CMAs for redundancy and backup purposes. Although described herein as preferably being mounted within the MBEC  50 , MES  110  may be affixed directly to the spacecraft structure  10  or to some other suitable intermediary sub-structure of the spacecraft. A non-limiting exemplary physical implementation of the MBEC  50  is illustrated in more detail in  FIG. 2 . An exemplary, non-limiting example of the unitary support structure  105  is more finely presented in  FIG. 3 . 
       FIG. 2  is an exploded view of an exemplary MBEC  50 . As noted above, the MBEC  50  preferably incorporates a hermetically sealed enclosure that may hold a vacuum in a pressurize environment or be pressurized in an evacuated environment. If pressurized, a helium gas is preferably used as the pressurizing gas. The MBEC  50  includes the MES base  52 , the MCE chassis  54  and a cover  56 . As discussed above, the MES  110  may comprise three or more CMA&#39;s  115 .  FIG. 1  illustrates an exemplary embodiment comprising four CMAs  115 . However, one of ordinary skill in the art will recognize that at least three CMAs  115  are required for spacecraft attitude control in three dimensions although any number may be utilized for back up, redundancy, singularity avoidance and other purposes. A preferred embodiment of a CMA  115  will be described below. 
     In some embodiments, the MES base  52  may also house a Momentum Control Electronics (MCE) module  201 , an embodiment of which will now be described.  FIG. 4  is an exemplary functional block diagram of the MCE  201 . The MCE  201  is configured to control the operation of the MES  110 , and may comprise a power conditioning electronics module  202 , an input/output electronics module  204 , an array steering law and control electronics module  206 , a gimbal drive electronics module  208  and a rotor drive electronics module  211 . 
     The power conditioning electronics module  202  converts spacecraft generated power to levels required by the MCS  100  electronics. For example, in some embodiments 28v DC power is converted into 12v and 5v DC power. The power conditioning electronics module  202  also filters out any electrical noise that may be present from other systems in the spacecraft. 
     The input/output electronics module  204  provides signal conditioning for internal analog and digital signals transmitted to various components and sub-components within the MCS  100  such as motor drives, thermisters, encoders, etc., as may be known in the art. The input/output electronics module  204  may also provide output signals to a functional monitor (not shown) aboard the spacecraft. 
     The CMG array steering law and control electronics module  206  executes instructions that are recorded on a computer readable storage medium that, when executed by a processor, coordinates the gimbal motion of the CMAs  115  within the MES  110  to produce a desired resultant torque to rotate the spacecraft. The instructions process a torque command that is supplied from the spacecraft (or an operator), and the response generates and supplies individual coordinated gyro rate commands simultaneously to each CMA  115 , via the gimbal drive electronics module  208  and the rotor drive electronics module  211 . By consolidating the CMG steering law and control electronics within the MCS  100  itself, the size and weight of the MCS may be further reduced by eliminating otherwise redundant electronic components. The steering law also reduces the spacecraft level integration and testing required. Additional disclosure concerning steering laws, singularity avoidance logic and their related signal processing may be found in co-owned, co-pending application Ser. No. 11/291,706 and U.S. Pat. No. 6,128,556 which are herein incorporated by reference in their entirety. 
     The gimbal drive electronics module  208  controls the gimbal positions of the various CMA&#39;s  115  by controlling the rotation rate of a gimbal motor shaft  232  via a gimbal motor rotor  236  in each CMA. In some embodiments the gimbal motor  233  may be a 2-phase electric motor with a 50:1 harmonic drive  237  (See,  FIG. 7 ). 
     The rotor drive electronics module  211  controls a momentum rotor  410  to maintain a constant speed and therefore a constant inertia. To implement this, the rotor drive electronics module  211  is preferably implemented using a negative feedback loop that maintains the momentum rotor speed at approximately 23,000 rpm by controlling power to a rotor spin motor  412  that is coupled to the momentum rotor  410  (See  FIG. 5 ). 
     In some embodiments, the CMG array steering law and control electronics module  206 , the gimbal drive electronics module  208 , and the rotor drive electronics module  211  may be consolidated into a single controller  212  (See,  FIG. 4 ) per CMA  115 . In such embodiments, each controller  212  receives torque commands from a flight computer and generates a gimbal rate command to the associated CMA  115 . In other embodiments, the multiple CMAs  115  may be controlled by a single controller  212  by combining all CMA controllers into a single programmable logic device, such as field programmable gate arrays (FPGA). Combining controllers allows a single torque command input from a flight computer to generate discrete gimbal rate commands to each CMA  115 . Integrating the controllers also reduces space and power requirements. The controller  212  may be any suitable controller and may be a low bandwidth controller, a bandpass controller or a high bandwidth controller. All or some of the controllers  212  may operate in conjunction with one or more processors (not shown). The processors may be any suitable type of processor and may be a single processor, multiple processors operating in concert, a parallel processor, a general processor or a special purpose processor. One of ordinary skill in the art will recognize that a programmable logic device a controller and a processor are each an example of a computer readable storage medium. 
     The MES  110  is installed into a single MBEC  50 . Such a configuration not only conserves space, but allows all of the CMA&#39;s to be evacuated (i.e. the removal of atmospheric gasses) in a single operation, instead of multiple operations. Further, the single MBEC  50  allows helium, or other inert gas, to be inserted into the MES via a gas port  51  (See,  FIG. 2 ). The inert gas mitigates heat build up within the CMAs  115  by creating convection within the MBEC  50 . Heat build up may result from a number of sources including, but not limited to, friction, solar radiation, motor losses, bearing drag and conduction. 
       FIG. 5  provides a side view and a cross sectional side view of an assembled CMA  115 . Each CMA  115  comprises an inner gimbal assembly (IGA)  210 , a torque module assembly (TMA)  230  and a signal module assembly (SMA)  260 . The CMA  115  is integrated via the external structures of the IGA  210  and TMA  230 , a harmonic drive between the TMA  230  and the output of the IGA  210 , and a flexible cable conduit, or a flexible spline  270 , connecting the IGA  210  and SMA  260 . The flexible spine  270  is preferably milled from a single piece of steel bar stock with one end having a spring like configuration to accommodate a small amount of flex and expansion along the longitudinal axis (C) of the CMA  115 . 
     In the non-limiting, exemplary depicted embodiment, the entire length of the CMA  115  is approximately 11.25 inches and is approximately 4 inches in diameter at its widest point. One of ordinary skill in the art will recognize that the size of the CMA  115  may be scaled to accommodate varying momentum and torque requirements. 
       FIG. 6  provides an exploded view of a CMA  115 . Each of the three major assemblies (IGA  210 , TMA  230  and SMA  260 ) will be discussed in turn. It should be noted that various structural components such as housings, interfaces, rings, and retainers are labeled as such for reference and for completeness, although they will not be discussed herein below. Such structural elements aid in the transfer of torque to the unitary support structure  105 . 
     Referring to  FIGS. 5 and 6  together, it is seen that the IGA  210  comprises the rotor spin motor  412 , which further comprises the momentum rotor  410 , the spin motor rotor  411 , the spin motor stator  413  and other structural components. The spin motor rotor  411  may be attached to the momentum rotor  410  or may be embedded into the momentum rotor  410 . The spin motor rotor  411  is driven by the spin motor stator  413 . The momentum rotor  410  is rotationally mounted within the spin bearings  421  and rotates about an axis (A) that is disposed perpendicular to the longitudinal axis (C) of the CMA  115 . In some embodiments the rotor spin motor  412  may be a 3-phase, 4-pole DC motor with a non-ferric motor stator. 
     At least one of the spin bearings  421  is held in place by a diaphragm  420 . The diaphragm  420  is an annular metallic disk that accepts the spin bearing  421  through a hole (not shown) machined through the center of the annular diaphragm  420 . The diaphragm  420  rigidly holds the spin bearing  421  in place laterally. However, longitudinal movement of the spin bearing  421  is accommodated by the flexing of the annular diaphragm  420  in the longitudinal direction along axis A. The diaphragm  420  may be physically tuned (e.g. stiffened) by selecting different thicknesses and different material for fabrication. The diaphragm  420  therefore acts a longitudinal vibration isolator for the momentum rotor  410  and accommodates thermal expansion. Because the momentum rotor  410  spins at rotational rates at or greater than 23,000 rpm, the vibrations and their harmonics exist at very high frequencies and therefore may be absorbed by a smaller sized annular diaphragm  420 . 
     In the exemplary depicted embodiment, the momentum rotor  410  is approximately four inches in diameter and is approximately two and one half inches thick but may be scaled in size as may be necessary for a particular application. Conventional momentum rotors have been significantly larger than four inches in diameter and have rotated at angular velocities (ω) in the vicinity of 6000 rpm in order to produce the proper amount of angular momentum (H) according the formula
 
H=Iω.
 
     Therefore, in order to maintain the same level of angular momentum while reducing the size (i.e. moment of inertia (I)) of the of the momentum rotor  410 , the angular velocity (ω) must be increased proportionally. As an example, reducing the diameter of the momentum rotor  410  (i.e. it&#39;s mass) by a factor of four would require the angular velocity be increased by a similar factor resulting in a much higher angular velocity (e.g. 23,000-24,000 rpm). One of ordinary skill in the art will recognize that the required increase in angular velocity depends on the specific geometry and the material composition of the momentum rotor  410  and the spin motor rotor  411 . The momentum rotor  410  itself may be made of any suitable metallic substance such as titanium or steel. 
     Conventionally, reaching the required higher angular velocities has proven elusive as conventional spin motors require exponentially more power at higher speeds due to the natural occurrence of back electro-motive force (EMF) in the ferric motor stators of heritage rotor drive motors. A back electro-motive force develops in an electric motor from circular flows of electrons in a conductor (i.e. eddy currents) due to changing magnetic fields that result from relative motion between the conductive motor stator and the magnetic field source from the stator. 
     Eddy currents produce magnetic fields in the opposite direction from those in the spin motor rotor  411  thereby partially canceling out the electric field created by the spin motor stator  413  driving the spin motor rotor  411 . Therefore, the power necessary to increase angular velocities to levels required by a smaller rotor mass increases exponentially to overcome the eddy currents developed in a ferric rotor motor stator. Such high power requirements make the use of heritage CMA spin motors impractical. 
     In order to minimize back electro-motive force even further, Litz wires may be used as the motor stator windings ( 419 ). Litz wires are specially constructed cables comprised of small individually insulated wires physically arranged so as to minimize any back EMF that may occur in the windings themselves. The winding of the Litz wires around the motor stator may be done in any suitable fashion. An exemplary winding may be found in co-owned U.S. Pat. No. 7,061,153 B1 and is herein incorporated by reference in its entirety. 
       FIG. 7  provides an exploded view of an exemplary TMA  230  which includes a gimbal motor  233  and its housing assembly, a 50:1 harmonic drive  237  and an electronic encoder  235 / 231 . The gimbal motor  233  receives a torque command from the gimbal drive electronics module  208  and in response rotates the shaft of the gimbal motor about axis C. This rotation, in turn rotates the momentum rotor  410  away from of its steady state rotational axis A. The movement from its steady state axis thereby creates a gyroscopic torque in a direction perpendicular to the axis of rotation of the momentum rotor  410  and axis C, which causes the satellite to rotate in reaction thereto. The gimbal motor  233  comprises a gimbal motor stator  238  and a gimbal motor rotor  236  which is attached to the shaft of the gimbal motor. 
     The TMA  230  also includes an optical encoder  231 / 235 , which is rigidly affixed to the shaft of the torque motor. The optical encoder disk  231  is connected to the end of a gimbal motor shaft  232  adjacent the SMA  260 . The optical encoder disk  231  is calibrated to the position of gimbal motor  233  and measures the absolute angular position and speed of the gimbal motor  233  which may be accurate to within 21 bits per every 360°. By extension, the optical encoder is also calibrated to the axis A of the momentum rotor  410 . 
     The optical encoder comprises an encoder electronics module  235  and a glass optical encoder disk  231  with a uniquely encoded opaque pattern engraved thereon. A light source (not shown) passes light through the optical encoder disk. The light source is blocked by the unique encoded pattern as the gimbal motor shaft turns. One or more light sensors ( 269 ) read the shadowing cast by the unique encoded pattern as the shadows pass the light sensor(s) thereby tracking the position of the gimbal motor  233 . The optical encoder thereby provides position feedback (See,  FIG. 1D ) to the gimbal drive electronics module  208 . 
     On the opposite end of the gimbal motor shaft  232  from the optical encoder disk  231  and proximate to the IGA  210 , is the harmonic drive  237  that drives the gimbal of the IGA  210 . In some non-limiting embodiments, the gearing ratio of the harmonic drive  237  may be 50:1. However, a different gearing ratio may be found useful with momentum rotors  410  of different sizes. The harmonic drive  237  is used to drive the spinning rotor away from its steady state axis A and thereby create the desired maneuvering torque. The maneuvering torque is then translated through the torque motor stator  238  into the torque motor housing  239  and then through the various housing and support components of the CMA  115 , to the unitary support structure  105 , through the isolators  60  and then to the spacecraft  10 . 
       FIG. 8  provides an exploded view of the SMA  260 . The SMA  260  houses a monofilament slip ring assembly  262  for transmitting power and controlling signals to the rotor spin motor  412  and the gimbal motor  233  (See  FIGS. 3 and 4 ). The monofilament  263  may be comprised of any suitable substance. In some embodiments the substance may be gold. 
     The SMA also contains a potentiometer  510  (See,  FIG. 9 ) which provides a coarse gimbal position indication and may also be used as a primary position sensor to determine a derived gimbal angle during initialization of a CMA  115 . The potentiometer  510  may also be used as a back up gimbal position indication up for the gimbal motor  233 . The potentiometer  510  comprises a dual resistive circuit  520 / 530 . The voltage along each resistive circuit  520  and  530  varies linearly with the distance from a 5v DC power source  522 / 524  and is determined by one of two mechanically slideable taps  540 / 60  that pick off the voltage associated with the current position of the tap along each resistive circuit  520  and  530 . The slideable taps are calibrated to the position of the gimbal motor shaft  232 . As such, the voltage varies with the angular position of the gimbal motor  233 . 
     As can be seen from  FIG. 9 , each resistive circuit  520 / 530  is configured as a circle to accommodate the rotation of the gimbal motor shaft  232 . Each resistive circuit  520 / 530  contains a gap in the circle located 180° from the other to allow connection to the gimbal drive electronics module  208 . The dual track design ensures that either of the primary or the secondary tap reads a voltage and provides a position indication for the full 360° rotation of the gimbal motor shaft  232 . 
       FIG. 10  illustrates the linear voltage readout from the potentiometer at every point along the primary tack  530  and the secondary track  520 . In this particular embodiment the voltage output is a 0-5v DC. As the slideable taps  540  and  560  move clockwise beginning at point  1 , the primary tap  540  is sensing a zero voltage while the secondary tap is sensing a voltage somewhat below 2.5v DC. At position  5 , the secondary tap is sensing a maximum voltage of 5v DC and the primary tap is sensing a voltage of somewhat less than 2.5v DC. At point  4 , the secondary tap is sensing a zero voltage and begins its linear increase. At point  3 , the primary tap  540  is sensing a max voltage and the secondary tap is sensing a voltage that is somewhat less than 2.5v DC. The voltage cycle continues or reverses as the gimbal motor  233  rotates. After having read the Applicants disclosure, one of ordinary skill in the art would recognize that the voltages being sensed by the primary and secondary taps can be processed to indicate the angular position of the gimbal motor shaft  232 . 
     The IGA  210  and the SMA  260  are connected by the flexible spline  270  (See  FIGS. 5 and 6 ). The flexible spline  270  connects the gimbal ring  423  of the IGA  210  to the slip ring shaft  264  of the SMA  260  such that the gimbal ring  423  and the slip ring shaft  264  turn at the same rate. As mentioned above, the flexible spline  270  has one end cut into a machine spring allowing flexibility due to thermal expansion and encoder mount restrictions. 
     The subject matter described above is provided by way of illustration only and should not be construed as being limiting. Various modifications and changes may be made to the subject matter described herein without following the example embodiments and applications illustrated and described, and without departing from the true spirit and scope of the present invention.