Abstract:
A moment control device for positioning a spacecraft which employs a plurality of spinning bodies operable to impart a desired torque to a space craft, the bodies being constructed in a unitary combination and the unitary combination being mounted to the spacecraft to be positioned.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to vehicle control and more particularly to attitude control apparatus for spacecraft. Yet more specifically, the present invention relates to a novel arrangement for reaction wheels (RW&#39;s), momentum wheels (MW&#39;s) and/or control moment gyros (CMG&#39;s) used to position a space vehicle and control its attitude. 
     2. Description of the Prior Art 
     RW&#39;s, MW&#39;s and CMG&#39;s have long been used to control the attitude of space craft. They were designed to replace reaction jets for cyclic type maneuvers and provided improved control, longer spacecraft life and reduced fuel requirements. In the prior art, a number, at least three and usually more for redundancy and fail safe operation, have been mounted at various locations about the spacecraft as dictated by volume, structure and thermal considerations and where sufficient space was available. The components, being remote from one another, have a number of disadvantages. For example, several boxes of electronics are required for each component and this introduces greater weight, volume, cost and heat generation. With the prior art each CMG, MW or RW has to be installed separately and many of the functions have to be calibrated and tested by the spacecraft integrator. The prior CMG art requires a gimbal rate sensor or tachometer which when combined with other data can be used to derive an approximation of the torque being delivered to the spacecraft. The sensor its self induces higher frequency signals which results in added and undesirable vibrations being transmitted to the spacecraft. Finally, the prior CMG art uses ring like mounting structures which are bolted to a plate structure in the spacecraft. This primarily two-dimensional structure is structurally inefficient. Each unit delivers torque to the spacecraft as its single function, but the net effect cannot be measured directly and efficiently. As will be shown, the present invention is more of a three-dimensional system having somewhat equal dimensions in all three spatial axes. 
     BRIEF DESCRIPTION OF THE INVENTION 
     The present invention overcomes the problems of the prior art by providing an integrated single unitary structure containing multiple spinning bodies, the electronics to control them and the intelligence to interface with the spacecraft on the sub-system level. Repackaging the system into a single integrated unit allows the manufacturer to design a single and more efficient structure to house the inner gimbal assemblies of the RW&#39;s, MW&#39;s or CMG&#39;s. With CMG&#39;s the outer gimbal base ring as a separate part, can be eliminated and the lower number of electronic boxes can be reduced resulting in reduced weight and the number of connections to the vehicle further reduces weight and improves heat conduction transfer to the outer surfaces of the vehicle. Further more, the unit manufacturer now has to deal with a single momentum control unit and, on delivery, the unit will be able to be plugged in to reduce the contractor cost and improve the spacecraft manufacturing schedule. The total volume occupied by the unitary structures will be smaller than that of separate mounted units. The arrangement and other design changes enabled by the grouping together converts the result into a higher level system offering many advantages including, better performance, the ability to measure the performance more accurately and more directly, lower weight, lower power, a smaller package, and lower cost. Using common circuits within the system reduces the number of electronic components. The number and resulting weight of cables and connector is also be reduced. Grouping them and adding a single six axes interfacing kinematics force measuring system, such as is proposed herein, enables a much more effective servo control system to regulate the resulting net torque and consequential motion of the spacecraft. This further enables the addition of a single isolation system as an integral part of the force measuring system to reduce or filter undesirable high frequency vibratory forces that are otherwise transmitted to the spacecraft. The present invention not only avoids this problem it also diminishes the effects of other sources of high frequency vibrations such as that caused by the bearings and rotor unbalance. The present invention allows the manufacturer to complete this process much more professionally before it leaves the factory and in parallel with the building of the spacecraft. By operating on a higher level set of requirements the manufacturer can relax sub-level requirement lowering his manufacturing cost. Installation into the spacecraft is much simpler. The result is lower spacecraft integrator cost and significantly reduced spacecraft manufacturing schedules which further reduces cost. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is one preferred embodiment of a unitary structure of the present invention housing a plurality of control moment gyros; 
     FIG. 2 is one preferred mounting structure attaching the unitary structure of FIG. 1 to a vehicle; and, 
     FIG. 3 is another preferred embodiment of the unitary structure of the present invention. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     Referring to FIG. 1, a unitary moment control unit (MCU)  10  is shown containing four control moment gyros (CMG&#39;s) identified by reference numerals  12 ,  14 ,  16  and  18 . While the structure of MCU  10  is shown as a symmetrical  14  sided structure, other configurations are possible and while CMG&#39;s are used for the preferred embodiment, other rotating mass members such as reaction wheels or momentum wheels may be used. It should also be understood, that while four CMG&#39;s have been shown, fewer may be used. Three CMGs are necessary for three axis control and the fourth CMG is for redundancy. Also more than four CMG&#39;s may be employed for fail safe and fail operational functions. An electronic package  20  having an input connection  22  and providing connection to each of the CMG&#39;s is shown by connectors  24 ,  26   28  and  30 . Having the electronics co-located with the CMG&#39;s reduces the amount of electronics required and allows the unit to be tested with all of the CMG&#39;s in place. 
     CMG  12  is mounted to the MCU  10  by a mounting member  40  on one end and by an unseen centrally located mounting member in a manner to be described in connection with FIG.  3 . CMG  14  is mounted to the MCU  10  by a mounting member  42  on one end and by the unseen centrally located mounting member. CMG  18  is mounted to the MCU  10  by a mounting member  44  on one end and by the unseen centrally located mounting member. CMG  16  is mounted similarly to the others by two unseen mounting members. 
     The MCU is held in a rigid arrangement made up of a plurality of terminal members shown as spheres  50  and a plurality of elongated joining members  52 . Electronics box  20  may be connected to one or more of the elongated members  52 . 
     It is seen that the orientation of the CMG&#39;s is such that they do no lie on any axis in common. Thus, the torque imparted by the CMGs will be resolved into a three orthogonal axes arrangement by using vector addition of their individual torque&#39;s. In this manner the space craft to which the MCU is mounted, as will be described in connection with FIG. 2, will be oriented as desired. If an CMG fails, the fourth CMG may be used to supply any missing torque. 
     The mounting of the MCU  10  to the spacecraft is important in order to impart the necessary torque and to minimize emitted vibration. The mounting should also be kinematic to minimize strain due to temperature changes This may be accomplished by a strut type element which has relatively high stiffness along its longitudinal axis and relatively low stiffness in the other axes. Co-pending application of David Osterberg entitled Load Isolator apparatus filed Jan. 29, 1997 with Ser. No. 08/790,647 and assigned to the assignee of the present invention describes a load isolator damper arrangement which may be used. The number of struts used to mount the MCU is also important. One of the most stable ways to mount a structure is by a hexapod arrangement also known as a Stewart Platform as will be described in connection with FIG.  2 . 
     In FIG. 2, a small portion of the structure of FIG. 1 is shown somewhat enlarged for clarity. Three mounting members  60 ,  62  and  64  which may be members such as shown as mounting members  40  of FIG. 1, are shown joined by elongated joining members  66 ,  68  and  70  which may be a suitable three of the joining members  52  of FIG. 1. A hexapod mounting consisting of six struts  72 ,  74 ,  76 ,  78 ,  80  and  82  are shown. Struts  72  and  74  each have one end connected to mounting member  60  while their other ends are pivotally connected to pivots  84  and  86  so as to be rotatable about axes  88  and  90  respectively. Struts  76  and  78  each have one end connected to mounting member  62  while their other ends are pivotally connected to pivots  94  and  96  so as to be rotatable about axes  98  and  100  respectively. Struts  80  and  82  each have one end connected to mounting member  64  while their other ends are pivotally connected to pivots  102  and  104  so as to be rotatable about axes  106  and  108  respectively. Pivots  84 ,  86 ,  94 ,  96 ,  102  and  104  are each connected to the spacecraft as shown by hatched lines  110 . Struts  72 ,  74 ,  76 ,  78 ,  80  and  82  are designed to include a predetermined desired amount of static stiffness and passive damping. The passive isolation system becomes a mechanical low pass filter which transmits desired torque&#39;s to the spacecraft while eliminating unwanted higher frequency vibrations. Furthermore, the passive isolation system reduces structural and bearing loads during launch, reduces weight and power consumption and allows the use of smaller bearings which emit less vibration and can be operated at higher speeds while providing longer life. By tuning the spin rotor bearing mount and adding passive viscous damping at the interface, each CMG can provide damping and a measure of vibration isolation. 
     By using force sensors  111  within the struts and feeding the information to a control system  112  the precision of the torque transmitted can be improved. While only two force sensors  111  are shown connected to control  112  for simplicity, each of the struts  72 - 82  would be similarly connected. By controlling the actual forces emitted by the MCU array rather than from each CMG an increase of dynamic range and accuracy of the entire momentum control system can be achieved. Finally, an actuator can be added to each strut to provide an active isolation control capability. This can be used to lower the frequency with which isolation and torque control can be provided. 
     As mentioned, while a hexapod arrangement is preferable, there may be situations where it is more practical to use more mounting members. For example, if a rectangular package is used, and eight strut arrangement with two struts at each of four comers might be preferred. 
     FIG. 3 shows an octopod mounting arrangement. 
     In FIG. 3, a unitary momentum control unit  120  is shown with four CMG&#39;s  122 ,  124 ,  126  and  128  mounted therein. CMG  122  is mounted at one end to a mounting member  130  similar to the mounting arrangement of FIG.  1 . The other end of CMG  122  is mounted to a central mounting member  134  and the spinning mass therein (not seen) rotates about an axis  136 . CMG  124  is mounted at one end to a mounting member  140  and the other end is mounted to the central mounting member  134  and its spinning mass (not seen) rotates about an axis  144 . CMG  126  is mounted at one end to a mounting member  150  and the other end is mounted to the central mounting member  134  and its spinning mass (not seen) rotates about an axis  154 . CMG  128  is mounted at one end to a mounting member  160  and the other end is mounted to the central mounting member  134  and its spinning mass (not seen) rotates about an axis  164 . The structure of MCU  120  is otherwise like the structure of MCU  10  in FIG.  1  and will not be described further except to note that mounting members  130 ,  140 ,  150  and  160  are mounted to the spacecraft (shown by hash marks  170 ) by eight struts  172  with two at each comer. The struts  172  may be the same as described in connection with FIG.  2 . The electronics box  20  of FIG. 1 has been omitted form FIG. 3 for purposes of clarity. 
     It is seen that the package comprising unitary MCU  10  is more spherical in general overall shape while the structure of the package of MCU  120  is more flat. Other shapes of structures may also be used so as to provide a shape best suited for the space availability of the spacecraft. It is seen that the unitary structure makes it easy for the manufacturer to test the dynamics of the system prior to mounting in the spacecraft and that a single mounting is all that is necessary to make it operational. 
     Many changes and modifications will occur to those having skill in the art. For example, as mentioned, the unitary structure is applicable to RW&#39;s and MW&#39;s as well as the CMG&#39;s used in the preferred embodiments. Accordingly, I do not wish to be limited to the specific structures used in connection with the preferred embodiments described herein.