Abstract:
A feedback system for use in controlling control moment gyroscopes that are used in positioning a spacecraft where the relative rate of rotation is determined from the difference between the rotation rate of the gimbals of the CMGs and the rotation rate of the frame about the gimbal axis. The inertial reference units on the spacecraft are utilized as the source of measured rates of rotation of the three spacecraft axes with respect to a predetermined axis and these are converted to the rates of rotation of the frame about the individual CMG gimbal axes.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     U.S. patent application entitled Relative Rate Sensor for Control Moment Gyroscopes, of Gerald K. Foshage, Ser. No. 08/821142 filed Mar. 19, 1997 and assigned to the assignee of the present invention. 
     BACKGROUND OF INVENTION 
     1. Field of the Invention 
     This invention relates to the production of feedback signals indicative of the rate of rotation of a system, such as a space craft for use in controlling a torque applied to the space craft with drive devices such as control moment gyros (CMGs). 
     2. Description of the Prior Art: 
     CMGs are well known in the art for use in controlling the torque applied to a space craft to control its orientation or rate of rotation with respect to reference axis or to another body such as earth. For example, a satellite in space may have its rate of rotation controlled about one or more axes (usually three) by fixing the gimbal frames of one or more CMGs (usually four to provide redundancy) to the frame of the spacecraft and applying a command rate to the gimbals of the CMGs which then create torques which are transferred to the frame of the spacecraft to obtain the desired rotation rate. In such systems it is desirable to provide closed loop feedback for better control by assuring that the actual rates are the same as the commanded rates. Such systems have employed tachometers to supply a signal indicative of the relative rate of rotation between the gimbal frame and the inner gimbal assembly. However, tachometers have greater size and weight than is desired, have low rate threshold (the signals can&#39;t be discerned from the noise), have variable rate ripple (voltage that varies without change in gimbal rate) and must be mounted colinearly with the gimbal. A better system has been described in the above referenced Foshage application where a pair of fiber optic gyros are employed, one to measure the absolute rate of rotation of the inner gimbal assembly and one to measure the absolute rate of rotation of the frame attached to the spacecraft. While this system provides excellent rate control, it requires the use of two separate absolute rate sensors (e.g. fiber optic gyros). 
     SUMMARY OF THE INVENTION 
     The present invention eliminates the need for one of the absolute rate sensors of the above referenced Foshage application by utilizing the outputs of the inertial reference units i.e., rate gyroscopes, (e.g. ring laser gyros) which are used on the spacecraft. The output of the spacecraft inertial reference units is transformed from three axis signals having reference to the space craft frame to signals relative to the axes of the CMG which is then subtracted from the absolute CMG gimbal rates determined by the absolute rate sensors mounted on the rotating CMG gimbals to provide relative rate feedback signals, as in the above referenced Foshage application. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a schematic block diagram of the system of the present invention 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     In FIG. 1, three CMGs 12, 14 and 16 are shown. A gimbal frame 18 for CMG 12 is shown mounted on spacecraft 20 and torque exerted on frame 18 will be transferred to spacecraft 20. CMGs 14 and 16 are also mounted on space craft 20 but, for convenience, only CMG 12 is so shown in FIG. 1. The torques exerted by CMGs 14 and 16 are shown by arrows 22 and 24 respectively. While 3 CMGs are shown, it should be understood that for redundancy, four or more such gyros are usually used. 
     The gimbal frame 18 of CMG 12 extends upwardly in FIG. I to provide a mount for a gimbal 26 for rotation about an axis 28. Gimbal 26 provides a mount for a rotor 30 to be rotatably driven by motive means (not shown) about an axis 32. A torque applied to gimbal 26 as, for example, by a connection shown as arrow 34, will cause a torque to be applied to gimbal frame 18 and thus to spacecraft 20 to cause rotation of spacecraft 20 about a first predetermined axis. Similarly, torques applied to the gimbals of CMG 14 and CMG 16 as, for example by connections shown as arrows 36 and 38 will produce torques by arrows 22 and 24 to be applied to spacecraft 20 about second and third predetermined axes respectively. In this manner, the rotation of spacecraft 20 may be controlled so that it will, for example, become or remain pointed in a predetermined direction or at a predetermined angle with respect to a reference axis. 
     The spacecraft 20 includes inertial reference sensors such as ring laser gyros and accelerometers which operate to produce signals relating to the spacecraft&#39;s rate of rotation about the inertial axes of the gyros and signals relating to the angular position of the spacecraft with respect to the reference axis. Signals indicative of the rate of rotation of the spacecraft are produced on a line 40 and those relating to angular position may be produced on a line 42 in FIG. 1. The rate of rotation signals on line 40 are with respect to a reference system on the spacecraft and accordingly, they are shown being presented to a TRANSFORM box 44 which, knowing the orientation of each of the three CMGs with respect to the spacecraft, operates to calculate the angular rates of rotation of the spacecraft about the gimbal axes of the CMGs 12, 14 and 16. In other words, using known math techniques such as Euler equations, the rate signals relative to one set of coordinates (the spacecraft&#39;s) are converted into rate signals relative to another set of coordinates (the CMGs). Accordingly, the TRANSFORM box will produce one output for each of the CMGs used, which output is indicative of the rate of the spacecraft rotation about that CMG&#39;s gimbal axis. This signal for CMG 12 is shown on a line 46 in FIG. 1. While the signals for CMGs 14 and 16 are shown on lines 47 and 48 respectively, for example. If more than three CMGs are used there would be additional signals from TRANSFORM box 44. 
     In FIG. 1, an absolute rate of rotation sensor 50, for example a fiber optic gyro, is shown mounted on gimbal 26 to produce a signal on a line 52 indicative of the rate of rotation of the gimbal 26 about the gimbal axis 28. Similar signals would be produced on lines (not shown) from CMGs 14 and 16. 
     As was the case in the Foshage application, the signal indicative of the rate of rotation of the frame about the gimbal axis and the signal indicative of the rate of rotation of the gimbal about the gimbal axis are subtracted in a circuit 54 wherein the gimbal rate signal is shown as positive and the frame rate signal is shown as negative. The difference is indicative of the relative rate of the CMG gimbal about its gimbal axis and is presented by the circuit 54 on a line 56 as a relative rate signal. Similar signals (not shown) would also be produced by other subtraction circuits for CMGs 14 and 16. The relative rate signal on line 56 may be used as a feedback signal for CMG 12, either directly or, as shown as being presented through a second summing circuit 58 and a line 60 to a control box 62. If the actual rate of rotation of the spacecraft 20 about the gimbal axes of CMGs 12, 14 and 16 are not the same as the desired rate of rotation of the gimbal about the gimbal axes as determined by the absolute rate sensors such as 50, then the feedback signals will operate through a control box 62 to produce a change the torque such as on connection 34 applied to the gimbals in such a way as to change the rate of rotation of the space craft until the two rate signals are equal and there is no feedback signal such as on connection 34. 
     Signals indicative of the angles between the axes of the spacecraft and some predetermined reference axis on line 42 are presented to CONTROLLER box 70 which also receives a signal on a line 72 indicative of a commanded or desired angle,Θ c . If the actual angle signals on line 42 are not the same as the desired angle signals on line 72, then CONTROLLER box 70 produces commanded relative rate signal on a line 74, for the CMG 12, and similar signals indicative of the commanded relative rate for CMGs 14 and 16 on lines 75 and 76. The signals indicative of commanded relative rate on lines such as 74, which occur when the actual angles of the spacecraft are not the same as the desired angles, are presented to the summing circuits such as 65 where the relative rate signals on lines such as 56 are subtracted therefrom to produce the feedback signals on lines such as 60 for the CMG CONTROL boxes such as 62. Thus, even when the commanded rates of rotation of the spacecraft and the gimbal are the same, signals on lines such as 74 will command that the spacecraft assume a new direction with respect to the predetermined direction and the CMGs will produce torques on their gimbals to change the space craft orientation. This will produce feedback signals on lines such as 60 until the rates of rotation are again satisfied and the new angle position is achieved, at which time the signals on lines such as 60 will go to zero. 
     It is seen that we have provided a feedback system for the CMGs of a spacecraft which is lighter and simpler than the prior art and is substantially as effective as the invention of the above mentioned Foshage application while eliminating one fiber optic gyro for each CMG. Many modification and changes will occur to those having ordinary skill in the art. For example, while three CMGs have been shown, any number may be employed and while the system has been shown for use on a spacecraft, any other applications requiring rate measurements and/or feedback may utilize the invention. We therefore do not wish to be limited to the specific embodiments used in connection with the description of the preferred embodiment.