Active spin axis control for spinning space vehicles

Abstract: Selective reconfiguration of existing spacecraft (10) on-board control and propulsion equipment (19-25, 30) is provided. The modification combines add-on switch means (41, 42) such as solid state multiplexers or relay banks and rate sensing gyros (43, 44) with existing accelerometers (64, 65), propulsion control electronics and thrusters (93-96). The switch means (41, 42) reconfigure the input circuits and the output circuits of the propulsion control electronics (60, 61). The switch means (41, 42) select the rate sensing gyro (43, 44) and the liquid apogee motors (91, 92) for control of spacecraft spin axis coning during perigee solid rocket motor (19) firing, and reselects the accelerometer and reaction control thrusters for on-orbit nutation control.

Description: BACKGROUND The present invention relates to spacecraft and solid rocket vehicles and, more particularly, to control of disturbance torques during solid rocket burn for a spinning vehicle. It is conventional to use multiple-stage rockets for putting artificial satellites into Earth orbit. Generally, two stages are used for the initial boost phase to achieve a low orbit outside the Earth's atmosphere. A third stage produces a highly elliptical orbit known as the transfer orbit, and a fourth stage is used when a high altitude circular orbit is required. A solid rocket motor in a spinning vehicle configuration is common for the third and fourth stages, at spin rates of 20 to 60 RPM, typically. During firing of a spinning solid rocket motor, the spacecraft is often subjected to disturbance torques that cause a coning motion known as nutation. As long as the nutation is small, it is not objectionable because it can be controlled or damped, and there is no danger of the satellite tumbling. However, if the nutation increases to the range between 8 to 15 degrees, there is cause for alarm. This coning motion is thought to be caused by the motion of the combustion products within the solid rocket motor during firing. One method of controlling this coning is by putting constraints on the manufacturer of the solid rocket motor and on the manufacturer of the spacecraft. The spacecraft mass properties and the solid rocket motor lateral force characteristics can be limited and specified in great detail to prevent or eliminate disturbance torques that cause coning motion. Another method of compensating for the coning motion is to provide dedicated control and propulsion equipment such as steering nozzles integral with the solid rocket motor that are automatically aimed in a direction that compensates for lateral disturbance torques. Another method is to provide a self-contained external bolt-on package that adds dedicated control and propulsion thrusters to the solid rocket motor for compensating lateral disturbance torques automatically. The present invention addresses the problem of nutation during firing of a spinning solid rocket vehicle. However, it reduces or avoids limitations or constraints on spacecraft mass properties and solid rocket motor lateral force characteristics. It does not require the use of dedicated control and propulsion equipment and it is more efficient in weight and cost than the alternative referred to above. The present invention provides a spacecraft on-board control and propulsion system comprising: a solid rocket motor; a first motion sensing device to sense nutation of the spacecraft when the solid rocket motor is firing; a second motion sensing device to sense nutation of the spacecraft when the solid rocket motor is not firing; a first propulsion device to apply thrust to control nutation of the spacecraft when the solid rocket motor is firing; a second propulsion device to apply thrust to control nutation of the spacecraft when the solid rocket motor is not firing; a propulsion control unit to generate propulsion control signals in response to an input from the first or second motion sensing device; and switch means adapted for external control to selectively couple the first motion sensing device to the input of the control unit and the first propulsion device to the output of the control unit when the solid rocket motor is firing and to selectively couple the second motion sensing device to the input of the control unit and the second propulsion device to the output of the control unit when the solid rocket motor is not firing, in order to provide nutation control both when the solid rocket motor is firing and when it is not. In another aspect of the invention, a method of controlling nutation of a spacecraft using an on-board control and propulsion system as above is provided, with steps as defined in claim 9. The switch means may be controlled by a time sequencing unit and by pulse commands from the ground control station. The switch means preferably determines gyro and thruster signal routing for the two states: subsystem configuration enabled (on) and disabled (off). The disabled state is the state that serves all mission requirements other than nutation control during perigee boost solid rocket motor firing and is the state that is operational throughout the life of the satellite. At the time of the perigee boost solid rocket motor firing, the subsystem is enabled, and the gyro signal is routed into the attitude control electronics unit, leaving the accelerometer temporarily disconnected therefrom. Also in this enabled state, the thruster signals are routed to the solenoid drivers for the liquid apogee motors. The various features and advantages of the present invention may be more readily understood with reference to the following detailed description taken in conjunction with the accompanying drawings, wherein like reference numerals designate like structural elements, and in which: FIG. 1 is a diagram illustrating the injection of an Earth satellite into orbit around the Earth; FIG. 2 is a side view of the satellite of FIG. 1, shown broken away and partly in phantom, and illustrating part of the on-board propulsion equipment carried thereby; FIG. 3 is a simplified block diagram of a conventional on-board control and propulsion system for the satellite of FIGS. 1 and 2; and FIG. 4 is a simplified block diagram of a redundant on-board control and propulsion system that has been modified in accordance with the invention. Referring now to the drawings, FIG. 1 is a diagram illustrating the injection of an Earth satellite 10 into a geosynchronous orbit 11. The orbit 11 is equatorial and is 35900 km (22,300 statute miles, or 19,300 nautical miles), above the Earth 12. There are several methods of launching an Earth satellite 10 into the geosynchronous orbit 11. The launch vehicles available include the Thor Delta, Atlas Centaur, Ariane, and Titan III C. When using the Titan III C, for example, the satellite 10 is first boosted to a low orbit 13 a few hundred miles (1 mile = 1.609 km) above the Earth 12. This low orbit 13 is typically a nominal 297 km (160 nautical miles) above the Earth 12 with an orbit inclination of 28.6 degrees when the launch is due East from Kennedy Space Center at Cape Canaveral, Florida. To place the spacecraft 10 in geosynchronous orbit 11, a solid rocket motor is fired at the perigee 14 of a transfer orbit 15, and liquid apogee motors are fired at the apogee 16 of the transfer orbit 15. Referring now to FIG. 2, the satellite 10 typically is provided with a perigee solid rocket kick motor 19 and with on-board control and propulsion equipment that includes at least two liquid apogee motors 20, at least one fuel tank 21, at least one radial spin up reaction control thruster 22. The spin up thruster 22 and spin down thruster 23 have nozzles aligned tangentially transverse to the spin axis, and are pointed in opposite directions. The liquid apogee motors 20, the spin up thruster 22, and the spin down thruster 23 include solenoid valves that are connected by tubing to the fuel tank 21. The satellite 10 is also equipped with an accelerometer 24, a solenoid drive unit 25, and an attitude control electronics unit 30. Referring now to FIG. 3, the satellite 10 is also equipped with a telemetry unit 26, a command receiving unit 27, and a time sequencing unit 28. More than one of each of the foregoing elements may be provided on board the satellite 10 to permit redundant operation which provides for a more reliable satellite 10. It should be understood that the liquid apogee motors 20, the spin up thruster 22, and the spin down thruster 23 are usually not operated during the transfer orbit 15 in conventional systems. However, in accordance with the principles of the present invention, apparatus is provided for reconfiguring the existing on-board control and propulsion equipment to permit automatic nutation control during solid rocket motor firing. That is to say, the liquid apogee motors 20, are operated selectively in a predetermined manner during the perigee kick solid rocket motor burn in the transfer orbit 15. This mode of operation is provided to correct for spin-axis coning caused by disturbance torques produced during the solid rocket motor firing. FIG. 4 is a block diagram showing an exemplary embodiment of a satellite on-board propulsion and control system that has been modified in accordance with the principles of the present invention. The exemplary satellite propulsion and control system has been illustrated in a general manner to show the unit interconnection scheme and the functional relationship between elements. FIG. 4 shows the addition of apparatus for reconfiguring existing on-board control and propulsion equipment for control of spacecraft spin axis orientation during solid rocket motor firing. The added apparatus selects rate sensing elements and one type of propulsion components for control during solid motor firing and reselects other sensing elements and other propulsion components for control during other mission phases. The overall subsystem is referred to as a Perigee Motor Firing Active Mutation Control, or PANC 40. All of the elements of the PANC 40 are redundant to provide reliable operation. The added components are mounted on the spacecraft and comprise first and second switch means 41, 42 and first and second rate sensing gyros 43, 44. The switch means 41, 42 may be solid state multiplexers or banks of relays, or the like. The gyros 43, 44 are mounted in such away that they are adapted to sense the rate of coning motion of the spin axis of the spacecraft. The switch means 41, 42 are shown in the disabled state which is functionally equivalent to the control subsystem in a space vehicle that has not been modified in accordance with the invention. The only other added item is the new wiring necessary to integrate the added components with the existing on-board propulsion and control system. A status line 45 connects the switch means 41, 42 to first and second telemetry units 46, 47 which telemeter status signals to a ground control station. The switch means 41, 42 are adapted to receive an "off" signal via control lines 50 from first and second command receiver units 51, 52 that receive control signals from the control station. The switch means 41, 42 are adapted to receive an "on" signal via control lines 53 from the command receiver units 51, 52 or alternatively, from first and second time sequenced units 54, 55. The PANC 40 includes first and second attitude control electronics units 60, 61 which are conventionally used in satellites such as Intelsat VI, for example, and are not modified in any way. Instead, the electrical harnessing interconnecting the existing units is modified to produce new paths for signal and control lines switched by the switch means 41, 42 and to provide new paths for control and status signals. The attitude control electronics units 60, 61 conventionally receive sensor input signals on first and second accelerometer input lines 62, 63. In an unmodified spacecraft, those input signals would always be provided by first and second accelerometers 64, 65. In the exemplary embodiment of a spacecraft modified in accordance with the present invention, the first accelerometer input line 62 is connected to the output of the first switch means 41. Signals from the first accelerometer 64 and from the first rate sensing gyro 43 are connected to the input of the first switch means 41. When the first switch means 41 is in the "off" or PANC disabled state, the output of the first accelerometer 64 is connected to the input of the first attitude control electronics unit 60. However, when the first switch means 41 is in the "on" or PANC enabled state, the first accelerometer 64 is disconnected, and the first rate sensing gyro 43 is connected to the input of the first attitude control electronics unit 60. Of course, the redundant units, namely, the second attitude control electronics unit 61, the second switch means 42, the second accelerometer 65, and the second rate sensing gyro 44 operate in a corresponding manner. The output control signal from the attitude control electronics unit 60, 61 is applied to the first and second switch means 41, 42 by first and second output control lines 66, 67. First and second solenoid drive units 70, 71 provide control for the entire propulsion and thruster system 72, which comprises first and second liquid apogee motors 91, 92, and first and second radial spin up reaction control thrusters 93, 94, and first and second radial spin down reaction control thrusters 95, 96. The first solenoid drive unit 70 is connected to the propulsion and thruster system 72 by a first solenoid drive bus 73, while the second solenoid drive unit 71 is connected to the propulsion and thruster system 72 by a second solenoid drive bus 74. The first solenoid drive unit 70 has radial thruster input lines 75 connected to one output of the first switch means 41, and has LAM thruster input lines 77 connected to another output of the first switch means 41. Similarly, the second solenoid driver unit 71 has radial thruster input lines 76 connected to one output of the second switch means 42, and has LAM thruster input lines 78 connected to a different output of the second switch means 42. In operation, the PANC 40 permits strong nutation damping during perigee kick motor burn. The switch means 41, 42 determine the routing of input signals and output signals for the attitude control electronics units 60, 61 for two states: PANC "on" or enabled; and PANC "off" or disabled. When the PANC 40 is "on" or enabled, a nutation signal from the gyros 43, 44 reaches the attitude control electronics units 60, 61. These units 60, 61 interpret the nutation signal and sends thruster firing signals to the liquid apogee motors 91, 92. With the PANC 40 in the "on" or enabled state, the signal from the gyros 43, 44 is routed into the attitude control electronics units 60, 61 via the accelerometer input lines 62, 63, leaving the accelerometers 64, 65 temporarily disconnected from the attitude control electronics units 60, 61. Also in the "on" or enabled state, the radial spin thruster control signals appearing at the output control lines 66, 67 of the attitude control electronics units 60, 61 are routed to the liquid apogee motor inputs to the solenoid drive units 70, 71 causing firing of the liquid apogee motors 91, 92. The input signal and output signal routings permit the attitude control electronics units 60, 61 to serve the PANC function even though the stock attitude control electronics units 60, 61 are originally designed to receive input signals from the accelerometers 64, 65, and are designed to fire the reaction control thrusters 93, 94, 95, 96 for on-orbit nutation control. In the "off" or PANC disabled state, the signals from the accelerometers 64, 65 are connected to the input of the attitude control electronics units 60, 61, and the gyros 43, 44 are disconnected. In the "off" or PANC disabled state the radial spin thrusters control signals are connected to the radial spin thrusters inputs of the solenoid drive units 70, 71. Accordingly, the control of the radial thrusters 93, 94, 95, 96 and the liquid apogee motors 91, 92 is restored to the original configuration. Hence, the PANC 40 when in the "off" or disabled state is functionally equivalent to the control ad propulsion system in a space vehicle which has not been modified in accordance with the invention. Although not indicated in FIG. 4, the switch means 41, 42 also control the on/off state of the electronics and mechanism in the gyros 43, 44. In addition, the gyros 43, 44 are permanently disconnected from the power bus upon perigee stage operation. Switch means in the perigee motor firing active nutation control subsystem determine gyro and thruster signal routing for two states: subsystem configuration enabled (on) and disabled (off). The disabled state of the perigee motor firing active nutation control system constitutes a satellite configuration that serves all mission requirements other than nutation control during perigee motor firing. This is the state that is selected throughout the operational life of the satellite. This state makes the satellite functionally equivalent with other satellites of the same series which lack the modified subsystem of the present invention. In this state an accelerometer signal is processed in an attitude control electronics unit, which determines and generates reaction control thruster firing control signals as necessary to damp any measured nutation exceeding a given threshold. At the time of the perigee motor firing, the attitude control electronics unit is internally configured to send thruster control signals to the reaction control thrusters. Neither the accelerometer, as a sensing device, nor the reaction control thrusters, as actuator devices, are adequate to implement the desired control during the perigee motor firing. When the subsystem is enabled, the gyro signal is routed into the attitude control electronics unit, leaving the accelerometer temporarily disconnected therefrom. Also in this enabled state, the reaction control thruster outputs are routed to the thruster valve drive unit solenoid drivers for the liquid apogee motors. These signal routings permit the attitude control electronics unit to serve the perigee motor firing active nutation control subsystem function even though it is designed to receive as an input the accelerator signal and use the reaction control thrusters for on-orbit nutation control. It is to be understood that the above described embodiment is merely illustrative of some of the many specific embodiments which represent applications of the principles of the present invention. Clearly, numerous and other arrangements can be readily devised by those skilled in the art without departing from the scope of the invention as defined by the accompanying claims.

Claims: A spacecraft on-board control and propulsion system (40) comprising: a solid rocket motor (19); a first motion sensing device (43; 44) to sense nutation of the spacecraft when the solid rocket motor (19) is firing; a second motion sensing device (64; 65) to sense nutation of the spacecraft when the solid rocket motor (19) is not firing; a first propulsion device (91; 92) to apply thrust to control nutation of the spacecraft when the solid rocket motor (19) is firing; a second propulsion device (93, 95; 94, 96) to apply thrust to control nutation of the spacecraft when the solid rocket motor (19) is not firing; a propulsion control unit (60; 61) to generate propulsion control signals in response to an input from the first or second motion sensing device; and switch means (41; 42) adapted for external control to selectively couple the first motion sensing device (43; 44) to the input of the control unit (60; 61) and the first propulsion device (91; 92) to the output of the control unit when the solid rocket motor (19) is firing and to selectively couple the second motion sensing device (64; 65) to the input of the control unit (60; 61) and the second propulsion device (93, 95; 94, 96) to the output of the control unit when the solid rocket motor (19) is not firing, in order to provide nutation control both when the solid rocket motor (19) is firing and when it is not. The system of claim 1 in which said switch means (41; 42) comprises a relay. The system of claim 1 in which said switch means (41; 42) comprises a solid state multiplexer. The system of claim 1 or claim 2 in which said first motion sensing device (43; 44) comprises a rate sensing electromechanical device. The system of claim 4 in which said electromechanical device (43; 44) comprises a rate sensing gyro. The system of any of the above claims in which said first propulsion device (91; 92) comprises a liquid apogee motor. The system of any of the above claims in which said second motion sensing device (64; 65) comprises an accelerometer. The system of any of the above claims in which said second propulsion device (93, 95; 94, 96) comprises radial spin thrusters. A method of controlling spacecraft nutation using a spacecraft on-board control and propulsion system according to any preceding claim comprising the following steps: sensing nutation of the spacecraft spin axis during solid rocket motor burn by means of the first motion sensing device (43; 44), the output of said first motion sensing device (43; 44) being connected to the input of the propulsion control unit (60; 61); applying thrust to the spacecraft from the first propulsion device (91; 92) in response to control signals generated by the propulsion control unit (60; 61) during said solid rocket motor burn to control nutation during solid rocket motor burn; and switching the input of the propulsion control unit (60; 61) from the output of the first motion sensing device (43; 44) to the output of the second motion sensing device (64; 65) and the output of the propulsion control unit (60; 61) from the first propulsion means to the second propulsion means to control nutation when the solid rocket motor is not firing. The method of claim 9 wherein the first motion sensing device is a rate sensing gyro, the second motion sensing device is an accelerometer, the first propulsion device is a liquid apogee motor, the second propulsion device is radial spin thrusters, and the step of applying thrust comprises applying thrust to the spacecraft from the first propulsion device (91; 92) in response to nutation above a predetermined threshold during said solid rocket motor burn.

Inventors: ALTOBELLI MARK R, ALTOBELLI, MARK R.
Applicants: HUGHES AIRCRAFT CO, HUGHES AIRCRAFT COMPANY
Application year: 1991
Main classification: B64G1
IPC codes: B64G1, G05D1