Stabilization of a spinning spacecraft of arbitrary shape

Abstract: A stabilization system (40) for a spacecraft which enables the spacecraft to be stabilized by rotation about any principal moment of inertia axis (x, y). The system includes two pairs of control thrusters (14, 16; 18, 20) for producing positive and negative control moments about two orthogonal axes (x, y) mutually perpendicular to the axis of rotation of the spacecraft. The control system (40) includes rate gyros (42, 46) which provide error signals which are sent to duty cycle modulators (54). The duty cycle modulators (54) provide a pulse width modulated control signal to the control thrusters (14 - 20) to approximate proportional control of spacecraft attitude.

Description: The present invention relates to a stabilization system for a spinning spacecraft providing rotational stability about an axis of rotation, including: measuring means for providing error signals relating to first and second orthogonal axes mutually perpendicular to said axis of rotation, said measuring means detecting the angular velocities of motion about said first and second axes, control thrusters providing positive and negative moments about said first and said second axis, respectively, wherein said spacecraft has a main thruster for orbit transfer, control means for firing said control thrusters in response to said error signals, whereby said spacecraft is maintained in stability. Such a stabilization system is known from document EP-A-0 225 683. The known system is a PWPF (pulse width pulse frequency)-based system which uses control thrusters arranged opposite to the main thruster in the spin-up phase. This invention relates thus to a system for stabilizing a spinning spacecraft exposed to disturbing moments. Spin stabilization is the simplest technique for controlling spacecraft attitude for spacecraft orbited around the earth. A spacecraft can be characterized as having three orthogonal moment of inertia axes which typically each have a unique moment of inertia value, thus defining a principal or maximum axis, a minimum axis, and an intermediate axis. Rotation about any axis presents stability concerns which are alleviated by the present invention. As an alternative to rotating a spacecraft about an unstable axis or when subjected to destabilizing forces, a three-axis active control system can be used. These systems are complex, and are subject to gyro drift error which can lead to attitude inaccuracies. Present nutation control systems for spinning spacecraft typically require nutation to reach a predetermined level before corrective thruster firing occurs and thus does not provide proportional attitude control. In some cases, this can cause loss of control. Further, a spin-stabilized spacecraft similar to that cited at the outset is disclosed in FR-A-2 373 823. Stabilization of rotational movement is effected by control thrusters arranged in parallel to the main thruster. Integrating gyros measure the angular deviation. An earth station transmits control signals used to adjust the gyros. In view of this it is an object of the present invention to provide an improved active stability control system. According to the system mentioned at the outset this object is achieved in that said measuring means provide error signals proportional to said angular velocities, and each of said control thrusters produces thrust acting in the direction of thrust of said main thruster. In accordance with this invention, an improved active stability control system is provided for spacecraft of various configurations including oblate, intermediate or prolate body forms . The system enables rotation about any of the principal moment of inertia axes of the spacecraft. Since the control system according to this invention requires a small expenditure of propellants, it is primarily useful for the transfer orbit phase of a mission rather than the continuous on-station phase. The present system provides essentially a two-axis control used with a spinning spacecraft. The system employs error signals which fire thrusters which are duty cycle modulated to provide the equivalence of proportional control over spacecraft attitude. Additional benefits and advantages of the present invention will become apparent to those skilled in the art to which this invention relates from the subsequent description of the preferred embodiments and the appended claims, taken in conjunction with the accompanying drawings. FIG. 1 is a pictorial diagram of a spacecraft including the active stabilization system of this invention with control thrusters positioned according to a first embodiment of this invention. FIG. 2 is a schematic diagram of a spacecraft like FIG. 1 but showing control thrusters positioned according to a second embodiment of this invention. FIG. 3 is a block diagram showing the control system for the stabilization system according to the present invention. FIG. 4 is a graph showing the output of the duty cycle modulators used with the present invention. The physical elements of the system according to the present invention will be described with reference to the Figures, which will be followed by a description of the theory of operation of the system. FIG. 1 illustrates a spacecraft 10 represented graphically having main thruster 12 provided for boosting the spacecraft to the desired orbital position. Orthogonal axes are shown with spacecraft 10 rotating about the z-axis. A number of control thrusters are shown which are controlled in accordance with this invention to maintain stability of rotation about the z-axis, including plus and minus x-axis thrusters 14 and 16, respectively; and plus and minus y-axis thrusters 18 and 20, respectively. As their description suggests, thrusters 14 through 20 produce thrust along vectors displaced from the spacecraft center of mass (C.M.) thus producing positive and negative moments about their respective axes. FIG. 2 shows an alternate embodiment of the present invention which is identical to the first described embodiment with the exception that spacecraft 24 has thrusters 26, 28, 30 and 32, which are moved to the outer corners of the spacecraft. Thrusters 26 through 32 are not located on the axes (unless the axes are rotated) and, as shown in the drawing, produce moments which are resolved along two axes. The second embodiment provides advantages in that the thruster outputs provide a greater control moment than the embodiment depicted by FIG. 1 due to the greater distance of their thrust vectors from the center of mass. FIG. 3 provides a graphical illustration of stabilization control system 40 in accordance with this invention. A pair of angular rate gyros 42 and 44 measure the transverse angular velocity (ω) about the x and y axes, respectively. The outputs from gyros 42 and 44 are separated into positive error channels 46 and 48, and negative error channels 50 and 52. Duty cycle modulators 54 provide a pulse-width modulated control signal having a duty cycle proportional to the magnitude of error signals along channels 46 through 52. This modulation provides the equivalence of proportional control moments by control thrusters even though such thrusters typically have only on or off states. Preferably, the frequency of control modulation is at least several times the spin frequency of spacecraft 10 in order not to excite pendulous motion of the fuel in internal tanks. A typical pulse-train output from a modulator 54 is shown in FIG. 4 which shows the duty cycle modulation output. FIG. 4 shows a total pulse period of T with a duration of t and a frequency of four times the spin period of spacecraft 10. Duty cycle modulators 54 directly control thrusters 14 through 20 as shown in the figure to provide moments which stabilize spacecraft 10. The following will discuss the theories of operation for control system 40. Let the moments of inertia of a spacecraft 10 about the three principal axes x, y, and z be called I x , I y , and I z , respectively, with z being the spin axis. Let I Z /I X be σ x and I z /I y be σ y . If both σ x and σ y are much greater than unity, spacecraft 10 is unconditionally stable. For values of σ x and σ y close to but greater than unity, fuel motion within spacecraft 10 can exacerbate wobble induced by cross products of inertia or by thrust misalignment of main thruster 12 and for some conditions result in a runaway amplification of the wobble angle. For values of σ x and σ y less than unity, nutational instability is caused by de-dampers on the spacecraft 10 unless some means of stabilization is employed. For spin about an intermediate moment inertia axis, such as occurs if σ x is greater than unity while σ y is less than unity, or vice versa, a rapid divergence of wobble angle occurs unless an effective means of stabilization is employed such as the system according to this invention. Euler's moment equations can be expressed as follows: sω x + (σ y -1) I y I x ω z ω y = L x I y - (σ x -1) I x I y ω z ω x + sω y = L y I y Let L x I y = M x - Kω x and L y I y = M y - Kω y then ω x = S+K M x - σ y -1 I y I x ω x M y S+K + σ x -1 σ y -1 ω z 2 ω y = S+K M y - σ x -1 I x I y ω z M x S+K + σ x -1 σ y -1 ω z 2 where ω x , ω y , and ω z are the angular velocities about the x, y, and z axes, respectively, and L x and L y are the external moments about the x and y axes, and M x , M y are the normalized external moments about the x and y axes, thus M x = L x I x M y = L y I y K is the feedback gain supplied by control system 40, and s is Laplace's differential operator. As discussed above, control system 40 measures the transverse angular rates of spacecraft 10 with rate gyros 42 and 44 and creates proportional control moments by duty cycle modulation of appropriate control thrusters. The behavior of the system with and without this feedback is explained by Euler's moment equations. When the feedback factor K is zero, i.e., control system 40 is inoperative, the determinant is s² + (σ x -1)(σ y -1) ω z 2 and the roots of the characteristic equation resulting from equating the determinant to zero are: s = ± jω z σ x -1 σ y -1 The behavior of spacecraft 10 can be considered for three cases. When both σ x and σ y are greater than unity (first case), the roots are imaginary when no energy dissipation is present. Energy dissipation causes the roots to have a negative component and spacecraft 10 is dynamically stable. Fuel displacement introduced by a small dynamic unbalance, however, causes further unbalance and as the values of either σ x or σ y approach unity, the situation can become statically unstable. When both σ x and σ y are less than unity (second case), the roots are also imaginary for the case of no energy dissipation. Energy dissipation causes the roots to have a positive component, however, and spacecraft 10 is unstable unless stabilized by other means. When σ x is greater than unity and σ y is less than unity (third case), or vice versa, one root is negative and the other positive. The wobble diverges rapidly unless some means is used to stabilize it. Now consider the behavior of the system when the thruster feedback provided by control system 40 is used to stabilize spacecraft 10. The roots of the characteristic equation for the first two cases are: s = -K ± jω z σ x -1 σ y -1 The feedback constant K causes the roots to have a negative component K, which can be as large as needed to overcome the instabilities associated with those cases. For the third case, the stability criterion (no roots with positive components) is: K 2 >(σ x -1)(σ y -1)ω z 2 which is easily met. Thus, control system 40 can be used to stabilize spacecraft 10 for all ratios of moments of inertia.

Claims: Stabilization system for a spinning spacecraft (10; 24) providing rotational stability about an axis (z) of rotation including: measuring means (42-52) for providing error signals relating to first and second orthogonal axes (x, y) mutually perpendicular to said axis of rotation (z), said measuring means (42-52) detecting the angular velocities of motion about said first and second axes (x, y), control thrusters (14-20; 26-32) providing positive and negative moments about said first and said second axis (x, y), respectively, wherein said spacecraft (10; 24) has a main thruster (12) for orbit transfer, control means (54) for firing said control thrusters (14-20; 26-32) in response to said error signals, whereby said spacecraft (10; 24) is maintained in stability; characterized in that said measuring means (42-52) provide error signals proportional to said angular velocities, and each of said control thrusters (14, 16; 18, 20) produces thrust acting in the direction of thrust of said main thruster (12). Stabilization system according to claim 1, characterized in that said control means (54) provide pulsed firing signals to said control thrusters (14-20; 26-32) to provide pulsed thrust outputs. Stabilization system according to claim 2, characterized in that said control means (54) adjusts the duty cycle of said firing signals to increase said duty cycle in response to increased magnitude of said error signals, thereby approximating proportional control of the attitude of said spacecraft (10; 24). Stabilization system according to any of the preceding claims, characterized in that said axis of rotation (z) is about a maximum, an intermediate or a minimum moment of inertia axis of said spacecraft (10; 24). Stabilization system according to any of the preceding claims, characterized in that said control thrusters (14-20; 26-32) each produce thrust in directions parallel to said axis of rotation (z). Stabilization system according to any of the preceding claims, characterized in that said measuring means (42-52) comprises a pair of rate gyros (42, 44) for measuring angular velocities of said spacecraft (10; 24) about said first and second axes (x, y) and duty cycle modulators (54) receiving signals from said rate gyros (42, 44) for providing pulse width modulated control signals for said control thrusters (14-20; 26-32). Stabilization system according to any of the preceding claims, characterized in that said measuring means includes means (46-52) for generating positive and negative error signals and said control means (54) includes means for directing said error signals to corresponding ones of said control thrusters (14-20; 26-32). Stabilization system according to claim 7, characterized in that said negative error signals are applied to said control thrusters (16, 20) for providing negative moments by a first pair of duty cycle modulators (54) and said positive error signals are applied to said control thrusters (14, 18) for providing positive moments by a second pair of duty cycle modulators (54). Stabilization system according to any of the preceding claims, characterized in that said control thrusters comprise a first pair of control thrusters (14, 16) for providing positive and negative moments about said first axis (x), a second pair of control thrusters (18, 20) for providing positive and negative moments about said second axis (y), and that said control means (54) fire one of said first pair of control thrusters (14, 16) in response to said error signals about said first axis (x) and fire one of said second pair of control thrusters (18, 20) in response to said error signals about said second axis (y). Stabilization system according to any of claims 1 to 8, characterized in that said control thrusters (26-32) are arranged offset from said first and second orthogonal axes (x, y) by 45°.

Inventors: ROSEN HAROLD A, ROSEN, HAROLD A.
Applicants: HUGHES AIRCRAFT CO, HUGHES AIRCRAFT COMPANY
Application year: 1989
Main classification: G05D1
IPC codes: G05D1, B64G1