Attitude Control Of Space Vehicles Research Articles

Computation of program controls performed nonstop by gyrodynes

A new method for calculating space vehicle (SV) attitude controls ensuring their effective implementation by a system of collinear pairs of single-gimbal forced unrestrained gyros (gyrodynes) has been proposed. The novelty of the method consists in a virtual kinematic configuration of the gyro system, i.e., the precession of gyro units in the collinear pairs of gyrodynes is coupled in a nonmechanical manner. In addition, the angular momentum of the system as a state variable for describing the dynamics of the SV permanent rotation was used for the first time at the stage of computing controls performed nonstop by gyrodynes. In the general formulation, when the desired final state of the SV is arbitrary, the SV attitude control problem can be reduced to a sequence of permanent rotations. The performance of the method is demonstrated as applied to the calculation of program gyrodyne controls with a permanent reduction in the SV angular velocity around its center of mass with a nonzero SV angular momentum after its discharge.

Integrated Method - the Optimum Way to Improve the Quality of Frequency Response Characteristics of the Space Vehicle Attitude Control System

The integrated method applied to the design of technical systems is a process in which various project, calculation and verification procedures are interconnected and interrelated. The results of procedures are used in a certain sequence, thus ensuring maximum reachable optimality of the system being designed.

Adaptive Dynamic Sliding Mode Control for Near Space Vehicles Under Actuator Faults

A novel adaptive dynamic sliding mode (ADSM) fault-tolerant control (FTC) methodology is developed for near space vehicle attitude control systems with actuator faults in this paper. The proposed ADSM approach combines dynamic sliding mode with adaptive control strategies that can make the systems stable and accurately track the desired signals in the presence of external disturbances, model parameter uncertainties, and even actuator faults. Firstly, the attitude dynamic model of X-33 and its faulty model are introduced, then the ADSM control and fault-tolerant control laws are designed for outer-loop and inner-loop, respectively. Finally, in comparison with one existing approach, the simulation results are provided to show the effectiveness of the proposed FTC scheme.

Controllability and observability of gyroelastic vehicles

Stored angular momentum devices such as reaction wheels and control moment gyros have been used extensively for space vehicle attitude control. They represent a potential source of actuation for vibration and shape control of large space structures where they can potentially be distributed in large numbers. The vibration characteristics of these gyroelastic vehicles are affected by the presence of stored angular momentum and, hence, so are the conditions for controllability and observability. In this paper, these conditions are derived for systems modeled as gyroelastic continua, i.e., vehicles with continuous distributions of mass, stiffness, and gyricity (stored angular momentum). The conditions are expressed in terms of the gyroelastic modes and cover the case of pointwise actuators and those modeled in a continuum fashion. A numerical example is used to show that the degree of controllability in the continuum case can be interpreted as that corresponding to the limit of a sequence of pointwise control problems. The observability conditions are developed for a general class of measurements. The concept of a gyroelastic node is introduced and related to the problem of locating sensors.

Flexible spacecraft attitude control using a simple P + D algorithm

Two philosophies for the attitude control of space vehicles with flexible appendages have appeared in the literature, these are: (i)control of the rigid centre body to lie within some bounded region while the flexural vibrations are allowed to behave in an uncontrolled manner;(ii)control of all plant states to some finite value as time tends to infinity.The former method does not require flexural information and therefore avoids the on-board computational burden associated with state estimation, or the cost of additional instrumentation. However, attitude degradation has been observed and reported indicating a clear need to compensate for structural vibrations.

Pulse Ratio Modulation: An Interesting Technique to Implement DC/DC Conversion

DC/DC converters may be considered as pulse modulators; the modulation laws usually implemented are those related to pulse frequency (PFM) and pulse width modulation (PWM) techniques. The pulse ratio modulation (PRM) technique, providing several advantages over conventional pulse modulation methods, has been developed for applications in a space vehicle attitude control system. Capable of operation with a high degree of precision and overcoming certain bandwidth problems associated with other modulation techniques, the modulator appears to hold excellent possibilities for applications in a broad range of control systems, such as DC and AC drives.In order to investigate the dynamical properties presented by a feedback control system employing this modulation technique, in the paper a mathematical model is presented, characterizing the PRM in terms of input output relationships. By using the state variables approach, the system is described through a vector recurrence equation and . linear bounds on the state characterizing the adopted modulation law. On the basis of this model both the steady state behaviour and the transient response are evaluated. As far as stability properties are concerned, the linearized model is developed in order to determine the local behaviour in correspondence with the steady state solutions, and a comparison is made with an equivalent PWM control system. Moreover a procedure, based on the use of projection operators in vector spaces, is presented to analyse the asymptotic stability in the large.

Optimum propellant usage for reaction jet systems of space vehicles

The on-off type control for reaction jet systems is proven to be the optimal fuel scheme. However, due to the nonlinear characteristics of this type control, no direct method of solution is known for this optimal process being applied to the attitude control of space vehicles. This paper will discuss the optimization of fuel usage for an attitude control of space vehicles. A computation technique is developed for the calculation of optimal control law. The method of calculus of variations is applied to the estimation of the changes of performance index as well as terminal constraints. Thus an algorithm is obtained by the steepest descent method. A numerical example is given in the paper.

A new method for performing digital control system attitude computations using quaternions

Space vehicle attitude control system performance is often limited by the computational speed of airborne digital computer hardware. The application of quaternions to digital atti- tude control problems involving coordinate system transformations can reduce computation time by more than 40% over the equivalent direction cosine matrix solution due to a new method of treating quaternions which allows the order of multiplication to be interchanged so as to isolate the most rapidly varying parameters. This paper presents a formulation for control equations in which the control error is expressed as a and all coordinate system transformations are performed using quaternions. The new principle of quaternion algebra which allows interchanging the order of multiplicatio n is developed and used to simplify the control equations. The control equations are applied to a three- gimbal IMU example and the results are compared to an equivalent direction cosine matrix solution. T HE problem of describing the relationship between two coordinate systems is one of the most basic concepts en- countered in the field of navigation and guidance. Until re- cently Euler angles formed the most widely used method of describing the rotation between two coordinate systems. The use of three Euler angles to fix the attitude of a body with re- spect to an inertial or reference coordinate system has the advantage of being well-defined geometrically and fairly simple to visualize. The 3X3 direction cosine matrix, how- ever, was more readily adaptable to high-speed digital com- putation and has replaced the Euler angle method in the solution of all but the most simple navigation problems. The direction cosine matrix approach is particularly useful to de- scribe several successive rotations of a body with respect to a fixed reference system. A third, but infrequently applied, approach to establishing body orientation utilizes the qua- ternion, first devised by Hamilton. The ap- proach makes use of Euler's Theorem which states that any real rotation of one coordinate system with respect to another may be described by a rotation through some angle about a single fixed axis. The is a compact form for rep- resenting the single fixed axis and angle referred to by Euler's Theorem. The may be handled much the same as the direction cosine matrix, in that successive rotations result in successive multiplication. The ad- vantage of the lies in its ability to define the ro- tational relationship between two coordinate systems using only four numbers as opposed to the nine elements of a direc- tion cosine matrix. This results in a similar simplification when the effect of several successive rotations is being com- puted. The principal impediment to use of quaternions has been that the direction cosine matrix, rather than the qua- ternion, is the desired end product of computation. Usually, the computation saved by using quaterion multiplication to perform successive coordinate system rotations is lost when the resultant must be put in the form of a direction cosine matrix. When the object of the computation is to obtain the control error it is not necessary to define the direction cosine matrix explicitly. In fact, optimum control, in the least angle sense, is performed by defining the control error in terms of the single axis and angle of Euler's Theorem. This may be done by formulating the problem completely in terms of quater- nions. The formulation is shown to be particu- larly advantageous due to a new method for treating quater- nions which allows the order of multiplicatio n to be inter- changed.

Magnetic components for the attitude control of space vehicles

For attitude stabilization of earth satellites, frequently, magnetic systems are used. A main part of such systems are magnetic torquers that, together with the geomagnetic field, produce moments to control the attitude or angular velocity of the satellite. Magnetic torquers consist of permanent magnets or of coils. The coils may be air cored or iron cored. A compromise between permanent magnets and coils is given by chargeable magnets (semi-permanent magnets), i.e., coils with a core consisting of a material with high remanence and medium coercive force. To damp angular oscillations and to remove spin rates, permeable rods are used which undergo magnetization changes in the geomagnetic field, thus transforming the kinetic energy of the satellite into hysteresis and eddy-current losses. Other damping devices consist of a secondary mass dissipating energy by moving relative to the satellite. Extensive testing is necessary to assure proper functioning and reliability of the devices described.

Performance characteristics of a ''damped'' mass expulsion attitude control system.

A new mass expulsion approach to space vehicle attitude control is introduced which shows promise of reducing propellant consumption and pulsing frequency plus providing better acquisition performance (time and propellant) and attitude accuracy than present minimumimpulse controllers. This attitude control logic is designed to adapt optimally to existing disturbance torques in damping the vehicle's position and rate, thereby driving the vehicle near the minimum propellant expenditure trajectory during each control cycle. Simple adaptive compensation based on the immediate past history of control pulses is included to correct for biased and time-variant disturbance torques, propulsive and sensing errors, and time-variant moments of inertia. The limit-cycle performance of the damped controller is compared with a pseudo-rate and a minimum-impulse controller in the presence of propulsive errors, sensing errors, and external disturbance torques. Results of a threeaxis dynamic simulation indicate that the damped controller uses only one-half of the propellant and only one-fifth of the pulses used by a minimum-impulse controller for a typical low-altitude, elliptical orbit.

Strapdown star tracker for space vehicle attitude control.

Strapdown star tracker for space vehicle attitude control, using scanning and error-signal determination method

Fuel minimization in flight vehicle attitude control

The problem of designing a single-axis rigid-body flight vehicle attitude controller to minimize control fuel consumption is formulated in the framework of system optimization theory. A performance measure which consists of the integral-square position and rate errors plus control fuel expended, with arbitrary relative weighting factors, is employed. The Pontryagin Maximum Principle is invoked to show that the controller which minimizes this measure is of the bang-off-bang type. The calculations are carried through and an explicit design solution given for the problem of a minimum-fuel space-vehicle attitude control.

Recent Advances in Attitude Control of Space Vehicles

Recent Advances in Attitude Control of Space Vehicles

Meteoric Effects on Attitude Control of Space Vehicles

Meteoric Effects on Attitude Control of Space Vehicles

Flywheel control of space vehicles

An analytic study of flywheel autopilots for attitude control of space vehicles is described, and the results of a three-axis analog computer program are presented. The objectives of this study were to determine the requirements, capabilities and limitations of flywheel autopilots as functions of desired accuracy and speed of response, disturbing moments, differential gravity restoring torques, component uncertainties, and vehicle initial attitude and attitude rate errors. The stability and accuracy degradations that resulted from not compensating for gyroscopic crosscoupling torques are summarized. Design tools are presented for synthesizing a three-axis autopilot in accordance with specified design criteria of size, power, accuracy, response rate and error disturbances.

Attitude control of space vehicles

The problem of sensing and controlling the angular orientation of a space vehicle is discussed. Optical, radio, and inertial methods of angle-sensing are described, and application of control torques by reaction jets, flywheels, radiation pressure, and gravitational gradient are compared. Requirements for angular accuracy likely to be imposed by maneuvering, performing measurements for navigation, making scientific measurements, radio communication, and solar cell orientation are described. Attitude control requirements and techniques for earth satellites and deep-space probes are compared. This discussion will not be published.