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Results for “SPACECRAFT STABILITY”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Effectiveness of large booms as nutation dampers for spin stabilized spacecraft

The issue of using long slender booms as pendulous nutation damping devices on spinning aircraft is discussed. Motivation comes from experience with the Galileo Spacecraft, whose magnetometer boom also serves as a passive nutation damper for the spacecraft. Performance analysis of a spacecraft system equipped with such systems are relatively insensitive to changes in the damping constant of the device. However, the size and arrangement of such a damper raises important questions concerning spacecraft stability in general.

Eke, F. O.↗
NASA NTRS1977

A charging model for three-axis stabilized spacecraft

A charging model was developed for geosynchronous, three-axis stabilized spacecraft when under the influence of a geomagnetic substorm. The differential charging potentials between the thermally coated or blanketed outer surfaces and metallic structure of a spacecraft were determined when the spacecraft was immersed in a dense plasma cloud of energetic particles. The spacecraft-to-environment interaction was determined by representing the charged particle environment by equivalent current source forcing functions and by representing the spacecraft by its electrically equivalent circuit with respect to the plasma charging phenomenon. The charging model included a sun/earth/spacecraft orbit model that simulated the sum illumination conditions of the spacecraft outer surfaces throughout the orbital flight on a diurnal as well as a seasonal basis. Transient and steady-state numerical results for a three-axis stabilized spacecraft are presented.

Massaro, M. J.↗
NASA NTRS1976

The motion and stability of spin stabilized spacecraft with hinged appendages

The dynamics and stability of a spin stabilized spacecraft with a hinged appendage system are treated analytically and numerically. The hinged system consists of a central hub with masses attached to massless booms of fixed length whose orientation relative to the main part can change. The equations of motion for the hinged system, with viscous damping at both hinge points, are linearized about the nominal equilibrium position where the booms are orthogonal to the nominal spin axis. Analytic stability criteria indicate that hinge damping must be present and that for limiting cases, where the spin axis is an axis of symmetry, certain conditions relating the hinge point offset coordinates to the moment of inertia ratio and end masses must be satisfied. The hinge damping is always required for the nominal deployment of hinge members.

Sellappan, R.↗
NASA NTRS1995

Telemetry down-link Doppler as an attitude sensor for spin stabilized spacecraft

The communications antenna on a spin stabilized spacecraft is seldom located on the spin axis, hence, the antenna is in motion relative to the center of mass of the spacecraft. The Doppler shift observed at the ground or space relay communications receivers will include oscillations whose frequency and amplitude are functions of the motion of the antenna and the attitude of the spacecraft relative to the line of sight (LOS). This functional dependence creates the possibility of estimating attitude parameters from Doppler measurements. This paper presents mathematical models of Doppler oscillations from spinning spacecraft, including the effects of nutation. Algorithms for estimating spin rate, attitude and nutation angle are described. Results of analysis of Doppler tracking of GOES-8 and WIND are also discussed.

Hendry, S. D.↗
NASA NTRS1976

The dynamics of spin stabilized spacecraft with movable appendages, part 2

The dynamics and stability of a spin stabilized spacecraft with a hinged appendage system are treated analytically and numerically. The hinged system consists of a central hub with masses attached to (assumed) massless booms of fixed length whose orientation relative to the main part can change. The general three dimensional deployment dynamics of such a hinged system is considered without any restriction on the location of the hinge points. The equations of motion for the hinged system, with viscous damping at both hinge points, are linearized about the nominal equilibrium position where the booms are orthogonal to the nominal spin axis for the case of two dimensional and three dimensional motion. Analytic stability criteria are obtained from the necessary condition on the sign of all the coefficients in the system characteristic equation.

Bainum, P. M.↗
NASA NTRS1975

The dynamics of spin stabilized spacecraft with movable appendages, part 1

The motion and stability of spin stabilized spacecraft with movable external appendages are treated both analytically and numerically. The two basic types of appendages considered are: (1) a telescoping type of varying length and (2) a hinged type of fixed length whose orientation with respect to the main part of the spacecraft can vary. Two classes of telescoping appendages are considered: (a) where an end mass is mounted at the end of an (assumed) massless boom; and (b) where the appendage is assumed to consist of a uniformly distributed homogeneous mass throughout its length. For the telescoping system Eulerian equations of motion are developed. During all deployment sequences it is assumed that the transverse component of angular momentum is much smaller than the component along the major spin axis. Closed form analytical solutions for the time response of the transverse components of angular velocities are obtained when the spacecraft hub has a nearly spherical mass distribution.

Bainum, P. M.↗
NASA NTRS1980

Maneuver reconstruction techniques for open-loop spin-stabilized spacecraft

The Pioneer missions were supported by spin-stabilized spacecraft designs using open-loop control and blow-down propulsion subsystems. Reliable estimates of the ever-changing performance inherent to these subsystems were needed to effectively design and reconstruct trajectory correction maneuver (TCM) strategies. These performance updates were obtained by adjusting model parameters to match independent telemetric and radiometric observations to define the simultaneous changes in attitude, velocity, and spin rate during a maneuver sequence.

Frauenholz, R. B.↗
NASA NTRS1975

Four methods of attitude determination for spin-stabilized spacecraft with applications and comparative results

The attitude of a spacecraft is determined by specifying independent parameters which relate the spacecraft axes to an inertial coordinate system. Sensors which measure angles between spin axis and other vectors directed to objects or fields external to the spacecraft are discussed. For the spin-stabilized spacecraft considered, the spin axis is constant over at least an orbit, but separate solutions based on sensor angle measurements are different due to propagation of errors. Sensor-angle solution methods are described which minimize the propagated errors by making use of least squares techniques over many sensor angle measurements and by solving explicitly (in closed form) for the spin axis coordinates. These methods are compared with star observation solutions to determine if satisfactory accuracy is obtained by each method.

Smith, G. A.↗
NASA NTRS1976

Optimal control of spin stabilized spacecraft with telescoping appendages

The control of a spin-stabilized spacecraft consisting of a rigid central hub and one or two movable offset telescoping booms (with end masses) is considered. The equations of rotational motion are linearized about either of two desired final states. A control law for the boom and mass position is sought such that a quadratic cost functional involving the weighted components of angular velocity plus the control is minimized when the final time is unspecified and involves the solution of the matrix Riccati algebraic equation. For three-axis control more than one offset boom (orthogonal to each other) is required. For two-axis control with a single boom offset from a symmetrical hub, an analytic solution is obtained; when this system is used for nutation decay the time constant is one order of magnitude smaller than previously achieved using non-optimal control logic. For the general case results are obtained numerically.

Bainum, P. M.↗
NASA NTRS1976

Optimal control of spin-stabilized spacecraft with telescoping appendages

The control of a spin-stabilized spacecraft consisting of a rigid central hub and one or two movable offset telescoping booms (with end masses) is considered. The equations of rotational motion are linearized about either of two desired final states. A control law for the boom end mass position is sought such that a quadratic cost functional involving the weighted components of angular velocity plus the control is minimized when the final time is unspecified and involves the solution of the matrix Riccati algebraic equation. For three axis control more than one offset boom (orthogonal to each other) is required. For two-axis control with a single boom offset from a symmetrical hub, an analytic solution is obtained; when this system is used for nutation decay the time constant is one order of magnitude smaller than previously achieved using nonoptimal control logic. For the general case results are obtained numerically.

Bainum, P. M.↗
NASA NTRS1976

Dynamics of spin stabilized spacecraft during deployment of telescoping appendages

The dynamics of a spin stabilized spacecraft with telescoping appendages are treated both analytically and numerically. Two types of telescoping appendages are considered: (a) where an end mass is mounted at the end of an (assumed) massless boom; and (b) where the appendage is assumed to consist of a uniformly distributed homogeneous mass throughout its length. For the telescoping system Eulerian equations of motion are developed. During all deployment sequences it is assumed that the transverse component of angular momentum is much smaller than the component along the major spin axis. Closed form analytical solutions for the time response of the transverse components of angular velocities are obtained when the spacecraft hub has a nearly spherical mass distribution. For the more general case, a series solution is obtained and this solution is limited by its radius of convergence. The comparison of the different approximate analytical methods with numerical integration results are studied and it is observed that the oscillatory nature of the responses of the transverse angular velocity components reduces rapidly with faster extension rates.

Sellappan, R.↗
NASA NTRS1993

A multimission three-axis stabilized spacecraft flight dynamics ground support system

The Multimission Three-Axis Stabilized Spacecraft (MTASS) Flight Dynamics Support System (FDSS) has been developed in an effort to minimize the costs of ground support systems. Unlike single-purpose ground support systems, which attempt to reduce costs by reusing software specifically developed for previous missions, the multimission support system is an intermediate step in the progression to a fully generalized mission support system in which numerous missions may be served by one general system. The benefits of multimission attitude ground support systems extend not only to the software design and coding process, but to the entire system environment, from specification through testing, simulation, operations, and maintenance. This paper reports the application of an MTASS FDSS to multiple scientific satellite missions. The satellites are the Upper Atmosphere Research Satellite (UARS), the Extreme Ultraviolet Explorer (EUVE), and the Solar Anomalous Magnetospheric Particle Explorer (SAMPEX). Both UARS and EUVE use the multimission modular spacecraft (MMS) concept. SAMPEX is part of the Small Explorer (SMEX) series and uses a much simpler set of attitude sensors. This paper centers on algorithm and design concepts for a multimission system and discusses flight experience from UARS.

Langston, J.↗
NASA NTRS1971

The attitude determination of a three axes stabilized spacecraft

An approach is presented for attitude determination of a three axis stabilized spacecraft. The employment of two independent vectors is described whose components with respect to two reference coordinate axes can be known. Such two coordinate axes are right-handed three mutually orthogonal axes systems of which one is fixed in the spacecraft and the other has known orientation in space. The relationship between the two systems are then expressed as mathematical equations with the use of the two reference vectors. By computing the derived equations, the attitude of the spacecraft can be obtained.

Fang, A. C.↗