Spinning Spacecraft Research Articles

The Spin Scan Camera System: Geostationary Meteorological Satellite Workhorse for a Decade

The spin scan camera was first flown on ATS-1 from 1966 to 1973 to continuously monitor the global atmospheric circulation. Later, modified versions with improved sensor capabilities were flown on ATS-3 and SMS. Today, the GOES series of spin scan cameras is in regular operational use. The value of the spin scan camera as a dependable meteorological workhorse requires proper functioning of the entire camera system, composed of four main elements: a) a spinning spacecraft whose highly stable and predictable motions generate a time divisible precision scan and therefore a metric image; b) a telescope having both on-axis image quality and a wide field of view; c) a data chain which in-corporates duty cycle improvement and uses the spacecraft as a communication link to distribute the image data to users; and d) image display and analysis techniques which permit organizing a large number of images in the time domain and efficiently selecting and measuring data of greatest importance.

Pitch angle measurements from satellites using particle telescopes with multiple view directions

An analysis technique for obtaining spherical harmonic coefficients from multiple particle telescopes mounted with different viewing directions on a spinning spacecraft is presented. The coefficients and their standard deviations are obtained by the method of least squares. The relation of these coefficients to physically meaningful parameters is discussed. Applications of these techniques to data from instruments on two spacecraft are illustrated.

Passive magnetic momentum wheel unloading

Rotational kinetic energy may be extracted from the momentum wheels of a spacecraft by converting the rotational energy to electrical energy with a dynamo effect using the Earth's magnetic field, and dissipating the resultant kinetic energy as heat. This is an extension of the magnetic spin and nutation damping schemes used with spinning spacecraft. A three-axis momentum wheel assembly can be constructed with the desired inertial and autounloading characteristics that will operate in low Earth orbit. For synchronous altitudes, such a system is also possible, but the wheel design is affected greatly. The passive system is essentially failure-proof and results in a simplification of onboard systems and ground operations.

Design and Performance Analysis of Suboptimum Attitude Control Procedures for a Dual Spin Spacecraft

The characteristic properties of optimum attitude control procedures for a dual spin spacecraft, with gas jets on the rotor part, have been analysed. The direct implementation of their optimum control procedures is considered non-feasible. Therefore, suboptimum control functions are devised. The design of these controls makes use of the Lyapunov stability theory in order to obtain near time optimum control functions. The implementation of the near optimum control procedures is found to be relay type control functions with a linear time variable feedback law. The loss in performance for the near optimum control has been analysed, and found acceptably low and the actual loss can be controlled by the selection of the control magnitudes. A performance index has been derived for which the obtained suboptimum control procedures are found to be truly optimum.

On dynamic behavior of stochastic systems with discontinuities in parameters

In this paper an analysis and stochastic characterization of optimally controlled systems in the presence of environmental disturbances, additive input and observation noise, and subject to discontinuous-type parameter uncertainties is considered. As an illustrative example, the control of a spinning spacecraft is considered and some quantitative aspects of the stochastic behavior of the control and estimation signals following discontinuities in the controlled spacecraft parameters are discussed.

Fine Pointing Control for a Large Spinning Spacecraft in Earth Orbit

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Feasibility Study of Developing Torodial Tanks for a Spinning Spacecraft

A study was made to determine the feasibility of developing toroidal propellant tanks for a bipropellant (N204/MMH) propulsion system to be used in a proposed advanced Pioneer spin-stabilized vehicle intended for a Jupiter-orbiter and possibly a Saturn-orbiter mission. The rationale for considering the use of two toroidal tanks rather than the proposed use of four spherical tanks includes the belief that a more symmetrical distribution of propellant mass and a smaller variation in the position of the vehicle center-of-mass during propellant consumption would result, reducing requirements for attitude-control propellants, for balance weight, and for other weights associated with the dynamics of the spinning spacecraft. Results lead to the conclusion that a toroidal tank containing an effective, passive surface tension propellant acquisition device could be fabricated with available manufacturing methods and could be used interchangeably for either fuel or oxidizer.

Thermally Induced Nutational Body Motion of a Spinning Spacecraft with Flexible Appendages

This paper deals with a thermally induced nutational body motion of a spinning spacecraft with flexible appendages. The explanation of this phenomenon is as follows: when this class of spacecraft is exposed to solar radiation, thermally induced bending vibrations of appendages occur at a spin frequency. These vibrations cause a periodic variation of the moments of inertia of the spacecraft at that frequency and this, in turn, produces self-excitation of nutational body motions through the instrumentality of the parametric excitation. The amplitude of the nutational body motion is determined by means of the method of averaging and the stability of the motion is examined in detail.

Rotational motion of a free body induced by mass redistribution

The effects of mass redistribution on the rotational motion of a free body, initially rotating uniformly about its axis of maximum moment of inertia, are studied using a simple three-body model composed of two points masses and a triaxial rigid body. An exact analytical solution for the motion of the original spin axis after redistribution of the masses is developed from the classical solution for the rotational motion of a free triaxial rigid body. Yhis motion in interpreted geometrically using the reciprocal inertia ellipsoid of the redistributed system and the angular momentum integral. Application of simple mass redistribution to the problem of spacecraft attitude control is discussed and results for a particular system are presented. T is a well-known fact1 that for a free, triaxial, rigid body, the condition of uniform rotation about its axis of in- termediate, principal moment of inertia is unstable. This in- stability manifests itself when small disturbances in the angular velocity components about the other principal axes occur; for then, the once-stationary axis of intermediate moment of inertia performs large-angle motions with respect to a nonrotating reference frame. Recently, Beachley and Dicker2 and Beachley,3 have suggested that this characteristic of rigid body rotational motion might be used to advantage to a spinning spacecraft in the sense that predictable large-angle motions of an axis fixed in the spacecraft might be induced by changing the mass distribution of the spacecraft.% A control system based on this idea might be more economical than a mass expulsion system which would rotate the angular momentum vector of the spacecraft. Changes in the mass distribution of the spacecraft may be accomplished by internal motion of control masses. The ef- fects of such internal mass motion on the rotational motion of a free system of bodies have been studied by several authors and a review of some of the work in this area is given by Lorell and Lange,5 who propose an automatic mass-trim system for spinning spacecraft which utilizes movable control masses. The primary purpose of this paper is not to study the effects of internal mass motion on the rotational motion of a free system of bodies, although these effects cannot be ignored en- tirely and will be considered. Our main goal is to obtain an analytical description of the spacecraft's rotational motion af- ter a redistribution of mass has occurred. Such a description has previously been obtained only in a limited way by numerical integration.3 We shall also provide geometrical in- terpretation of the analytical results and consider the use of simple mass redistribution as a means for spacecraft attitude control. In this paper we shall consider a system similar to those discussed in Ref. 3 and employ some results from the classical theory for the rotational motion of a free, triaxial, rigid body.

A Rotor-Fixed Modal Simulation Model for Flexible Rotating Equipment

A transient, flexible rotor formulation is derived on the basis of a representation previously employed to simulate the motion of flexible spinning spacecraft. The distributed parameter characteristics of the rotor are approximated by modeling the rotor as an elastically connected group of n rigid bodies. The elastic rotor deflections of the component rigid bodies are defined in terms of a rotor-fixed frame of reference; hence, during constant synchronous whirling the elastic deflections appear to be constant. The model is initially simplified by the traditional small deflection assumptions of the theory of elasticity, and is additionally simplified by the use of modal coordinates. Modal coordinates dramatically reduce the dimensionality of the model, and significantly clarify the dynamic analysis of the problem. Required data input to the model, and typical model output are demonstrated for the Mark 15-F turbopump of the Rocketdyne J-2 engine system. The model is shown to correctly demonstrate the form of unstable rotor whirling associated with internal hysteresis damping when operating above the first bending-mode critical speed.

On the measurement of energetic particle flux anisotropies with a class of spinning detectors

Some of the charged particle telescopes on spinning spacecraft operate in an analysis-readout mode in which a Poissonlike statistical bias can significantly distort both the amplitude and the direction of the measured flux anisotropy. Briefly, if the telescope accepts the first event after telemetry readout for analysis and then accepts no further events for analysis until after the next readout, then, if the conventional computational techniques are followed, the anisotropy amplitude deduced from the analyzed events will be reduced from the true anisotropy, and the deduced direction of maximum flux will lag the true direction (in the sense of the spacecraft spin). For high enough fluxes the rates will be reduced to the sampling rate (a well-known effect), but for weak anisotropies the lag will approach 90° (an effect not discussed heretofore). However, these distortions may be computed if the incident rate of unanalyzed events is measured independently, and hence the measurements may be corrected for the statistical bias in a straightforward manner.

Nutation dampers vs precession dampers for asymmetric spinning spacecraft.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

A Magnetic Attitude Control System for an Axisymmetric Spinning Spacecraft

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Large deformation, deep penetration theory for a compressible strain-hardening target material

1 lorillo, A. J., Damping Dynamics of Axisymmetric Rotor Stabilized Satellites, ASME Winter Meeting, Chicago, 111., Nov. 1959. 2 Likins, P. W., Criteria for Dual Spin Journal of Spacecraft and Rockets, Vol. 4, No. 12, Dec. 1967, pp. 1638-1643. 3 Mingori, D. L., Effects of Energy Dissipation on the Attitude of Dual-Spin Satellites, AIAA Journal, Vol. 7, No. 1, Jan. 1969, pp. 20-27. 4 Pringle, R., Jr., Stability of the Force-Free Motions of a Dual-Spin AIAA Journal, Vol. 7, No. 6, June 1969, pp. 1054-1063. 5 Velman, J. R., Simulation Results for a Dual-Spin Proceedings of the Symposium on Attitude Stabilization and Control of Dual-Spin Spacecraft, Aug. 1-2, 1967, Rept. SAMSOTR-68-191, Nov. 1967, U. S. Air Force. 6 Johnson, C. R., Tacasat I Nutation Dynamics, to be published in AIAA Progress in Astronautics and Aeronautics: Communications Satellite Technology for the 70' s, Vol. 25, edited by N. E. Feldman and C. M. Kelly, MIT Press, Cambridge, 1971.

Attitude control of a sun-pointing spinning spacecraft by means of solar radiation pressure

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Stability of the force-free motions of a dual- spin spacecraft.

References 1 Anderson, J. D., of the Masses of the Moon and Venus and the Astronomical Unit from Radio Tracking Data of the II Spacecraft, Doctoral Dissertation in Astronomy, UCLA; also TR 32-816, July 1967, Jet Propulsion Lab. 2 Vegos, C. J. and Trask, D. W., Ranger Combined Analysis Part II: Determination of the Masses of the Earth and Moon from Radio Tracking Data, Jet Propulsion Laboratory Space Programs Summary 37-4/> 3 Null, G. W., Gordon, H. J., and Tito, D. A., The IV Flight Path and Its Determination from Tracking Data TR 32-1108, Aug. 1967, Jet Propulsion Lab. 4 Pease, G. E., private communication. 5 Pease, G. E., Mariner V Real Time Cruise and Encounter Orbit Determination Results, Jet Propulsion Laboratory Space Programs Summary 47-50, Vol. II, March 1968, pp. 125-141. 6 Esposito, P. B., On the Nature of the Gas Leak on the Pioneer VI Spacecraft, private communication, TM 311-2, June 1967, Sec. 311, Jet Propulsion Lab. 7 Melbourne, W. G., Muhleman, D. O., and O'Handley, D. A., Radar Determination of the Radius of Venus, Science, Vol. 160, No. 3831, May 1968, pp. 987-989.

Attitude stability criteria for dual spin spacecraft.

Attitude stability eriteria are developed for spacecraft, which consist of two primary bodies capable of unlimited relative rotation about a common axis. Such vehieles, typified in existing hardware by the Orbiting Solar Observatory (OSO) satellites, have applicability to missions for which the simplicity, reliability, and longevity of spin stabilization combine with a requirement for unidirectional pointing of a component, such as an antenna or a solar panel. The utility of dual spin systems is substantially increased by the results of attitude stability analysis, which reveals the possibility of obtaining passive stable spin-axis attitude by spinning a vehicle about its axis of greatest or least inertia, providing only that an effective energy dissipation device is attaced to a counter-rotating platform which has greatly reduced spin. This result is coutrary to common interpretation of the familiar major axis requirement for stability, according to which energy dissipation in a vchicle rotating about its axis of least inertia must produce instability. Stability eriteria are developed first by Routhian analysis of a system with energy dissipation capability in only one of the two primary bodies, and subsequently in more general (but less rigorous) terms for a fully dissipative vehiele. Numerical integration and model studies provide corroboration.

Nutational stability of an axisymmetric body containing a rotor

T7K)R many satellite and space-probe missions, full three-F axis attitude control is not required, and simple (passive) spin stabilization is sufficient to insure satisfactory system performance. In many applications, in fact, the need for stabilization is not dictated by primary mission requirements at all but arises indirectly, perhaps as a means of minimizing or removing thermal and antenna design problems. Spin control is ideal for this type of mission. In other caess, considerable improvement in system performance is afforded, e.g., in a high-altitude communication satellite, increased antenna gain can be obtained using an antenna with a toroidal radiation pattern. Fully passive techniques for three-axis attitude control have only limited application, and it is generally necessary to employ far more sophisticated, active stabilization methods such as mass expulsion or flywheel control systems. However, there is one active, but fairly simple, three-axis control technique that is a natural extension of spin stabilization. In this technique, control of the third axis (control about the spin axis) is achieved by despinning a portion of the spinning spacecraft. That is, instead of the vehicle being a single body, it is constructed of two bodies (I and II in Fig. 1) constrained to rotate about a common axis. Suppose that II is the despun portion. A motor (not shown) is included which can change the relative rate about OZ. If the motor is servo-controlled, say by a sensor mounted on one of the bodies, the rate of body II can be maintained at zero or can be varied to permit a given axis in II, normal to OZ, to track an external reference. Thus, in effect, the despun component is three-axis stabilized by means of a single-axis active controller. Control of spin-axis attitude, if required at all, would normally be by ground command to activate, for example, pulse jets, or, in the case of earth orbiters, a magnetic coil mounted on the spinning component. A wide range of configurations can be devised which use this basic two-body construction. One application is id the use of a small directional antenna as the despun portion of a spin-stabilized communications satellite. In this application, the spin axis would be maintained normal to the orbit plane, and the antenna controlled to point along local vertical, allowing maximum gain to be obtained. Another type of single-axis control can be provided if the major part of a spacecraft is despun, and a high-speed internal rotor is employed to provide the gyroscopic stiffness. In fact, this case can be considered a derivative of a three-flywheel system from which two wheels are removed and the third operated about a high bias rate instead of null. Because of the high speed of the internal rotor, spin-axis pointing can again be controlled by ground command. The two-body configuration does appear to offer definite advantages as a means of achieving simplified three-axis attitude control of spacecraft; for this reason, it has received, and is receiving, serious attention in the industry. Indeed, one spacecraft of this construction, OSO I (built by Ball Brothers for NASA), was launched over two years ago and has operated successfully in orbit. In this paper the nutational stability of the two-body system is considered, and the constraints that the requirement for stability places on the moments of inertia of the configuration are established. SYMBOLS: O X , , O Y , , O Z ARE A X E S FIXED IN BODY I