SPACECRAFT

Abstract: A spacecraft having a sun ray blocker device (111, 112, 141, 271, 301, 411, 511, 611, 811, 921, 951, 1800, 2100, 2700) for shading a thermal radiator surface (11, 12) of the spacecraft in which the sun ray blocker device is movable in relation to the thermal radiator surface to keep the surface substantially in shade without substantially blocking thermal radiation from the thermal radiator surface to deep space. Preferably a sun-facing side (111a, 112a) of the sun ray blocker device is thermally insulated from an opposed side (111b, 112b) to reduce thermal radiation from the sun ray blocker device to the thermal radiator surface and the sun ray blocker device is also preferably deployable in orbit after launch.

Description: SPACECRAFT The present invention relates to a spacecraft and particularly to a spacecraft having a sun ray blocKer device for shading thermal radiator surfaces on said spacecraft from solar heating. The following patents are generally representative of the prior art in the broad fields of solar array related sun shields, solar array deployment mechanisms, and the thermal control of radiator surfaces for various types of spacecraft. United States Patent 4,133,502 to Andrew Anchutin describes a plurality of arrays of solar cells which are symmetrically stored about a spacecraft during launch to provide symmetrical loading. When the spacecraft is in operational configuration, the solar, arrays are deployed adjacent each other on one side of the spacecraft to eff ctively form a single array and the single array may be oriented to face the Sun by a common drive mechanism. United States Patent 4,508,297 to Guy G. Mouilhayrat et al, describes an equatorial orbit satellite with solar panels having blades with a median line inclined at a certain angle relative to the equatorial plane. Thus, the field of vision of the antennas is free and disturbing torques become acceptable. United States Patent Number 5,372,183 to Harold P. Stric berger describes a spacecraft adapted for operation in a low inclination angle earth orbit which comprises north, south, east and west panels defining a spacecraft interior volume. The north and south panels are oppositely disposed with respect to each other and the east and vest panels are oppositely disposed with respect to each other. The spacecraft interior volume generally and preferably lacks structural elements that substantially restrict thermal radiation among the panels. The north and south panels, to which spacecraft equipment is usually mounted, each include conductive heat pipes for reducing the temperature dif erence across each panel. The exterior surfaces of the north, south, east and west panels have a covering, preferably of optical solar reflectors (OSRs) , for radiating thermal energy therefrom, wherein the OSRs have a solar absorptivity that is substantially less than their thermal e issivity. The interior surfaces of the north, south, east and west panels have a covering for ef ectively radiating thermal energy between and among the panels across the interior volume. United states Patent 4,725,023 to Haruo shiki describes a geostatic satellite which comprises a spinning drum for stabilization which spins around an axis of rotation which is parallel to the axis of the Earth. A paddle member loaded with solar cells is directly rotatable about the same axis and is controlled such that the solar calls face the sun. A de-spun platform supports communication gear and maintains the gear pointed to a relatively fixed point on Earth. A shading device for shading the electronics laden de-spun platform from the Sun is attached to the paddle member and rotatable therewith. Thereby, the shading device will always be disposed between the Sun and the de-spun platform. However, the shading device also blocks thermal radiation from the platform and also itself heats up in sunlight and radiates heat towards the platform, decreasing the efficiencies of heat transfer from the spacecraft to space. Notwithstanding the prior art, the present invention is neither taught nor rendered obvious thereby. It is an object of this invention substantially to reduce or eliminate the direct and indirect solar heating of certain spacecraft radiator-panels, and to also minimize the magnitude of any reduction in the radiative- view-factor of the (shielded) radiator panel. In order to achieve that objective, the materials and design selected for the sun ray blocker device, which will be discussed below, should ideally provide all of the following: minimum blockage of the field-of-view to deep space of its associated radiator surface(s) , low absorption of the solar energy incident on its front (sunward) surface, high radiation of absorbed thermal energy back to space, and high insulation of heat between the front (sunward) and back (anti-sunward) sides of the sun ray blocker device. It is also desirable to provide a sun ray blocker device that is capable of greatly reducing or eliminating solar energy incident on those sides of certain spacecraft relative to which the Sun direction makes a low angle. The types of spacecraft to which the present invention applies include some spacecraft for operation in equatorial or low inclination orbits, and in sun synchronous orbits with low orbit-sun angles. In the case of three-axis stabilized, Earth-pointing, geostationary spacecraft for example, these shaded sides are either or both of the north and south main-body panels. In the case of the sun synchronous spacecraft for example, the shaded sides are either or both of the sides or main-body panels that face out along the pitch axis (i.e. that face parallel to the orbit normal and anti-normal) . The present invention can also be applied to types of spacecraft, other than geostationary and sun synchronous types, upon which the solar illumination is incident at low angles relative to thermal radiator surface. In those spacecraft it is those main thermal radiator surface that can be shaded by the present invention device. According to a first embodiment of the invention there is provided a spacecraft for orbiting a sunlit celestial body, the spacecraft including a thermal radiator surface for radiating heat from the spacecraft into space, and a sun-ray blocker device mounted on said spacecraft for shielding said thermal radiator surface from rays of sunlight, characterised in that said sun ray blocker device is locatable for placing in shadow substantially the whole of the thermal radiator surface from sunlight without substantially impeding thermal radiation from said thermal radiator surface into space. , Preferablyan effective radiation view factor for thermal radiation from the thermal radiator surface ( 1 1,12,1804, 2121,2721) to deep space is significantly greater than corresponding geometrical radiation view factor to deep space, and in particular, for a geostationary spacecraft where the geometrical radiation view factor to deep space is 0.65, the effective radiation view factor to deep space is at least 0.87. A pref rred inventive feature will be seen to reside in the sun ray blocker device including at least one sun blocker panel having a sun-facing surface and an opposed anti-sun-facing surface, wherein the εun-facing surface is thermally insulated from the opposed surface. conveniently the sun-facing surface is thermally insulated from the opposed surface by a multi-layer insulation blanket. Advantageously the sun-facing surface has a low solar energy absorptivity of less than 0.5. Advantageously the sun-facing surface includes a solar cell panel for supplying electrical power to the spacecraft. Preferably the sun-facing surface has a high thermal e issivity of higher than 0.7. A preferred inventive feature will be seen to reside in the sun ray blocker device being moveable between a stowed, non-operative position and a deployed operative position. Conveniently the sun ray blocker device includes an attachment arm for attaching the sun blocker panel to the spacecraft. Advantageously the attachment arm is attached by a hinge means to the sun blocker panel and/or by a second hinge means to the spacecraft. Advantageously the sun ray blocker device includes a motor or moving the sun ray blocker device between the stowed position and the deployed position. Preferably locating means are provided for locating the sun ray blocker device with respect to the thermal radiator surface which include adjustment means to maintain the majority of the thermal radiator surface in shadow irrespective of changes in the attitude and/or orbit of the spacecraft. Advantageously the adjustment means includes carriage means (i80i, 1901, 2001) for carrying the sun blocker panel (1800, 1900, 2000) and transport means (1802, 1808, 1903, 1904, 2003, 2006) for moving the carriage with respect to the spacecraft. Conveniently the transport means includes rail means (1802) nd the carriage means (1801) includes drive means to drive the carriage along the rail means. Preferably the transport means includes an annulus rotatable in circular path defined by bearing means, the annulus being driveable by drive means to move the carriage along the path defined by the bearing means. Alternatively the transport means includes rail means (1902) and belt means (1903) connected to the carriage means (1901) , the belt means being driven by drive means (1904) to move the carriage means along the rail means (1902) . Conveniently the adjustment means includes a variable length attachment arm (2101, 2703) or attachment of the sun blocker panel (2100, 2700)to a solar cell array assembly (2103, 2701) for rotation with the solar cell array assembly, such that the distance of the sun blocker panel from the solar cell array assembly may be varied during rotation. Alternatively the spacecraft has a solar cell array adapted for tracking movements of the sun relative to the spacecraft, wherein the adjustment of the location of the sun ray blocker device in relation to the thermal radiator surface is synchronised with the tracking movement of the solar cell array, when in normal operation. Conveniently the sun ray blocker device is mounted on the solar cell array or on means carry said solar cellar array. Advantageously the solar cell array tracks the movement of the sun by rotation of the solar cell array about an axis of rotation of the solar cell array such that the sun blocker panel also rotates about said axis of rotation of the solar cell array. Conveniently the thermal radiator surface is orthogonal to the axis of rotation of the solar cell array so that the sun blocker panel rotates about an axis normal to the thermal radiator surface. Advantageously adjustment means includes a variable length attachment arm for attachment of the sun blocker panel to the spacecraft. Advantageously the attachment arm is a scissor arm (2703) . Alternatively the attachment arm (2101) is formed of articulated portions (2104, 2105, 2106) hich may be mutually articulated during rotation to vary an effective length of the attachment amn. Conveniently adjustment means for attachment of the sun blocker panel to a solar cell array assembly are such that a distance between the sun blocker panel and the solar cell array assembly may be varied during rotation of the sun blocker panel. Conveniently means are provided for adjusting the size of the sun blocker panel. Conveniently the spacecraft includes control means for controlling the spacecraft so as to maintain an angle between a sun line and the thermal radiator surface below a predetermined angle by adjustment of the orbit and/or attitude of the spacecraft in use. Preferably the predetermined angle is 60 degrees. More preferably the predetermined angle is 45 degrees. Host preferably the predetermined angle is 23.5 degrees. Advantageously the control means is adapted to maintain the thermal radiator surface substantially parallel to a plane of an orbit of the spacecraft. Alternatively the control means is adapted to maintain the spacecraft in a sun synchronous orbit. Alternatively, the control means is adapted to maintain the spacecraft in an equatorial or low- inclination orbit. According to another embodiment of the invention there is provided in a three axis stabilised spacecraft for orbiting about a planet and having at least one solar cell assembly having at least one solar cell panel, and being a north solar cell panel assembly or a south solar cell panel assembly, said at least one solar cell panel assembly being mounted on an axle so as to be controllably rotated from said spacecraft about an axis of rotation so as to face the sun, said spacecraft having a nadir panel which is generally pointing to the centre of the planet, an opposite panel known as a zenith panel, which faces away from the centre of the planet and Sharing the same planar normal vector as said nadir panel, an east panel and a west panel, said east panel and said west panel having their planar normal vector laying on a orbital plane pointing to the velocity vector of the spacecraft generally tangential to the direction of travel on the orbit, and a north panel and a south panel, said north panel and said south panel having their planar normal vector generally perpendicular to the orbit plane and parallel to the axis of rotation of the planet, said solar cell panel extending outwardly from said spacecraft, the improvement which comprises: attaching at least one sun ray blocker device to said at least one solar cell panel, said at least one device being either a north blocker device or being a south blocker device and corresponding to said at least one solar cell panel, each of said at least one sun ray blocker device being positioned forwardly from and offset relative to a solar cell surface of a solar cell panel and at a predetermined angle to either of said north panel and said south panel, said north panel or said south panel, said sun ray blocker device being positioned so as to cast a shadow on at least a majority of the exposed surface of its corresponding north or south panel during solar exposure thereto. According to another embodiment of the invention there is provided in a three axis stabilised low inclination orbit spacecraft for orbiting about the earth and having two sets of solar cell array assemblies having solar cell arrays, one set being a north solar array assembly and the other being a south solar array assembly, said assemblies each being mounted on an axle so as to be controllably rotated from said spacecraft about an axis of rotation so as to face the sun, said spacecraft having an earth panel which is generally pointing to the centre of the earth, an opposite panel known as a zenith panel, which faces away from the centre of the earth and sharing the same planar normal vector as said earth panel, an east panel and a west panel, said east panel and said west panel having their planar normal vector laying on an orbital plane pointing to the velocity vector of the spacecraft generally tangential to the direction of travel on the orbit, and a north panel and a south panel, said north panel and said south panel having their planar normal vector generally perpendicular to the orbit plane and parallel to the axis of rotation of the earth, said solar cell panel extending outwardly from said spacecraft, the improvement which comprises: attaching at least one sun ray blocker device to each of said north solar array and said south solar array, one device being a north device and another device being a south device, each of said sun ray blocker devices being in the form of a panel and being positioned forwardly and offset relative to the solar cell surface of a solar ray and at a predetermined angle to said north panel and said south panel, said north blocker device being positioned so as to cast a shadow on at least a majority of the exposed surface of said north panel during solar exposure thereto, and said south blocker device being positioned so as to cast a shadow on the exposed surface of said south panel during solar exposure thereto. The present invention preferably provides a sun- synchronous sun ray blocker device (not to be confused with sun synchronous orbits referred to elsewhere Herein) for use in a spacecraft designed to orbit around a planet with solar incidence at low angles to their thermal radiator surfaces, i.e. with sun directions close to the planes of the individual radiator surfaces. Preferred embodiments of the present invention are spacecraft for operating in an orbit plane oriented at a low angle (or within a range of low angles) to the sun direction, the said spacecraft having a thermal radiator surface that is oriented approximately parallel to the orbit plane and a solar array assembly that is rotated about an axis approximately perpendicular to the orbit plane nominally at the orbital rate. Examples of appropriate orbits are: (a) low inclination orbits around the Earth (including nominally equatorial orbits) , and (b) sun synchronous- orbits with low orbit-sun angles (which around Earth and Mars, for example, are nominally polar orbits) . The term "spacecraft 11 as used herein includes satellites and other space bound vehicles. Mounted on any spacecraft to which the present invention is applied is at least one device for blocker sun rays and thereby preventing them from directly impinging on a radiator surfaces of the spacecraft. In many embodiments of the present invention the individual spacecraft will have at least one solar array assembly (comprised of solar cell panels and rotary axial booms) which may be used as mounting support for the sun ray blocker device(s) , so that the combination assembly of solar array assembly and sun blocker device(s) is operationally controllably rotated together as an integral unit to track the Sun throughout the orbital revolutions of the spacecraft, said solar array assembly being mounted on the spacecraft so that operationally in orbit it can be rotated about an axis that is maintained oriented approximately perpendicular to the orbit plane in a manner such that the solar-cell side of the solar cell panels is maintained sun facing and substantially perpendicular to the Sun direction. Because then the sun ray blocker device may rotate integrally with the solar array assembly, it is able to prevent sun rays from directly impinging on all or part of an associated thermal radiator surface(s) , whose plane is maintained approximately parallel to the orbit plane, thus creating a continuous steady and benign thermal environment for the thermal radiator surface. spacecraft operated in orbits with low orbit-sun angles (the angle between the orbit plane and the Sun) are prime candidates for application of the current invention device. Various different frequently-utilized types of orbits feature low orbit-Sun angles. Currently, among the most utilized types of orbits with low orbit- Sun angles are (a) low inclination and nominally- equatorial orbits, including geosynchronous orbits, and (b) the subset of sun synchronous orbits with low orbit- Sun angles. Sun synchronous orbits maintain a little- varying orbit-sun angle as the planet revolves around the Sun. The Earth revolves around the sun once per year. One type of spacecraft operated in a nominally equatorial orbit around a planet, e.g. the Earth, or in particular, a geosynchronous orbit, is frequently used for the purposes of telecommunications, broadcasting, monitoring ecological conditions, global positioning, remote sensing, surveillance and weather forecasting. Another type of satellite operated in nominally sun synchronous orbits around planets, e.g. the Earth, with low orbit-Sun angles, is frequently used for the purposes of weather monitoring and remote sensing of the planet and its atmosphere. Some of the benefits of these sun synchronous orbits are: low spacecraft altitudes, frequent over-flight of the planet within close proximity of virtually all latitudes and longitudes, and near constant angle of solar illumination on the day side of the orbit. Means of adjusting the attitude and orbit of spacecraft are well known, for example, are described in "Principals of Communication Satellites" by Gary D Gordan and Walter L Morgan published by John Wiley & Sons 1993, pages 12-14, 55-58 and in "Spacecraft Attitudes, Termination and Control" by James R Wertz published by Kluwer Academic Publishers 1978. Attitude and orbit control may for example be provided by the use of thruεters and/or momentum or reaction wheels. Typically, the attitude (i.e. the orientation) of these types of satellite is controlled so that as the satellite orbits the planet part of its payload equipment steadily faces approximately toward the center of the - planet, while the solar arrays are maintained sun pointing. The attitude (orientation) control systems of such spacecraft belong to various classifications that are well known within the space industry. For example, two of the more currently prominent types of attitude control system are commonly referred to as "three-axis- stabilized" control systems and "spin stabilized" control systems. The present invention device functions independently of the type of attitude control system, and independently of the orientation of spacecraft equipment other than the orientation of the thermal radiator surfaces that the device shields from solar energy. In these types of spacecraft, the performance of the present invention device is generally better the closer the shadowed thermal radiator surface is to being parallel to the orbit plane (which in these types of spacecraft is maintainable at a low angle to the sun line) . Hereinafter, the concept of a "model spacecraft" is defined and employed in order to avoid the distraction of multiple lengthy descriptions of diverse spacecraft to which the present invention device may be applied. The model spacecraft is used herein, somewhat like a tailor' s dummy, in order to facilitate the illustration and explanation of features, functions, and examples of applications of the present invention device. By definition the model spacecraft has a basic, deployed (i.e. unfolded), structural configuration that is typical of many current three axis stabilized satellites, and a corresponding operational mode that is typical of a geostationary Earthpointing spacecraft, tfote that this definition was selected on the basis of current estimates of the most frequent future application of the present invention device. The definition of the model spacecraft could equally well have been based on typical characteristics of another relevant type of spacecraft, for example an Earth-pointing spacecraft in a sun synchronous orbit with a low orbit-Sun angle. Referring to FIGURES 1 and 5, the basic structural configuration of the model spacecraft is based on a main body in the form of a hollow, right parallelepiped. For the stated purposes of using the concept of the "model spacecraft" herein, it is useful to consider the main body as being comprised of six principal, planar, structural panels. The external surfaces of one opposing pair of the six panels that form the main body of the model spacecraft constitute the mounting sites for the thermal radiator sur aces that are shielded from direct solar heating by means of the present invention device. Mounted on one or each of these two radiator-bearing panels, and extending perpendicularly outwards therefrom, is a solar array assembly, comprised of a rotary solar array boom to which are attached solar cell panels. That is not to say that application of the present invention device is limited to spacecraft resembling the structural configuration and operational mode of the model spacecraft. For example, the present invention device is also applicable to: sun synchronous spacecraft in orbits with low orbit-Sun angles; spin stabilized spacecraft; spacecraft with polyhedral and/or irregular structures; spacecraft that are not nadir pointing; spacecraft with solar arrays that deploy and subsequently lie along axes that are not perpendicular to the radiator-bearing panels; etc. Much of the text herein that supports the accompanying claims is written with reference to the model spacecraft. Regardless, the supporting text also applies to applications of the present invention device to other suitable types of spacecraft. For example, the principal relevant difference between many suitable sun synchronous spacecraft ( for polar orbits at Earth and Mars at least, where the polar axes lie close to the planes of sun-synchronous orbits around them) and the model spacecraft is that the plane of the thermal- radiator surface(s) that is shaded by the present invention device is approximately parallel to the axis of rotation of the planet (rather than perpendicular to the axis of rotation of the planet as for geostationary spacecraft like the model spacecraft) . Accordingly, the supporting text describing spacecraft like the model spacecraft is easily read as it relates to these suitable sun synchronous spacecraft, for example by substituting £n "pitch axis panel" or "orbit normal panel" to replace "north/south panel", and substituting in "velocity panel" or "roll axis panel" to replace "east/west panel". In order to provide functional services in an operational orbit, the model spacecraft has one of the six structural panels of its main body continuously facing the planet, e.g. the Earth. That panel is referred to as the earth panel or the nadir panel. A vector that is outward-from and normal-to the earth panel is parallel to the (body fixed) yaw axis of the spacecraft. In the model spacecraft the yaw axis is maintained nominally parallel to the nadir direction, i.e. is nominally pointed toward the center of the planet. Because the model spacecraft operates in a geosynchronous orbit, which is nominally circular, the yaw axis of the model spacecraft is maintained nominally perpendicular to the velocity vector of the spacecraft. A vector that is outward-from and normal-to the plane of the structural panel opposite the nadir panel is parallel to the negative yaw axis. That main-body panel of three axis stabilised Earth pointing geostationary spacecraft like the model spacecraft is usually referred to as the zenith panel or anti-earth panel. Another opposing pair out of the six structural panels comprising the main body of the model spacecraft are oriented so that, nominally or approximately, vectors that are outward-from and normal-to their planes lie in the orbital plane and are perpendicular to the yaw axis and to the nadi and zenith directions. These outward normal vectors are parallel and anti-parallel to the positive and negative (body fixed) roll axes of the spacecraft. Because the geosynchronous orbit of the model spacecraft is circular, the roll axes of the model spacecraft nominally coincide with the velocity and anti- velocity vectors of the orbital motion. For geostationary spacecraft the velocity of the spacecraft is eastward; and consequently these two main-body panels of spacecraft like the model spacecraft are generally referred to as the east and west panels. Accordingly, the remaining two structural panels comprising the main body of the model spacecraft are oriented so that their planes are nominally or approximately parallel to the orbit plane. Vectors that are outward-from and normal-to the planes of these panels are parallel to the positive and negative (body fixed) pitch axes of the spacecraft. Since the geosynchronous orbit of the model spacecraft is nominally equatorial, "the pitch axes of the model spacecraft are approximately parallel and anti-parallel to the spin axis of the Earth; and accordingly these two main-body panels of spacecraft like the model spacecraft are referred to as the north and south panels. To avoid unnecessary further repetition, to illustrate and explain the features and functions of the present invention device, reference shall be made to application to the (previously defined) model satellite, which is operated in a three-axis-stabilized, Earth- pointing, geosynchronous mode. (That is not to say that application of the present invention device is limited to spacecraft that resemble the previously described model spacecraft and/or are operated in the corresponding mode. For example, the present invention device is also applicable to spin stabilized spacecraft with irregularly shaped structures that are neither nadir pointed nor geosynchronous. ) Through each orbital revolution of a spacecraft like the model spacecraft, which in a preferred embodiment is around the Earth, the Sun sequentially directly illuminates the east, zenith, west, and nadir main-body panels. While illuminated (or insolated) thus these main- body panels absorb incident solar energy and their temperatures increase, which significantly reduces their net radiative cooling capability. If not countered by some means this can signi icantly limit the quantity of equipment (which dissipate heat into the spacecraft) that can be carried on board, and/or can result in undesirably elevated temperatures of associated spacecraft equipment. The north and south panels, however, generally face deep space during the entire orbit and only directly receive solar illumination and solar energy at relatively low co¬ incidence angles on a seasonal basis. Because the direct input of solar energy into the north and south panels is relatively low to zero, these panels are the principal sites on spacecraft like the model spacecraft for the locations of thermal-energy radiator surfaces. The north panel is directly heated by the sun for a duration of about 6 months (from about March 21st to about October 21st) at an incidence angle, defined as the angle between the panel plane and the sun vector, which seasonally increases from o degrees (when the sun vector is edge-on to the panel) to about 23.5 degrees followed by a decrease to O degrees again while the Sun is on the north side of the earth equator, i.e. during the northern spring and summer. The south panel is directly heated by the Sun for the remainder of the year, i.e. during the southern spring and summer, in a similar fashion and concomitantly with the north panel. These relatively low solar incidence angles favor use of the north and south panels for locating the principal thermal radiator surfaces of the spacecraft. At a maximum incidence angle of 23.5 degrees for the solar vector relative to the north and south panels the incident solar energy is approximately 40% of that for normal (perpendicular) incidence. In the prior art numerous design practices have been employed to the surface treatment of the north and south panels in an effort to reduce the absorbed solar energy, thereby allowing more internal heat dissipation without raising the operational temperature level of the equipment that is thermally coupled to the panels. One example, optical surface reflectors (OSRs) , which have a high ratio of thermal emisβivity versus solar absorptivity, have been widely used as the surface treatment of spacecraft thermal radiators. However, the seasonal solar heating of the spacecraft through OSRs still constitutes a significant amount of heat input to the spacecraft, which forces the spacecraft designer to lower the level of internal power dissipation to maintain an acceptable operating temperature for the spacecraft equipments. Solar energy absorbed by a spacecraft like the model spacecraft through its north and south panels has two obvious undesirable impacts on the performance of the spacecraft. (1) It reduces the allowable level of internal power dissipation, which directly relates to the "value" of a spacecraft. The revenue from a spacecraft, especially a commercial communications spacecraft, is fundamentally limited by its capacity for power dissipation. A reduced allowable power dissipation level directly results in lower potential for revenue generation, which reduces the value of the spacecraft. (2) The operating temperatures of the internal equipment are increased, and as a result the reliability of those components may be reduced. The reliability also relates to the life of a spacecraft, which directly relates to its "value" as well. If the undesired solar heating were to be reduced, higher operational payload power would be allowable within the spacecraft and/or lower operating temperatures of the spacecraft equipments would be achieved. Therefore, by virtue of the present invention, the spacecraft could be operated at a higher efficiency, with higher reliability, and would thereby generate revenue at a faster rate, all of which improvements would increase its value. There is another important factor that affects the capability of a thermal radiator surface to reject heat to deep space: the "effective" radiative view factor (ranging from 0 to l) from that panel to deep space. The ideal radiative-view-factor enabling a panel to reject maximum heat into deep space is unity (l) . A device or means situated between the radiator surface and deep space could block the radiator's view to deep space and thus reduce the heat-radiating capability of the radiator. The sun ray blocker device of this invention is mounted on the spacecraft, for example conveniently attached to the solar array assembly/assemblies of the spacecraft and rotating therewith. Since the primary function of the sun ray blocker device used in this invention is to provide a significantly more benign thermal environment for the principal thermal radiator surfaces (or panels) , basically by shading them, the spacecraft should have at least one such surface. In the f case of three-axis-stabilized Earth-pointing geostationary spacecraft, for example, there are two principal thermal radiator surfaces - the north and south panels; and accordingly at least two separate sun ray blocker devices can be included, one to shade each of these panels. Thus, the sun ray blocker device in the present invention follows the movement of the Sun with respect to the thermal radiator panel(s) that it shades. In the case of a three-axis-stabilized Earth-pointing geostationary spacecraft, like the model spacecraft for example, the sun ray blocker device casts its shadow onto its associated thermal radiator surface, which is on either a north or a south panel, seasonally - through the six month long northern spring and summer in the case of the north panel, and through the six month long southern spring and summer in the case of the south panel. The (counter-productive) reduction in the radiation-view- factor of the thermal radiator surface caused by the presence of the associated present invention device is small; and the net effect of this reduction combined with the (beneficial) shading of the panel is a great improvement in the radiative efficiency of the radiator surface. In addition to the foregoing, some of the considerations, advantages and parameters for the present invention device are as follows (others will become self- evident from the subsequent discussion of the FIGURES) : Variety in the operational form and size of the sun ray blocker device is permissible provided that its sun blocker panel follows the movement of the Sun with respect to the spacecraft, and it blocks the Sun's rays by casting a shadow onto the spacecraft main body at appropriate times, and it produces close to the minimum reduction in the effective radiative view factor to deep space of the thermal radiator that it shields, and it satisfies other system requirements of the spacecraft (for example clear ield of view requirements) . The material and/or the construction of the sun blocker panel of the sun ray blocker device is preferably highly thermally insulating between its sun and anti-sun sides in order to provide the greatest practical effective radiative view factor and radiative ef iciency of the radiator-surface shielded by the sun blocker panel. In its fully deployed con iguration the sun blocker panel may be mounted through a wide range of orientations relative to the radiator surface that it shields (for example, the angle 501 in FIGURE 12a below does not have to be 90 degrees i.e. a right angle) as long as it casts shadow providing adequate coverage of the associated thermal radiator surface(s) on the spacecraft. The ideal width of the sun blocker panel is greater than either the width or the length of the radiator surface that it shields. However, the dimensions of the sun blocker panel may be limited by other constraints. For example, in the launch configuration the folded dimensions of the sun blocker panel may be limited by launch-envelope constraints, i.e. the si2e of the volume allowed for the spacecraft by the launch vehicle during launch. Therefore, it may be necessary to make the sun ray blocker panel deployable to enable it to be folded for launch and deployed in orbit. This can be achieved by hinged deployment, slide extension, pre-offset or any other means to increase the width of the sun blocker panel (see FIGURES 16a, b and σ, and 17a, b and c discussed below) . The mechanisms for extending, deploying, and supporting the sun blocker panel may involve various techniques and devices that are well known in the current state of the art of the design of mechanisms for spacecraft. For example the techniques and devices employed could involve mechanisms constructed from well known device types such as: hinges, flaps, slides, spring motors, wax motors, detentes, cable/bolt cutters, split nut releases, pin pullers, hook and pin releases, etc. Alternatively, so-called "active" devices such as electrical motors may be used at the discretion of the spacecraft designer. For example, one or more electrical motor (for example a stepper drive motor) could be employed to produce the motions resulting in extension (and possibly also retraction) of the sun blocker panel. Such active control could be utilized to facilitate certain operations of the spacecraft, for example station-keeping and attitude control operations for which displacing the sun blocker panel from the exhaust plume fields of rocket thrusters would be beneficial. The present invention device is applicable to spacecraft other than those spacecraft, like the model spacecraft for example, which operate in the low- inclination or equatorial orbits that have been described thus far herein. It is applicable to the broad class of spacecraft for which the solar illumination (insolation) is incident at low angles relative to the planes of the surface(s) of their thermal-radiator surface(s) . A certain subset of spacecraft belonging to the set of spacecraft that are well known in the space industry as "sun synchronous" fulfill this requirement for low solar incidence angles on at least one thermal radiator surface; and the present invention is applicable to them. Within this subset of sun synchronous spacecraft is an even smaller but well known subset comprised of those spacecraft that operate in orbits with low orbit-Sun angles and in which the thermal radiator surfaces are utilized while oriented close to parallel to the orbit plane. A sun ray blocker device according to this invention is applicable to those spacecraft, to provide them with a shaded, benign, and desirable thermal environment for their thermal-radiator surfaces basically by protecting them against direct solar heating. Note, however, that when the angle of incidence of direct sunlight on the thermal radiator surface is zero the sun ray blocker device is unnecessary. Heretofore the structural configuration and orientation of spacecraft to which the current invention device is applicable have mainly been described with reference to three-axis-stabilized Earth-pointing spacecraft for operating in low inclination or equatorial orbits, like the model spacecraft for example. The fundamental difference between those preceding descriptions and the structural configurations and orientations of the sun synchronous spacecraft to which the current invention device is applicable stems from the orientation of the orbit with respect to the axis of rotation of the planet. Within the space industry, sun synchronous orbits are widely referred to as being "polar", since the orbit plane of a sun synchronous orbit, around Earth and Mars at least, lies within several degrees of the axis of rotation of the planet; and therefore nominally includes the planetary poles. Therefore, for the aforementioned particular subset of sun synchronous spacecraft to which the current invention device is applicable, the panels and the radiator surfaces on them that are thermally protected by the current invention device are generally not, strictly speaking, "north" and "south" panels. However, herein the terms "north" and "south" are occasionally used for convenience to indicate the panels that are thermally protected by the sun blocker panel on spacecraft in sun synchronous orbits as well as on spacecraft like the model spacecraft, for example, in (nominally) equatorial orbits. The rationale is that in the particular, suitable, well known, and currently populous, aforementioned, subset of sun synchronous spacecraft the planes of thermal radiator panels shielded by the present invention device are also approximately perpendicular to the axis of the orbit (as for spacecraft like the model spacecraft in its orbital configuration and orientation) . For both these types of spacecraft we could instead meaningfully refer to the protected panels and radiator surfaces as "pitch-axis" or "orbit normal" panels and surfaces, because the pitch axis of the spacecraft (which is parallel to the orbit normal) is nominally/approximately perpendicular to them and thereby defines their orientation. Depending upon the requirements of the propulsion subsystem and/or the attitude control subsystem of the spacecraft, the spacecraft designer may elect to provide only one sun ray blocker device, i.e. on only one of the two sides of the spacecraft that face approximately along the pitch axis (e.g. on the north or the south panel for the model spacecraft) . In any particular application there may be a preference for one side of the spacecraft over the opposite side because of other system requirements. For example, in a potential embodiment of the present invention device on a particular current design of geostationary spacecraft, the south side is preferred because of field of view requirements for attitude-control thrusters on the north side. Again, if the spacecraft designer elects to do so, sola cells can be mounted onto the external surfaces of the sun blocker panel to provide additional power to the spacecraft. BRIEF DESCRIPTION OF THE DRAWINGS The present invention should be more fully understood when the specification herein is taken in conjunction with the drawings appended hereto showing exemplary embodiments of the invention wherein: FIGURE l is a simplified perspective view of a prior art three axis stabilized Earth-pointing geosynchronous spacecraft; FIGURE 2 shows an east-panel based view of the prior art spacecraft illustrated in FIGURE 1 orbiting in a low inclination or an equatorial orbit; FIGURE 3a shows a north-panel based, top view of the prior art spacecraft illustrated in FIGURE 1 orbiting about the Earth at different times of the day, and FIGURE 3b illustrates orbit-plane based views of that spacecraft at its noon, 6 a.m. , and midnight positions and also establishes sun angles for different seasons of the year; FIGURES 4a and 4b show the variation in the solar- incidence angle on the north and south panels, respectively, of the prior art spacecraft illustrated in FIGURE 1 orbiting Earth, through one calendar year; FIGURE 5 illustrates a perspective view of a spacecraft configuration according to the present invention, based on the prior art spacecraft illustrated in FIGURE 1; FIGURES 6a, 6b and 6c illustrate top views of a present invention arrangement as applied to the prior art spacecraft illustrated in FIGURE 1. The views shown are simultaneously parallel to both the orbit-plane and the plane of the solar cell panels and the sun blocker panels of the sun ray blocker device. Hereinafter this view direction is also referred to as "top view". FIGURES 6a, 6b and 6c show that as the spacecraft revolves around the orbit the earth panel always faces the Earth, and the cell-side of the solar array panels together with the front (sunward) sides of the sun blocker panels of the sun ray blocker devices always face the Sun; FIGURES 7, 8 and 9 illustrate top views of present invention devices utilizing different attachment arrangements; FIGURES 10a, 10b and 10c illustrate portions of top views of one of the solar array assemblies of a prior art spacecraft before, during, and after its deployment; FIGURES Ila, lib, lie, 12a, and 12b show, in top view, aspects of the deployment and the function of present invention devices as applied to the prior art spacecraft shown in FIGURES 10a, 10b, and 10c; FIGURES 13 and 14a show partial top views of two alternative present invention devices; and FIGURE 14b shows a partial back (anti-sun) side view of the arrangement shown in FIGURE 14a; FIGURES 15a and 15b show partial top views of an alternative present invention device in its fully deployed and partially deployed configurations, respectively; FIGURES 16a, b, and c show a view of an alternative present-invention device from the front (sunward) direction with the sun blocker panel fully deployed, and two views from a direction orthogonal to both the front view and the (previously defined) top view directions with the sun blocker panel folded and deployed, respectively; FIGURES 17a, b, and c show a different alternative present-invention device in the same views and deployed states as those shown in FIGURES 16a, b, and c; FIGURE 18 shows a further embodiment of the invention; FIGURE 19 shows another embodiment of the invention; FIGURE 20 shows another embodiment of the invention; FIGURES 21 to 24 and 26 show another embodiment of the invention; FIGURE 25 shows details of the embodiments of FIGURES 21 and 24 and 26; FIGURES 27 to 30 show another embodiment of the invention; and FIGURES 31 and 32 illustrate alternative shapes for sun blocker panels used in the present invention. Referring now to FIGURE 1, there is shown an oblique view of a fully deployed (i.e. fully unfolded from its launch configuration) spacecraft (or satellite) 1, like the previously described model spacecraft for example, which is represented by a main body 10 which contains six external panels: 11, 12, 13, 14, 15 and 16, a group of antenna reflectors 20, 21, 22 and 23, and two solar array assemblies, consisting of two solar arrays (one or more solar cell panel) 100 and 101 and their supports 100a and 101a by which they are connected to the main body 10, which are extended northward and southward from the main body out of the north and south panels 11 and 12, respectively. The number of antenna reflectors is driven by the need of the telecommunications application and is a matter of design. In this example, four reflectors are shown and are represented by two deployable large reflectors 20 and 21 mounted on east and west panels 15 and 16, respectively. Two non-deployable reflectors 22 and 23 are mounted on nadir panel 14. While orbiting in a low inclination orbit about Earth, the spacecraft is controlled in such a way that the earth or nadir panel 14 is pointing in the general direction of the center of the Earth, thus allowing the antenna reflectors to perform telecommunications functions with Earth. Opposite to the earth panel 14 is the zenith panel 1 . The solar arrays 100 and 101 may contain multiple panel elements (typically two to eight or more on each side - a four panel-element example is shown in FIGURE 1) or may contain as few as one panel element. However, usually solar arrays that are comprised of multiple solar cell panels are utilized, in order to provide sufficient electrical power for the spacecraft's use. The size and number of the solar cell panels is driven by mission power requirements, and is constrained by, among other factors, the capability of the attitude control subsystem to maintain pointing stability and also by the capability of the thermal control subsystem to manage the heat dissipated on board the spacecraft. Once the size and number of the panel elements is defined, generally it is desired to maximize the electrical power generated by the solar cells which are mounted on one side of the array panels by facing the cell side of the array toward the sun as directly and as long and continuously as possible. ith spacecraft main body 10 maintaining its earth panel 14 pointing to the Earth continuously, the line between the spacecraft and the Sun will cone around the north- south axis of a spacecraft like the model spacecraft once every orbit, making the Sun appear to circle about the main body 10 as it does so. In order to maintain both solar arrays of a spacecraft like the model spacecraft pointing directly to the Sun they are driven by motor systems which rotate the arrays about the north-south axis, as indicated by the arrow R in FIGURE 5 with respect to the main body 10 at a speed such that the cell side of the array always faces the Sun while the spacecraft orbits the Earth, i.e. the solar arrays rotate about the north-south axis sun synchronously with the Sun to achieve optimum sun exposure for maximum power generation. Reference is made to FIGURE 2, a top view of p ior art spacecraft or satellite 1 of FIGURE 1, wherein the aforesaid seasonal exposures are illustrated. (Parts identical or very similar to those in FIGURE l are identically numbered throughout the FIGURES herein and are not all repeated, to reduce redundancy. This applies to all of the following FIGURES that illustrate the same spacecraft or the same parts or components, or ones very similar.) The north panel ll and the south panel 12 (FIGURES l and 2) are maintained oriented parallel to the orbital plane of the satellite, which is co-planar or nearly co-planar with the equatorial plane of the Earth. While the spacecraft is orbiting the Earth, these panels (11 and 12) will not receive daily solar input like the other panels (earth panel 14, zenith panel 13, east panel 15 and west panel 16) . Those two panels 11 and 12, however, will be subjected to direct solar heating on a seasonal basis, at incidence angles which will peak at 23.5 degree at the northern summer- and northern winter- solstices respectively, as shown. FIGURE 3a shows a north-based top view of a spacecraft 1 orbiting Earth at different local times of day and illustrates the constancy with which the nadir panel 14 faces Earth 300 throughout the orbital revolutions. (The solar cell panels are shown edge-on out of the paper. ) FIGURE 3b shows a partial side view of the spacecraft 1 of FIGURES 1 and 2 at midnight, 6 a.m. , and noon orbital positions, and also the approximate sun angles at the northern summer and northern winter solstices at midnight and noon. FIGURES 4a and 4b show the profile of the solar incidence angle on the north and south panels, respectively, such as panels 11 and 12 of spacecraft 1 shown in FIGURE l, for one calendar year. It can be seen from FIGURES 4a and b that sunlight is incident on each of the north panel and south panel for a portion of the calendar year. These periods are nominally 21 March through 21 September for the north panel, and 21 September through 21 March for the south panel. Therefore, the sun ray blocker devices of the current invention perform their shading functions for their respective radiator panels for those periods only. FIGURE 5 illustrates one preferred embodiment of the current invention, which eliminates or greatly reduces the seasonal solar input on the north and south panels 11 and 12, thus providing more efficient thermal radiators for the spacecraft. In this present invention embodiment, the sun ray blocker devices are comprised of two sun blocker panels 111 and 112 and mounting, supporting, and deployment mechanisms by means of which the blocker panels are integrated with and deployed with the structures and mechanisms that support and cause the solar array to rotate. The radiators on the north and south panels 11 and 12 have dedicated sun ray blocker devices 111 and 112 attached to the north and south array assemblies 100 and 101, respectively, as shown in FIGURE 5. After the thus modified spacecraft l has been launched into the operational orbit and its appendages have been fully deployed, the sun blocker panels 111 and 112 will achieve their final positions in front of the cell side of the solar arrays with their surfaces more or less parallel to the plane of the solar arrays. The south blocker device 112 is positioned such that during the time between the northern autumnal and northern spring equinoxes, when otherwise there would exist a potential for solar heating of the south panel 12, the south blocker device 112 will cast a shadow over the south panel 12 thereby eliminating the potential for such solar heating. The north blocker device 111 performs a similar function relative to the north panel 11 during the time between the northern spring and northern autumnal equinoxes. When the solar array assemblies 100 and 101 are maintained directly sun pointed, by virtue of their being rotated, the sun blocker devices will likewise be maintained directly sun v pointed and thereby interposed between the Sun and the north and south panels that they shade. The materials used for the sun blocker panels ill and 112 are selected to minimize the heat transferred from their sun facing surfaces ilia and 112a to their anti-sunward surfaces 111b and 112b. This may be achieved by including insulating material(s) and constructions in the composition of the sun blocker panels. For example, the panels may include known thermally insulating materials and assemblies of materials, such as multilayer insulation (MLI) blankets which utilize layered films of metallized Mylar separated by fabric netting. These materials and constructions are well known in the space industry and have typical heat resistance values of 0.007 to 0.01 Watt/deg.C/sq.in. The sun blocker panels of the present invention device will generally experience a sizable temperature difference, for example possibly greater than 100 degree c, between surface Ilia and surface 111b and between surface 112a and surface 112b when the satellite is in its normal orientation in the mission orbit, except when the spacecraft is passing through the Earth's shadow. To obtain the maximum sun blocking effect, the panels of the sun ray blocker devices 111 and 112 are configured (sized, oriented, and positioned) in such a way that at the summer and winter solstices, when the Sun is about 23.5 degree from the orbit plane, the sun blocker devices will cast shadows that entirely cover the radiator surfaces on their respective thermal radiator surfaces on the spacecraft panels ll and 12. Accordingly, if the radiator surfaces are rectangular the shadows must be at least as wide as the diagonals of the rectangles. FIGURES 6a, 6b and 6c show top partial views of a present invention arrangement as the spacecraft orbits Earth and the main body 10 is rotated at the orbital rate so that the earth panel 14 always faces Earth, and the sun blocker panels 111 and 112 always face the Sun (which is at the left in the FIGURES) . These FIGURES are drawn in the inertial frame of reference of the solar array assemblies 100 and 101. Thus, if one were to stand on either of the solar array assemblies 101 and 102 one would see main body 10 rotate one revolution per orbital revolution around the Earth. FIGURE 7 is a top partial section view showing more details of a present invention spacecraft. In this context the phrase "top view" denotes a view parallel to the planes of the sun blocker panels 111 and 112 and also parallel to the orbit plane. Note that in Figure 7 through Figure 15b various examples of embodiments of the present invention are depicted together with generic partial views of a spacecraft main body and a solar array assembly (labeled 400 and 408, respectively, later in FIGURES 10 and FIGURES 11) . Additionally, FIGURES 8 and 9 show alternative embodiment arrangements in top partial section views. In FIGURE 7, the spacecraft has main body 10, north panel 11, and solar cell panel support 223 with attached solar cell panel 225. In this case, there is a connecting solar array boom-and-yoke 219 and hinges at hinge points 221 and 227. Together this solar cell panel support 223, a solar cell panel 225, a solar array boom- and-yoke 219, and the hinges at hinge points 221 and 227 comprise part of a solar array assembly. The solar array boom-and-yoke 219 fold forwardly against north panel 11 and the solar cell panel support 223 together with the solar cell panel 225 folds down at hinge point 227 in an accordion-like fashion for launching. During launch, ascent, and orbit achievement the solar array assembly is in its folded-closed configuration. After achievement of the mission orbit it is electro-mechanically and/or mechanically deployed (unfolded) to allow the solar cells to be maintained directly sun-pointed. Attached to solar cell panel support 223 is a two-section connecting arm having a short inner portion 209 and an outer portion 207 connected by hinge(s) at hinge point 215. The anti- sunward side 111b of sun blocker panel ill is connected to outer arm portion 207 by hinge(s) at hinge point 203. Optional solar cells 201 are functionally positioned on the sunward surface Ilia of the sun blocker panel 111. Hinge points 203 and 215 provide for folding of the solar blocker panel 111 and its hinged arm 207 against the solar cell panel 225 in a compact and stiff configuration suitable for launch and subsequent deployment. The electromechanical and/or mechanical designs and methods for deploying (opening) and closing solar array assemblies are commonly used in contemporary spacecraft. The same or similar mechanisms are used to deploy the sun ray blocker devices of the present invention. These mechanisms and methods for deployment and closing are well within the skills of the artisan. In FIGURE 7, there is an imaginary plane 250 extending off the surface of north panel 11. In its deployed configuration the sun blocker panel ill may touch or extend through this imaginary surface, and consequently may provide additional shading for the earth, west, zenith and east panels as they rotate with respect to the Sun. FIGURE 8 shows an alternative embodiment where sun ray blocker device 271 does not intersect imaginary plane 250. Further, it has a single connecting arm 205 with hinge points 203 and 217 at opposite ends to form an assembly and is connected directly to the substrate of solar cell panel 225. It may be folded and stowed for launch and deployed or unfolded in orbit in a similar way to the sun ray blocker device in FIGURE 7. In FIGURES 7 and 8, the sun ray blocker devices cast their shadows over the major part of the outer surface of north panel 11 and, in these embodiments, completely shadow that surface during the times when otherwise they would be exposed to the Sun. Further, the solar cells 201 may be inσluded to produce additional solar power for the spacecraft. In FIGURE 9, identical parts to FIGURES 7 and 8 are identically numbered. Sun ray blocker device 301 is connected directly to solar cell panel support 223 by hinge(β) at hinge point 309 so as to fold over up-close against solar cell panel 225 in the launch configuration. In this embodiment, sun ray blocker device 301 is not parallel to the solar array, yet still effectively shades north panel 11. FIGURES 10a, 10b and 10c depict a typical prior art sequence of deployment of a solar array assembly, which is part of the transformation of the spacecraft from its launch configuration to- its configuration for normal operations in orbit. For simplification in this document, only one (the north) solar array assembly is shown in the FIGURES. These particular FIGURES show a satellite with a main body 400, and a solar array assembly 408 comprised of four solar cell panels, with solar cell surfaces 400a, mounted on solar cell panel supports 408 which are interconnected by hinges at three hinge points 403, 404, and 405 and connected to the main body 400 by a single boom 419 and hinge(s) at hinge points 401 and 402. FIGURE 10a depicts the solar array assembly folded and stowed for launch. FIGURE 10b depicts it in the process of being deployed (unfolded) . Figure 10c depicts its fully deployed state. If a multiple-arm boom design is desired by the spacecraft designer, various embodiments can be designed to satisfy performance requirements using greater numbers of arms and hinge points. FIGURES ila, lib and lie illustrate the deployment sequence of one possible design embodying the present invention. Components in FIGURES ila, lib, and lie that are identical to components in FIGURES 10a, 10b, and 10c are numbered identically to their identical parts, in addition to the prior art solar array assembly that was previously depicted in FIGURES 10a, 10b and 10c, FIGURES ila, lib, and lie also depict the present invention sun blocker panel 411 connected to the solar array boom 419 by an arm 430 and hinges at hinge points 406 and 407. Alternatively, by design the sun blocker panel 411 could be hingedly attached via the arm 430 to a convenient different location on the solar array assembly. FIGURE Ila depicts the solar array assembly and the sun ray blocker device folded and stowed for launch, FIGURE lib shows them partially deployed (unfolded) . FIGURE lie depicts their fully deployed state.FIGURES 12a and 12b show sun blocker panels which are not parallel to the plane containing the solar cell panels yet which still provide proper shading of the north or south panel. Components in FIGURES 12a and 12b and subsequent figures that are identical to components that appear in previous figures are numbered identically with their corresponding or very similar components or are left un-numbered to avoid unnecessary repetition. Alternatively, by design the sun blocker panel 111 could be hingedly attached via the arm 430 to a convenient different location on the solar array assembly. FIGURE 13 depicts yet another alternative embodiment of the present invention. The sun blocker panel 511 is connected to the solar array boom 219 by hinge(s) at hinge point 507 for its stowing folded and subsequent deployment. FIGURES 14a and b show an arrangement similar to that in FIGURE 13, with identical parts identically numbered, however, more hinges at hinge points 606 and 607 are used with blocker panel 611 as required by design for folding the panels prior to deployment. FIGURE 14b represents a partial view of the anti- sun side of the spacecraft looking toward the Sun (i.e. a side view relative to the top view shown in Figure 14a) . FIGURES 15a and 15b show one embodiment in which sun blocker panel 811 utilizes separate active motors 306 and 307 which are used to actively deploy and/or retract the sun ray bloαker device. This arrangement allows satellite operators to use deployment motors that are separate from the solar panel deployment motors so as to permit them to retract the sun blocker panels to prevent their interference, if any, in satellite operations such as in the use of propulsion systems during spacecraft performance of station keeping or attitude control maneuvers. In some spacecraft designs the required size (dimensions and/or area) of a sun blocker panel in its fully deployed configuration may exceed the constraints of its "launch envelope" i.e. the constraints of the maximum-allowable space allocated to the sun ray blocker device in the launch configuration of the spacecraft when the solar array and the sun ray blocker device are in their launch configuration. Therefore, for compatibility with the constraints of the size of the corresponding launch envelope it may be necessary for the sun blocker panel of the sun ray blocker device to be comprised of several (i.e. more than one) pieces, instead of being one single integral piece, which are folded together in the launch configuration and are subsequently deployed (unfolded) in orbit to form effectively one continuous sun blocker panel. FIGURES 16a, 16b and 16c, and FIGURES 17a, 17b and 17c, respectively depict two examples from the many possible designs for sun blocker panels which fold and deploy. Parts a, b, and c of the FIGURES 16 and 17 show each of the two designs in a view from the front (Sun) direction with the sun blocker panel fully deployed, and two views from a direction orthogonal to both the front view and the top view directions with the sun blocker panel folded and deployed, respectively. (As defined earlier herein the phrase "top view" denotes a view that is simultaneously parallel to the plane of the sun blocker panels 921 or 951 and the orbit plane.) This allows the sun ray blocker device to increase its dimensions using hinge(s) and/or strut(s) and/or flap(s) etc. at hinge points 925 and 927 or a slide-out design. Referring collectively to all FIGURES 16, sun blocker panel 921 has a center section 923 with hinge(s) and/or strut(s) and/or flap(s) etc. at hinge points 925 and 927 and outer, swing up panels 929 and 931 which may be designed to deploy (swing up) automatically. In all FIGURES 17, sun blocker panel 951 has main section 953 with slide-out extensions 955 and 957 that may be designed to deploy (slide out) automatically. (Automatic hinging and automatic sliding or telescoping is well within the purview of the artisan in the spacecraft industry and need not be further elaborated upon herein. ) The embodiments of the present invention device illustrated in FIGURES 18 through FIGURE 30 are as generally applicable as the other embodiments described herein. However, they also function efficiently in cases where a sun blocker device cannot be attached to an axle located near the center (1811, 2123, 2722) of an associated thermal radiator surface (1804, 2121, 2721) . One such case is that in which the axis of rotation (1803, 2131, 2701) of a solar array assembly extends outward from the associated thermal radiator surface (1804, 2121, 2721) at a location that is signi icantly offset from the center (1811, 2123, 2722) of the thermal radiator surface. In that case, designs with an attachment arm of fixed length between the sun blocker panel and the solar array axis could be unsuitable, because the motion of the sun blocker panel about the center of the thermal radiator surface would be eccentric. .Another such case is that in which there are stay- out zones inboard of the periphery of the associated thermal radiator surface - through which objects such as a supporting boom (for example for a sun blocker panel) are not allowed to pass. This could be the case, for example, when certain attitude- or orbit- control thrusters (1810) are also located on the panel upon which the thermal radiator surface (1804) is located. The arrangements illustrated in FIGURES 18 through FIGURE 30 may be employed to overcome these constraints, whilst still maintaining a sun blocker panel at a substantially uniform distance from the center of the associated thermal radiator surface. Selection between the embodiments shown in FIGURES 18-30 for any particular application may involve trade-offs between many additional performance-requirements of the spacecraft- system, including for example: mass, strength, stiffness, flatness, circularity, simplicity, and reliability. Figure 18 shows an embodiment of a sun blocker device in which the sun blocker panel 1800 is mounted on a carriage 1801 with wheel-sets or bearing-sets 1808 and 1830 by means of an attachment arm 1805 in which the carriage may be driven around a closed rail 1802, the carriage 1801 being attached to the rail 1802 by rolling or sliding means that also react against and thereby limit rotations of the carriage 1801 (and thereby the sun blocker panel) about axes passing through the points of contact of the carriage and the rail, in one of many potential embodiments, for example, this may be achieved using wheel or bearing sets 1808, 1830 that are adequately spaced both along-track and cross-track on both sides of the rail 1902, and which are also cambered at an adequate angle to the plane of the baseplate. Attached to the carriage is at least one generally-radial boom or strut 1805, an outer end of which is attached to the sun blocker panel at hinge point 1812 and an inner end of which is attached to the carriage 1801 at hinge- point 1813, and the carriage is rollingly or slidingly mounted on the rail 1802 by the wheels or bearings 1808 and 1830. At least one of these wheels 1830 or bearings is provided with a motor to drive the carriage along the rail 1802, for example by friction, or by the engagement o ' a toothed wheel or a worm-drive in a rack. Electrical power may be supplied to the motor, via brushes for example. The attachment arm 1805 and the sun blocker panel 1800 can be folded at the hinge points 1812 and 1813 to achieve a stowed configuration of the sun blocker device for launch, during which the folded device may be temporarily caged securely for proper management of launch-induced dynamic environments and loads. The sun blocker panel may be further folded for launch as illustrated in FIGURES 16 and 17. Following launch the attachment arm 1805 and the sun blocker panel 1800 can be deployed for subsequent operation in orbit, including sun-tracking travel around rail 1802. It will be appreciated that the rail 1802 need not be circular as shown in FIGURE 18, but in the case of a significantly rectangular thermal radiator surface, for example, the rail could be elliptical and in either case may be diverted to avoid obstacles mounted on the spacecraft. Alternatively, as illustrated in FIGURE 19 the attachment arm 1805 could be mounted on a solid rotatable wheel instead of on a carriage and rail. In the embodiment illustrated in FIGURE 19 the wheel is a ring or annulus 1902 floating in circumferentially located bearing-sets 1903 and controlled and driven by a motor 1930 mounted on the baseplate under 1804. In a similar alternative embodiment, illustrated in FIGURE 20, a carriage 2001, similar to that provided in the embodiment illustrated in FIGURE 18, is provided; but i ithis embodiment the carriage 2001 is driven around a closed rail 2002 not by a motorized wheel, but by an endless belt 2003, chain, or cable attached to the carriage 2001, the belt being driven by a motor 2030 that is, mounted to the baseplate under 1804 and which engages the belt 2003, chain, or cord. A tensioning device 2040 is .also provided to engage the belt 2003 and tension the belt while not impeding the passage of the carriage around the rail 2002. Again, as in the embodiment illustrated in FIGURE 18 the rail 2002 need not be circular. Figures 21 through 30 illustrate embodiments in which a sun blocker panel is mounted on the spacecraft via an attachment arm from an axis 2131, 2701 that is offset from the center 2123, 2722 of an associated thermal radiator surface. The axis 2131, 2701 could be concentric with or identical to the axis of rotation of a solar array assembly. Figures 21 through 26 illustrate an alternative embodiment in which a sun blocker panel 2100 is attached to an axle at axis 2131. The axle may be either concentric with or identical to an axle of a solar array assembly. The sun blocker panel is attached to the axle by an articulated attachment arm 2130 that included three articulated portions 2132, 2134, 2137. The inner end of the inner-portion 2132 is fixed radially to the axle at axis 2131. The middle-portion 2134 is pivoted at its inner end to inner-portion 2132 at pivot-point 2133, and the outer end of middle-portion 2134 is pivoted to the inner end of outer-portion 2137 at pivot point 2135. At its outer end the outer-portion is attached to the sun blocker panel at hinge point 2138 and near its inner end the outer-portion is hinged at hinge point 2136 to allow folding and stowing for launch followed by deployment in orbit. As depicted in FIGURE 21 through FIGURE 26 the inner-portion 2132 and the outer portion 2137 of the attachment arm 2130 turn anti-clockwise at the same rate, the outer-portion 2137 carrying the sun blocker panel with it, whereas the middle portion 2134 rotates clockwise at the same rate. The length of the inner-portion 2132 is approximately equal to the offset of the axis of rotation 2131 from the center 2123 of the thermal radiator surface. In principle the length of the middle-portion 2134 may be longer or shorter than the length of the inner portion 2132. However, in the case that the axis of rotation 2131 is occupied by an obstruction such as the axle of a solar array assembly then the middle- portion 2134 must be shorter than the inner-portion 2132 for clearance of the solar array axle at axis 2131, as can be seen in FIGURE 23 in which the attachment arm 2130 is approaching its closest to the axle at axis 2131. By articulating the articulated portions through rotation of the arm 2130 about the axis of rotation 2131 the sun blocker panel can be maintained at a substantially constant distance from the center of the associated thermal radiator surface 2121, to describe a substantially circular path 2140 around the spacecraft. It will be evident that in the case of a thermal radiator surface that is significantly far from being radially symmetric the length of the articulated portions of arm 2130 could be adapted to achieve a wide range of desired paths around the thermal radiator surface. As shown in FIGURE 21, the articulated portions 2132, 2134, 2137 are arranged to the full reach of attachment arm 2130 in a straight line when the sun blocker panel is passing a side of the thermal radiator surface furthest from the axis of rotation 2131. As shown in FIGURES 22 and 23 the attachment arm 21 0 has an effective length equal to the sum of the lengths of an outer 2137 and an inner 2132 articulated portion when the sun blocker panel 2100 is at an intermediate distance from the axis of rotation 2131. As shown in FIGURE 23, the effective length of the attachment arm 2130 is at its minimum when the sun blocker panel is at its closest to the axis of rotation 2131, at which point its length is equal to the sura of the lengths of the inner 2132 and outer 2137 portions less twice the length of the middle- portion. The inner articulated portion 2132 of the attachment arm 2130 rotates about axis 2131. The means of attachment of the inner articulated portion 2132 may be independent of a solar array axle along axis 2131, the inner portion 2132 then being mounted to a concentric tubular axis around a central solar array axle. Alternatively, the inner articulated portion may be fixed solidly to a solar array axle along axis 2131. In the illustrated embodiment, for a geostationary spacecraft for example the inner and outer articulated portions rotate anti-clockwise at one revolution per day, and the middle articulated portion rotates clockwise at one revolution per day. This rotational relationship may be achieved by diverse means, such as: separate motorized pivots at pivot points 2131, 2133, and 2135; or by a system of belt-linked pulley wheels at pivot points 2131, 2133 and 2135, driven by a single motor or by an axle along axis 2131. The articulated portions 2132, 2134, and 2137 may be sprung together, so that in a failure mode the attachment arm 2130 automatically extends to its greatest length. In that case, any failed pivot points can be made to fail free, for example using comraandable frangible-links in the associated pulley wheels or drive motor, allowing spring-driven extension of the arm 2130. FIGURES 25 and 26 illustrate a means of articulating the articulated portions 2132, 2134, 2137 with respect to each other, utilizing a driving force from a solar array axle at axis of rotation 2131. A cylinder 2501 is provided, mounted co-axial with the solar array axle at axis 2131 but fixed to a base panel 2121. The inner articulated member 2132 is fixed to the solar array boom -Sl¬ at axis 2131 so that the inner articulated member 2132 rotates at the same rate as the solar array boom at axis 2131. A middle articulated portion 2134 is pivotβdly fixed to an outer end of the inner articulated portion 2132 at a pivot point 2133 and a pulley wheel 2502 of the same diameter as cylinder 2501 is fixed to an inner end of the middle articulated portion 2134. A toothed belt 2506 is looped around the cylinder 2501 and the pulley wheel 2502 so that as the solar array shaft 2131 and the inner articulated portion 2132 rotate anti-clockwise in a direction of the arrow 2507, the toothed belt 2506 causes the pulley wheel 2502 and the middle articulated portion 2134 to counter-rotate at the same rate in the direction of arrow 2508. As shown in FIGURE 26, a second equal sized pulley wheel 2601 is fixed to an outer end of the inner articulated portion 2132 on a side of the inner articulated portion 2132 opposite to that on which the cylinder 2501 is fixed to base panel 2121, and a third equal sized pulley wheel 2602 is fixed to an inner end of an outer articulated portion, such that the third pulley wheel 2602 and the outer articulated portion 2137 are together pivotedly attached to the outer end of the middle articulated portion 2134. A second toothed belt 2603 loops around the second pulley wheel 2601 and the third pulley wheel 2602 so that as the middle articulated portion 2134 rotates in the direction of arrow 2508 the toothed belt 2603 causes the outer articulated portion 2137 and the third pulley wheel 2602 to counter-rotate at the same rate in the direction of arrow 2604. Thus, the outer articulated portion 2137 rotates in the same sense as the inner articulated portion 2132 whereas the middle articulated portion 2134 counter-rotates. In a further embodiment illustrated in FIGURES 27 through 30, a sun blocker panel 2700 is attached to an axle 2701 of a solar cell array by means of a scissor attachment arm 2730. The scissor arm is comprised of a first articulated arm 2704, 2708 and a second articulated arm 2705, 2709, comprised of inner articulated portions 2704, 2705 and outer articulated portions 2708, 2709 respectively. The inner articulated portions are connected by hinges at hinge points 2702, 2703 to the solar array boom 2701 respectively and the outer ends of the outer articulated portions 2708, 2709 are connected by hinges 2710, 2711 to the sun blocker panel 2700 such that when the articulated arms are extended to the full length they are still not parallel to avoid their locking up. A lanyard 2712 is located in between the articulated arms 2704, 2708 and 2705, 2709 and extends between the sun blocker panel 2700 and the solar array axle 2701. The inner articulated portions 2704, 2705 and the outer articulated portions 2708, 2709 are sprung at hinge points 2706 and 2707 so as to automatically extend the articulated arms 2704, 2708 and 2705, 2709 to their full extent as limited by the lanyard control. The articulated arms 2704, 2708 and 2705, 2709 thereby form a parallelogram, the shape of which may be controlled by retracting or deploying the lanyard 2712. Alternatively the shape of the parallelogram could be controlled by motorized hinges, or alternatively by a retractable and deployable lanyard between hinge points 2706 and 2707 with sprung hinges 2702, 2703, 2710 and 2711 instead of at 2706 and 2707. In the embodiment of the invention illustrated in FIGURES 21 through 24, the distance of the sun blocker panel 2700 from the solar array boom 2701 can be varied as the sun blocker panel 2700 rotates about the solar array boom 2701 to maintain the sun blocker panel at a constant distance from the spacecraft as illustrated by the path 2712. The embodiments described in FIGURES 18 through 20 have the advantage that the attachment arm does not obscure thrusters 1810 present on the face of the spacecraft, that the sun blocker panel shades. A sun blocker panel 3100, 3200 is not necessarily rectangular in shape. As shown in FIGURE 31, the sun blocker panel 3100 has trapezoidal first-and second- extensions 3101, 3102 hingedly attached to a main body 3103 of a sun blocker panel 3100. The first extension 3101 is extended by unfolding the extension through rotation in the direction of the arrows 3104 and the second extension is extended by unfolding the extension through rotation in the direction of the arrows 3105 from a position flat against the main body 3103. As shown in FIGURE 32 a rectangular main body 3202 of the sun blocker panel 3200 may have substantially triangular extensions 3201, 3202 which may be extended and retracted from the main body by sliding translation of the extension 3201 in the direction of double-handed arrow 3204 and unfolding the extension 3202 in the direction of arrows 3205. The descriptions of designs for the structural support and the deployment of sun ray blocker devices written herein are examples from thousands of possible structural support and deployment designs which can be used for this purpose and are within the scope of the present invention. This paragraph describes an example to demonstrate the geometrical approach to calculating the dimensions of a sun blocker panel for providing total shadow coverage to a quasi-rectangular shaped radiator surface. The example used is that of a radiator surface on a north or south panel of a geostationary spacecraft, like the previously defined "model" spacecraft for example, at the summer or winter solstice, when the incidence angle of the sun's rays (measured from the plane of the benefited thermal radiator surface) is at a maximum, using a quasi- rectangular (for this simple illustration at least) shaped sun blocker panel whose plane is perpendicular to the plane of the associated thermal radiator surface (referring to FIGURE 12a, angle 501 is then 90 degree) . For this example, take the north or south radiator- surface of the spacecraft to be rectangular, of length and width A and B, respectively. Then, the length and width dimensions, L and W, respectively, of the sun- exposed surface of the fully-deployed sun blocker panel (which is shown in FIGURES 16a and 17a) should be as follows: L is greater than or equal to V(A 2 +B 2 ) , and W is greater than or equal to 0.435x /(A 2 +B 2 ) based on the corresponding orbit-sun angle of 23.5 degree. However, if only a portion of the surface area on the north or south panel needs to be shadowed, i.e. high heat-dissipating equipments were to be mounted in certain localized areas of the north or south panel, the sun ray blocker device can be tailored to shade only those areas and may accordingly be smaller. In addition, if a sun blocker panel whose plane is not perpendicular to the plane of the associated thermal radiator surface were to be selected by the spacecraft designer, the minimum value of the width, W, may be greater or less than 0.435x V(A 2 +B 2 ) depending on the size of angle 501 in FIGURE 12a. If angle 501 is greater than 90 degree, W may be greater than 0.435χ (A 2 +B 2 ) ; if it is less than 90 degree, W may be less than 0.435x /(A 2 +B 2 ) . If additional shading to the other four panels, earth, zenith, east and west panels, is desired, the width (W) of the sun blocker panel can be increased to extend past the imaginary plane 250 toward the center of the satellite as shown in FIGURE 7. Thus, by the foregoing descriptions contained herein it can be seen that by virtue of the present invention losses in the efficiency of the cooling- of the thermal radiator panels of a spacecraft caused by solar heating can be eliminated or minimized via various sun blocking arrangements. Obviously, numerous modifications to and variations on the present invention are possible in light of the above teachings. For example, as a practical matter, a designer might counterweight or counterbalance the rotating axles or arms to overcome the weight imbalance caused by sun ray blocker devices of the present invention without exceeding the scope of the present invention. It is therefore understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described herein.

Claims: CLAIMS 1. A spacecraft for orbiting a sunlit celestial body (300) , the spacecraft including a thermal radiator surface (11,12,1804, 2121, 2721) for radiating heat from the spacecraft into space, and a sun ray blocker device (111,112, 141, 271, 301, 411, 511, 611, 811, 921, 951,1800,2100, 2700) mounted on said spacecraft for shielding said thermal radiator surface (11,12,1804,2121,2721) from rays of sunlight, characterised in that said sun ray blocker device is locatable for placing in shadow substantially the whole of the thermal radiator surface (11,12,1804,2121,2721) from sunlight without substantially impeding thermal radiation from said thermal radiator surface (11,12,1804,2121,2721) into space. 2. A spacecraft as claimed in claim 1, wherein an effective radiation view factor for thermal radiation from the thermal radiator surface (11,12,1804, 2121,2721) to deep space is significantly greater than corresponding geometrical radiation view factor to deep space, and in particular, for a geostationary spacecraft where the geometrical radiation view factor to deep space is 0.65, the effective radiation view factor to deep space is at least 0.87. 3. A spacecraft as claimed in claims l or 2, wherein the sun ray blocker device includes at least one sun blocker panel (111, 112) having a sun-facing surface (Ilia, 112a) and an opposed anti-sun-facing surface (111b, 112b) , wherein the sun-facing surface (Ilia, 112a) is thermally insulated from the opposed surface (111b, 112b) . 4. A spacecraft as claimed in claim 3, wherein the sun- facing surface (Ilia, 112a) is thermally insulated from the opposed surface (111b, 112b) by a multi-layer insulation blanket. 5. A spacecraft as claimed in claims 3 or 4, wherein the sun-facing surface (ilia, 112a) has a low solar energy absorptivity of less than 0.5. 6. A spacecraft as claimed in any of claims 3 to 5, wherein the sun-facing surface(Ilia, 112a) includes a solar cell panel for supplying electrical power to the spacecraft. 7. A spacecraft as claimed in any of claims 3 to 6, wherein the εun-facing surface (ilia, 112a) has a high thermal emissivity of higher than 0.7. 8. A spacecraft as claimed in any of claims 3 to 7, wherein the sun ray blocker device (111,112, 141, 271, 301, 411, 511, 611, 811, 921, 951, 1800, 2100, 2700) is moveable between a stowed, non-operative position and a deployed operative position. 9. A spacecraft as claimed in any claims 3 to 8, wherein the sun ray blocker device (111,112, 141, 271, 301, 411, 511, 611, 811, 921, 951,1800, 2100, 2700) includes an attachment arm (207, 205, 430, 230) for attaching the sun blocker panel (111, 112) to the spacecraft. 0. A spacecraft as claimed in claim 9, wherein the attachment arm (207, 205, 430, 230) is attached by a hinge means (406, 306) to the sun blocker panel (111, 112) and/or by a second hinge means(407, 507, 607, 307) to the spacecraft. 11. A spacecraft as claimed in any of claims 8 to 10, wherein the sun ray blocker device (111,112; 141; 271; 301; 411; 511; 611; 811; 921; 951) includes a motor for moving the sun ray blocker device (111,112; 141; 271; 301; 411; 511; 611; 811; 921; 951) between the stowed position and the deployed position. 12. A spacecraft as claimed in any of claims 3 to 11, wherein locating means are provided for locating the sun ray blocker device (111, 112; 141; 271; 301; 411; 511; 611; 811; 921; 951) with respect to the thermal radiator surface (11,12) which includes adjustment means to maintain the majority of the thermal radiator surface (11,12) in shadow irrespective of changes in the attitude and/or orbit of the spacecraft during normal operations. 13. A spacecraft as claimed in claim 12, wherein the adjustment means includes a variable length attachment arm (2130, 2730) or attachment of the sun blocker panel to the spacecraft. 14. A spacecraft as claimed in claim 13, wherein the attachment arm is a scissor arm (2730) . 15. A spacecraft as claimed in claim 13, wherein the attachment arm (2130) is formed of articulated portions (2132, 2134, 2137) which may be mutually articulated 5. A spacecraft as claimed in claim 13, wherein the transport means includes rail means (2002) and belt means (2003) connected to the carriage means (2001), the belt means being driven by drive means (2030) to move the carriage means along the rail means (2002) . 16. A spacecraft as claimed in claim 13, wherein the carriage means includes an annulus (1920) rotatable in a circular path defined by bearing means (1903) the annulus being driveable by drive means (1930)to move the carriage along the path defined by the bearing means. 17. A spacecraft as claimed in any of claims 12 to 16, having a solar cell array (100, 101, 408) adapted for tracking movements of the Sun relative to the spacecraft, wherein the adjustment of the location of the sun ray blocker device (111,112, 141, 271, 301, 411, 511, 611, 811, 921, 951) in relation to the thermal radiator surface (11,12, 1804, 2121, 2721) is synchronised with the tracking movement of the solar cell array, when in normal operation. 18. A spacecraft as claimed in claim 17, wherein the sun ray blocker device (111,112; 141; 27i; 301; 411; 511; 611; 811; 921; 951) is mounted on the solar cell array (100, 101408) or on means carrying said solar cell array. 19. A spacecraft as claimed in claim 18, wherein the solar cell array tracks the movement of the sun by rotation of the solar cell array about an axis of rotation of the solar cell array (100, 101, 408) such that the sun blocker panel (111, 112) also rotates about said axis of rotation of the solar cell array. 20. A spacecraft as claimed in claim 19, wherein the thermal radiator surface (11,12) is orthogonal to the axis of rotation of the solar cell array so that the sun blocker panel (111, 112) rotates about an axis normal to the thermal radiator surface. 21. A spacecraft as claimed in claim 18-20, wherein the adjustment means includes a variable length attachment arm (2130, 2730) for attachment of the sun blocker panel to the spacecraft. 22. A spacecraft as claimed in claim 21, wherein the attachment arm is a scissor arm (2730) . 23. A spacecraft as claimed in claim 21, wherein the attachment arm (2130) is formed of articulated portions (2132, 2134, 2137) which may be mutually articulated during rotation to vary an effective length of the attachment arm. 24. A spacecraft as claimed in any of claims 21-23, wherein the adjustment means for attachment of the sun blocker panel (2100, 2700)to a solar cell array assembly(2131, 2701) are such that a distance between the sun blocker panel from the solar cell array assembly (2130, 2730) may be varied during rotation of the sun blocker panel. 25. A spacecraft as claimed in any of claims 3 to 24, wherein means (929, 931, 955, 957) are provided for adjusting the size of the sun blacker panel (111, 112) . 26. A spacecraft as claimed in any of the preceding claims, including control means for controlling the spacecraft so as to maintain an angle between a sun line and the thermal radiator surface (11,12) below a predetermined angle by adjustment of the orbit and/or attitude of the spacecraft in use. 27. A spacecraft as claimed in claim 26, wherein the predetermined angle is 60 degrees. 28. A spacecraft as claimed in claims 26 or 27, wherein the control means is adapted to maintain the thermal radiator surface (11,12) substantially parallel to a plane of an orbit of the spacecraft. 29. A spacecraft as claimed in any of claims 26 to 28, wherein the control means is adapted to maintain the spacecraft in a sun synchronous orbit. 30. A spacecraft as claimed in any of claims 26 to 28, wherein the control means is adapted to maintain the spacecraft in an equatorial or low-inclination orbit. 31. In a three axis stabilised spacecraft for orbiting about a planet and having at least one solar cell assembly having at least one solar cell panel, and being a north solar cell panel assembly or a south solar cell panel assembly, said at least one solar cell panel assembly being mounted on an axle so as to be controllably rotated from said spacecraft about an axis of rotation so as to face the Sun, said spacecraft having a nadir panel which is generally pointing to the centre of the planet, an opposite panel known as a zenith panel, which faces away from the centre of the planet and sharing the same planar normal vector as said nadir panel, an east panel and a west panel, said east panel and said west panel having their planar normal vector laying on a orbital plane pointing to the velocity vector of the spacecraft generally tangential to the direction of travel on the orbit, and a north panel and a south panel, said north panel and said south panel having their planar normal vector generally perpendicular to the orbit plane and parallel to the axis of rotation of the planet, said solar cell panel extending outwardly from said spacecraft, the improvement which comprises: attaching at least one sun ray blocker device to said at least one solar cell panel, said at least one device being either a north blocker device or being a south blocker device and corresponding to said at least one solar cell panel, each of said at least one sun ray blocker device being positioned forwardly from and offset relative to a solar cell surface of a solar cell panel and at a predetermined angle to either of said north panel and said south panel, said north panel or said south panel, said sun ray blocker device being positioned so as to cast a shadow on at least a majority of the exposed surface of its corresponding north or south panel during solar exposure thereto. 32. In a three axis stabilised low inclination orbit spacecraft for orbiting about the earth and having two sets of solar cell array assemblies having solar cell arrays, one set being a north solar array assembly and the other being a south solar array assembly, said assemblies each being mounted on an axle so as to be controllably rotated from said spacecraft about an axis of rotation so as to face the sun, said spacecraft having an earth panel which is generally pointing to the centre of the earth, an opposite panel known as a zenith panel, which faces away from the centre of the earth and sharing the same planar normal vector as said earth panel, an east panel and a west panel, said east panel and said west panel having their planar normal vector laying on an orbital plane pointing to the velocity vector of the spacecraft generally tangential to the direction of travel on the orbit, and a north panel and a south panel, said north panel and said south panel having their planar normal vector generally perpendicular to the orbit plane and parallel to the axis of rotation of the earth, said solar cell panel extending outwardly from said spacecraft, the improvement which comprises: attaching at least one sun ray blocker device to each of said north solar array and said south solar array, one device being a north device and another device being a south device, each of said sun ray blocker devices being in the form of a panel and being positioned forwardly and offset relative to the solar cell surface of a solar ray and at a predetermined angle to said north panel and said south panel, said north blocker device being positioned so as to cast a shadow on at least a majority of the exposed surface of said north panel during solar exposure thereto, and said south blocker device being positioned so as to cast a shadow on the exposed surface of said south panel during solar exposure thereto.

Inventors: KASKIEWICZ PAUL, LIU LINCHIH OLIVER, WU ALBERT T, KASKIEWICZ, PAUL, LIU, LINCHIH, OLIVER, WU, ALBERT, T.
Applicants: KASKIEWICZ PAUL, LIU LINCHIH OLIVER, TURBOSAT TECHNOLOGY INC, WU ALBERT T, KASKIEWICZ, PAUL, LIU, LINCHIH, OLIVER, TURBOSAT TECHNOLOGY, INC., WU, ALBERT, T.
Application year: 1999
Main classification: B64G1
IPC codes: B64G1