Patent Publication Number: US-2020299005-A1

Title: Satellite control system using electrically controllable variable reflection glass panels

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. patent application Ser. No. 15/074,786, entitled “SATELLITE CONTROL SYSTEM USING ELECTRICALLY CONTROLLABLE VARIABLE REFLECTION GLASS PANELS,” by Harold A. Rosen, PhD, filed Mar. 18, 2016 and issued May 12, 2020 as U.S. Pat. No. 10,647,450, which application is hereby incorporated by reference herein 
    
    
     BACKGROUND 
     1. Field 
     The present disclosure relates generally to spacecraft attitude and orbital control systems, and specifically a spacecraft attitude control system for use in earth orbiting satellites and spacecraft orbit and attitude control of spacecraft of a novel very long baseline gravitational wave detection satellite configuration located near the sun-earth Lagrangian points L3, L4 and L5, utilizing the control of solar radiation pressure by the use of electrically controllable variable reflection glass panels to provide the required torques and forces. 
     2. Description of the Related Art 
     Earth orbiting satellites are known and have been utilized for a large number of purposes, such as communications, navigation, weather observation, and research. The orbital elements of these satellites vary greatly, depending on the purpose of the satellite. Satellites in general are not restricted to earth orbits as they may orbit the sun or other planets. Regardless of mission, most satellites undergo several distinct phases requiring substantial forces and torques: launch, injection into an intermediate orbit, and injection into the operational orbit. The control forces and torques needed for these phases are significant. However, once the satellite is placed into its operational orbit, the torques required for attitude control are more easily manageable, and, in the case of the novel gravitational wave detection satellite configuration, the forces for orbital control are also more easily manageable. 
     Typically, the attitude of satellites is controlled by actuators that comprise chemical or electrical thrusters. Such satellites may include electrically powered reaction wheels (also known as momentum wheels) to aid in attitude control. The problem with using such thrusters can be the transient attitude disturbances they create and the resulting disturbance of the pointing of narrow communication beams or even narrower beams of optical payloads. Additionally, the thruster fuel usage may adversely affect the launch mass of the satellite if the satellite is intended for long duration operation. Adding to these problems, momentum wheels used for attitude control require periodic thruster actuation to keep their angular momentum within prescribed limits. 
     The orbital control of satellites has also used chemical and electrical thrusters, but require greater use of fuel and/or power over the satellite lifetime, and thus are often the limiting factor in satellite life. 
     What is needed is a system and method for controlling the attitude of satellites that minimizes transient attitude disturbances and permits more precise control. In orbits where satellite orbital control needs are minimal, what is needed is a system and method that controls forces in such a way that essentially allows for unlimited satellite lifetime, and also does not contaminate any sensitive instruments onboard the satellite. 
     SUMMARY 
     To address the requirements described above, this document discloses the use of panels of variable reflectivity to control the attitude and orbit of spacecraft. This is accomplished by controllably varying the reflectivity according to an error signal obtained from sensor measurements. 
     The electrically controllable variable reflection glass panel comprises two thin outer layers of highly transparent glass; a middle layer, of liquid crystal droplets embedded in a polymer, whose transparency is electrically controlled by transparent electrodes; and a reflective film on the rear surface of the outer layer that is farther from the sun. This enables an electrically controllable variable reflected light. The magnitude of the reflected light is controlled by impressing a low, variable voltage on the external electrode contacts, creating an electric field in the polymer layer that causes the layer to go from a least to a most transparent arrangement in response to that field. This results in a variable solar radiation pressure on the panel that is used for controlling the satellite. 
     Several embodiments are disclosed, including a first embodiment in which the control panels are affixed to the solar panels of a conventional geostationary satellite, and a second embodiment in which the control panels are affixed to a cylindrical body of a scientific geostationary satellite. In these embodiments, the control panels are used solely for satellite attitude control. In a third embodiment, in which the control panels are affixed to the body of a satellite and the satellite is placed near a Lagrangian point, the control panels are used for both satellite attitude and orbit control. 
     The features, functions, and advantages that have been discussed can be achieved independently in various embodiments or may be combined in yet other embodiments of which further details can be seen with reference to the following description and drawings. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Referring now to the drawings in which like reference numbers represent corresponding parts throughout: 
         FIG. 1  is a view of an electrically controllable variable reflection glass panel; 
         FIGS. 2A, 2B, 2C and 2D  are diagrams depicting vectors that show how the variation of solar radiation pressure affects the magnitude of the control force exerted on the electrically controllable variable reflection glass panel; 
         FIG. 3  is a diagram presenting a typical laboratory measurement of the output voltage, representing the light reflected by the control panel, versus the voltage applied to its external electrode contacts; 
         FIG. 4  is a diagram illustrating the arrangement of the control panels on a conventional geostationary communication satellite; 
         FIG. 5  is a diagram illustrating the arrangement of the control panels on a geostationary cloud-monitoring satellite; 
         FIGS. 6A-6C  are diagrams showing multiple views of the arrangement of the eight control panels on a gravitational wave detection satellite located near a Lagrangian point; 
         FIG. 7  is a diagram illustrating the relevant Lagrangian points (L3, L4 and L5) in the sun-earth system and the nearby locations and orientations of the three gravitational wave detection satellites; and 
         FIG. 8  is a diagram of a six degree of freedom controller that can be used with the eight control panel embodiment. 
     
    
    
     DESCRIPTION 
     In the following description, reference is made to the accompanying drawings which form a part hereof, and which is shown, by way of illustration, several embodiments. It is understood that other embodiments may be utilized and structural changes may be made without departing from the scope of this disclosure. 
     Overview 
     Disclosed is a control system for a satellite that modulates the solar radiation pressure with electrically controllable variable reflection glass panels to generate forces and torques needed for satellite control. The property controlled by the so-called smart glass is its light transmission in response to low power electrical input. Smart glasses having these properties have been used to electrically control the light level in homes, offices, and airplanes. As opposed to these systems, in the system herein used for satellites, a reflective film is added to the assembly and no external transmission is possible. In addition, measurements of the light reflection versus voltage are made in order that the system can be electronically linearized, making it possible to control the system without cross-coupling among the orbit control functions and attitude control functions. 
     To generate the torques needed for attitude control and the forces needed for orbit control, this invention makes use of a reflective film on the glass surface away from the sun of the electrically controllable variable reflection glass panel, causing the radiation pressure vector to vary with the reflectivity. In addition, the geometric arrangement, size, and orientation of the electrically controllable variable reflection glass panels are designed to yield the desired forces and torques for the particular mission. 
     Three exemplary applications are described in this disclosure. They include a geostationary communications satellite application, a cloud-monitoring earth satellite application, and an application directed to a very long baseline constellation of gravitational wave detection satellites located near the L3, L4, and L5 earth-sun Lagrangian points. In the first two applications, attitude control is improved, but gravitational and other forces inherent in those orbits render the electrically controllable variable reflection glass panels impractical for orbital control and maintenance. However, in the third application, the electrically controllable variable reflection glass panels provide not only precision attitude control but also orbit control as well. Orbit control without the use of thrusters in this application is practical because the gravitational perturbations in the orbit used are much less than for earth-orbiting satellites. For example, in the orbits described herein inclination control, the largest component of orbit control for earth-orbiting satellites, is not needed. Likewise, East-west drift and eccentricity corrections also are not needed because of the stability inherent in the Lagrangian points. As described below, to compensate for the solar radiation pressure in the radial direction, the three satellites are located at points slightly closer to the sun than the Lagrangian points. (At the Lagrangian points, there are no forces, as the gravitational and centrifugal forces cancel each other. The satellites depart from it in the radial direction by an amount that provides a force that is equal to and opposite in direction to the solar radiation pressure.) 
       FIG. 1  is a view of an electrically controllable variable reflection glass panel  100  that comprises two thin outer layers of highly transparent glass  102 A,  102 B (alternatively referred to hereinafter as outer layers  102 ); a middle layer  104  comprising liquid crystal droplets embedded in a polymer, whose transparency is electrically controlled by transparent electrodes  106 ; and a reflective film  108  on the rear surface of the outer layer  102 B that is farther from the sun. This enables an electrically controllable variable reflected light. 
     The magnitude of the reflected light is controlled by impressing a low, variable voltage on external electrode contacts  110  which are electrically coupled to the respective transparent electrodes  106 , creating an electric field in the polymer layer  104  that causes the rod-like liquid crystal droplets within the layer  104  to go from a least to a most transparent arrangement in response to that field. This results in a variable solar radiation pressure on the panel  100  that is used for controlling the satellite. 
       FIGS. 2A, 2B, 2C and 2D  are diagrams depicting vectors that show how the variation of this solar radiation pressure affects the magnitude and direction of the force exerted on the electrically controllable variable reflection glass panel. The incident force vector, F i , is fixed for a given panel, independent of the light reflected from the electrically controllable variable reflection glass panel. The control force, F r , which varies with the light reflection, is always opposite to the direction of the reflected light. 
     Turning first to  FIGS. 2A and 2B , the solar radiation pressure results from the change in momentum of the impinging photons: when they are absorbed and come to a halt their velocity change is equal to c, the velocity of light. When incident radiation  202  reflected by a the panel reflective surface  108  normal to the direction of the incoming light (e.g. parallel to surface normal  204 ), the photons undergo a change of  2   c , since the reflected photons travel toward the sun at the speed of light. At the distance from the sun to the earth, one astronomical unit (AU) or 1.5×10 8  kilometers, the average solar power flux density P is 1366 watts per square meter. P varies inversely with the square of the distance from the sun. The pressure p is this flux divided by the speed of light c, or P/c. At one AU, the pressure p is 4.5×10 −6  Newtons per square meter. The corresponding force F i  has a magnitude of Ap Newtons, where A is the area of the surface panel reflective surface  108  projected perpendicular to the sunline  206  in square meters. If (as illustrated in  FIG. 2B ) the surface  108  is normal to the sunline  206  and reflects all the incoming light, p is twice as much, or 9.0×10 −6  Newtons per square meter. The diagrams in  FIGS. 2A and 2B  illustrate these cases, with  FIG. 2A  depicting the case where the panel  100  is set for zero reflection and  FIG. 2B  depicting the case where the panel  100  is set for full reflection. 
     While twice as much radiation pressure is obtained in these cases where the incident radiation is normal to the panel front surface, the amount that can be controlled is the same as when no reflection occurs since the change in momentum due to the incident photons is a constant. However, because the control panels will likely be connected to a solar panel substrate to meet the required panel stiffness, the reflective layer is still useful in that it forms a barrier that isolates the panel from its mounting surface. 
       FIGS. 2C and 2D  illustrate that a major benefit of having a reflective layer (e.g. layer  102 B with the reflective film  108 ) is when the reflective surface  108  is not normal to the direction to the sun (e.g. the sunline  206 ). This results in a control force component that can be perpendicular to the sunline  206 . The magnitude of this force is affected by the amount of light reflected by the control panel  100 . This is shown by the vector diagrams for a surface  108  whose normal is rotated by a thirty-degree angle from the sun direction. 
     In  FIG. 2C , it is assumed that all the impinging light (e.g. incident radiation  202 ) is absorbed by panel  100 , whereas in  FIG. 2D , it is assumed that all the impinging light is reflected by the panel  100 . The angle θ of the reflected light  208  from the surface normal  204  is equal to the angle θ of the surface normal  204  to the sunline  206 . In the case shown in the vector diagram presented in  FIG. 2D , the value of θ is thirty degrees, and the reflected light  208  is N or sixty degrees from the sunline  206 . 
       FIG. 3  is a diagram presenting a typical relationship between light reflected by the control panel, represented by the output voltage, versus voltage applied to the contacts  100 . Because of inherent system non-linearities, the light reflected from the panel  100  will not typically be a linear function of the voltage applied to the contacts  110 . As illustrated, the normalized light reflected from the control panel typically varies non-linearly with the voltage applied to the contacts  110 . Since a linear control system is desirable, this non-linearity may be compensated for. In one embodiment, this may be performed via electronic circuitry in the satellite, for example by use of a linearizing function implemented in a processor with memory. For example, before the satellite is launched, the actual control panel  100  output may be measured and used to generate a function or lookup table that converts a voltage command to the control panel  100  to a linearized voltage command that is actually provided to the control panel  100 . This function may be stored in a memory of the satellite controller. This permits the desired satellite attitude and orbit controls to be designed with precision and without cross-coupling among the six degrees of freedom involved in attitude and orbit control. 
     The voltage commands to the control panels are biased so that a nominal command in response to an error signal of zero is associated with a midrange of reflectivity of each of the control panels, allowing both positive and negative panel correction forces. 
     First Application—Geostationary Communication Satellite 
       FIG. 4  is a diagram illustrating how the attitude control during the operational phase of a conventional geostationary communication satellite  400  can be improved by use of electrically controllable variable reflection glass panels rather than conventional means such as thrusters and momentum wheels. The geostationary satellite  400  shown has two earth-pointing reflector antennas  402 A and  402 B whose beams, ideally, always point to the specific areas on earth being served. The satellite  400  also includes north and south solar panels ( 404 A and  404 B, respectively) coupled to the satellite body  406  that rotate about their longitudinal axis relative to the satellite 360 degrees each day so that they can face the sun and generate the needed power. The view shown in  FIG. 4  is for the midnight solar panel  404  position. The satellite  400  also includes a satellite controller having a processor, memory and sensor package  410  for performing the orbital and attitude control functions. 
     The two solar panels cannot be exactly alike, and their differences result in a small but persistent torques imposed on the satellite  400 . Modern satellites use reaction wheels to maintain the desired satellite attitude, but also then require periodic use of thrusters to center the wheel speed, resulting in transient attitude disturbances. The transients associated with solar panel  404  imbalance can be avoided by using the north and south electrically controllable variable reflection glass control panels ( 408 A and  408 B, respectively) to continuously produce a torque opposing the disturbing torque. The torque command can be calculated (using the appropriate circuitry) and applied as a plus or minus signal, respectively, to the two panels  408 A and  408 B. In this way, a pure torque is always applied. This technique can also compensate for other sources, such as antennas, of radiation pressure imbalances. Further, although the control panels  408  are illustrated as being equidistant from the center of mass of the satellite  400 , other embodiments are possible wherein the control panels are not equidistant, but are also of different surface area, so that the effective torques applied to the satellite body  406  by each control panel  408  about the x-axis are equal. 
     Second Application: Geostationary Cloud-Monitoring Satellite 
       FIG. 5  is a diagram illustrating the application to a geostationary cloud-monitoring satellite  500  having an optical payload  502  with a very narrow beam, demanding highly accurate and transient-free attitude control. Because the payload  502  requires very little power, the satellite power needs can be provided by solar cells  504  on its cylindrical surface rather than by external solar panels. In this application, the electrically controllable variable reflection glass panels are located on the northern and southern ends of the cylinder of the satellite body ( 504 A and  504 B, respectively) and are cylindrical rather than planar in shape in order to present the same view to the sun as the satellite performs its 24 hour period rotation. Alternatively, this could be provided by a multiplicity of small planar control panels arrayed around the satellite body to approximate the cylindrical surfaces of the north control panel  504 A and the south control panel  504 B. 
     Third Application: Gravitational Wave Detection Satellite 
       FIGS. 6A-6C  are diagrams showing one embodiment in which variable reflection control panels are to be used to control both attitude and orbit in particular satellite orbits. This configuration permits all six degrees of freedom—three translational and three rotational—of the satellite to be controlled solely by use of the variable reflection control panels (without the need for thrusters), and provides a tremendous advantage in a space-based system for the detection and measurement of low frequency gravitational waves from the far reaches of the universe. 
     This requires a very long baseline space antenna. This very long baseline can be provided by the use of three satellites orbiting the sun at the Lagrangian points L3, L4 and L5 of the sun-earth system. 
       FIG. 7  is a diagram illustrating these Lagrangian points. Because the gravitational disturbances are so small at these locations, being due primarily to the attractions of the other planets, orbit control is possible by the use of control panels alone on the satellite  600  and thus no thrusters are needed for this function. In addition, as in the cases of the earth-orbiting satellites, attitude control is possible as well. 
       FIG. 6A  is a diagram illustrating an embodiment of a satellite  600  using eight electrically controllable reflection glass panels  1 - 8  on the satellite body  602 . The control panels  1 - 8  are arranged into four pairs arrayed symmetrically about the sunline  206  and the center of mass of the satellite  600 . As illustrated in  FIG. 6A , control panels  1  and  2  together comprise a first pair of control panels. A second pair of control panels  5  and  6  are located on a side of the center of mass of the satellite  600  diametrically opposing the first pair of control panels  1  and  2 . Similarly, control panels  3  and  4  comprise a third pair of control panels, and control panels  7  and  8  are located on a side of the satellite  600  center of mass diametrically opposing the third pair of control panels. 
     As illustrated in  FIG. 6A , the normals  204  of each of the control panels  1 - 8  are offset in angle from the sunline  206  by an angle θ. In the illustrated embodiment, the offset angle θ is thirty degrees, which results in only a small loss of projected area and provides a good compromise among the force components perpendicular to and parallel to the sunline  206 . Further, the reflected light  208  from each control panel  1 - 8  does not strike the solar panel  604 , any other control panel, or other part of the satellite  600 . 
       FIG. 6B  is a diagram illustrating the same satellite  600  as seen from the point of view of the sun. The satellite always faces in this direction as it traverses its yearlong circular orbit. 
       FIG. 6C  shows a side view of the same satellite  600 . As can be seen in the drawing, the control surfaces are all rotated thirty degrees from the sun normal direction. This arrangement of panels allows control of each of the six degrees of freedom while eliminating cross-coupling. 
       FIG. 8  is a diagram of a six degree of freedom controller that can be used with the eight control panel embodiment illustrated in  FIGS. 6A-6C . Sensor measurements are used to provide error signals that are then used to generate control system forces (F x , F y , and F z ) and torques (T x , T y , and T z ) that drive the error signals to zero. Those desired system forces and torques are used by the linear combiner  802  to determine the panel forces (F 1 -F 8 ) needed from each of the eight control panels. Each of these each panel forces (F 1 -F 8 ) is converted into associated linearized voltages (V 1 -V 8 ) provided to the electrodes  106  of the associated control panels. This closed loop process is performed continuously over time and with a continuous range of reflectivity levels, and hence drives the error signal to zero with minimal transients. 
     It is also noted that although an embodiment of the satellite  600  having eight control surfaces  1 - 8  has been discussed for purposes of illustration, a six degree of freedom control of the orbit and attitude of the satellite  600  may be obtained with fewer (e.g. six) control surfaces. 
     CONCLUSION 
     This concludes the description of the preferred embodiments of the present disclosure. The foregoing description of the preferred embodiment has been presented for the purposes of illustration and description. It is not intended to be exhaustive or to limit the disclosure to the precise form disclosed. Many modifications and variations are possible in light of the above teaching. It is intended that the scope of rights be limited not by this detailed description, but rather by the claims appended hereto.