Patent Publication Number: US-2021179298-A1

Title: System and method for determining an initial orbit of satellites post deployment

Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application claims priority to US Provisional Application No. 62/946,497 filed Dec. 11, 2019, and US Provisional Application No. 62/947,029 filed Dec. 12, 2019, the contents of each of which are herein incorporated. 
    
    
     BACKGROUND OF THE INVENTION 
     The present invention is directed to a system and method for tracking objects launched from a platform, and more particularly, to determine and establish the initial orbit of satellites deployed from a platform travelling through outer space soon after deployment; i.e., soon after the moment of release from the platform, within seconds after deployment; the time it takes to travel the 0.5 to 20 meters to be within the sensor range. 
     Ever since humanity entered the space-age in the late 1950s it has been launching satellites and other objects into space to then be maintained in orbit about the earth. In the early days, and most of today&#39;s launches, single rockets launched a relatively small number of objects, typically one object per rocket. Therefore they have been easily tracked by earth-based radar and optical systems. Satellite operators typically need a few passes of one satellite through the field of view of the earth-based radar and optical systems, taking hours to days, to determine orbital parameters of their satellite. 
     However, in the past decade the advent of small satellites led to the launch of tens to more than a hundred satellites by a single rocket. A notable example is the Indian Space Research Organization (ISRO) Polar Satellite Launch Vehicle (PSLV) flight C37 2017 launch of 104 satellites. The satellite Cartosat-2D, 712 kg and the size of a small car, was C37&#39;s primary payload and the other one hundred and three secondary payload satellites were small satellites, of the nanosatellite class, with sizes as small as a loaf of bread and mass of 4 kg. As a result, the prior art earth-based sensors tasked with observing satellites after release, so that their orbital parameters can be determined, are overwhelmed. Consequently, the uncertainties of the initial orbit parameters and epoch of these orbits can be large and make the initial orbit determination problematic and resource intensive. By orbit parameters it is meant either the three-dimensional position vector and three-dimensional velocity vector or six orbital elements such as semi-major axis, eccentricity, inclination, argument of periapsis, right ascension of the ascending node, and true anomaly This uncertainty can lead to extensive use of ground based sensors to search for the satellites and establish orbital parameters. This is further exacerbated by the plans for a few low earth orbit communications networks which contemplate the launching of hundreds if not thousands of small satellites substantially simultaneously at a rapid cadence; as a result of substantially simultaneous launch from a single vehicle. 
     As seen in  FIGS. 1A, 1B  a launch vehicle  12  has an upper stage, platform,  12   a  carrying a payload P of P 1 -P N  satellites to be launched therefrom. As seen in  FIG. 1A , launch vehicle  12  launches from earth  10  and during ascent to orbit OR it exposes its payload P for launch as known in the art. As seen in  FIG. 1B , each payload establishes a respective orbit O 1 -O N  upon deployment from launch vehicle launch platform  12   a.  It is desirable to determine the orbit parameters of each launched satellite P N  relative to the earth  10  as quickly as possible so that satellite operators have a timely and accurate set of orbit parameters for each satellite and catalogs that contains orbit parameters can be updated rapidly and accurately. 
     The problem becomes that as the number of launch satellites increases, as they establish orbits, earth-based initial orbit determination becomes difficult, inaccurate, and sometimes impossible utilizing the ground based prior art systems. Ground based active sensors such as radars and laser rangefinders get overwhelmed as too many objects appear as “chaff,” and the multiple inputs saturate and confuse the sensors. Additionally, prior art passive optical sensors, such as telescopes, have too narrow a field of view to capture the large number of satellites being launched. They require several passes of each satellite to accurately make observations required to perform initial orbit determination. Both senor types quickly get overwhelmed when dealing with the number of satellites now contemplated to be launched. 
     Accordingly, it is desired to provide an initial orbit determination system which overcomes the shortcomings of the prior art and enables more timely and accurate tracking of multiple payloads immediately after deployment from the platform (rocket). 
     SUMMARY OF THE INVENTION 
     A system for determining an initial orbit of an object launched from an orbiting launch vehicle includes a sensor affixed to the launch vehicle. A command and data handling subsystem that includes a computer and one or more digital and analog interfaces receiving inputs from the sensor. A navigation subsystem, connected to the command and data handling subsystem determines the orbital parameters of the launch vehicle relative to earth, the orientation and angular rates of the launch vehicle with respect to a celestial reference frame, and transmits them to the command and data handling subsystem. A communications subsystem is also connected with the command and data handling subsystem and it is used to transmit and receive messages between the command and data handling subsystem and an earth-based communication system of a ground station. The sensor is an active device for transmitting electromagnetic signals toward the object launched from the launch vehicle, and receiving the signals reflected therefrom by the object that was launched. 
     The command and data handling subsystem processes the reflected signals of the sensor and determines the range, azimuth, and elevation angles of the launched object. The command and data handling subsystem determines the position and velocity vectors of a launched object relative to the platform. The command and data handling subsystem further receives the output of the navigation subsystem and combines the relative position and velocity vectors with the orientation and orbital parameters of the platform and determines the orbital parameters of the launched object with respect to earth. Finally, the orbital parameters, that represent the initial orbit of the launched object relative to earth, are transmitted to a ground station by the communication subsystem of the invention. 
     In one embodiment of the invention the transceiver is a phased array device and more particularly, a radar sensor. In another embodiment the sensor is a flash lidar. The navigation subsystem and command and data handling subsystem may be integrally formed with the launch vehicle. In effect the launch vehicle is a platform 
     In another embodiment of the invention the system includes a discrete platform. The platform is mounted on the launch vehicle. The sensor, navigation determination subsystem, and command and data and subsystem are disposed on the platform. The sensor has a wide field of view and a close range. 
     Because the platform, and hence the sensor affixed to it, can rotate after the object is launched the invention can include as many sensors as needed to obtain a cumulative field of view extending up to a full sphere (4π steradian) with a range out to about 1 kilometer. 
     In yet another embodiment of the invention the launch vehicle rotates relative to the launched payload. When the cumulative sensor field of view is less than a full sphere, the reflected signal is received periodically. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present disclosure will be better understood by reading the written description with reference to the accompanying drawings, in which like reference numerals denote similar structure and refer to like elements throughout in which: 
         FIGS. 1A, 1B  are schematic drawings showing a launch vehicle for launching a multiple satellite payload into orbit as known in the art; 
         FIG. 2  is a schematic diagram of the system for determining initial orbit of a satellite constructed in accordance with one embodiment of the invention; as deployed on a launch vehicle; 
         FIG. 3  is a block diagram of a system for determining the initial orbit of the satellite constructed in accordance with the invention; 
         FIG. 4  is a plot of the field of view of one embodiment of the sensor operating in accordance with the invention, the thin lined portions of the path being the path of the launched object with respect to the sensor that rotates with the launch platform, and the thicker lined portions of the path being the path of the launched object in the field of view of the sensors, thus illustrating timing of the sensed object within the field of view of a sensor constructed in accordance with the invention; 
         FIG. 5  is a graph of the detection of the reflected signal by the sensor, as a function of the velocity and distance of the object detected at each instance of detecting the reflected signal within the field of view; 
         FIG. 6  is a flowchart for tracking the initial orbit of satellites deployed from a platform travelling through outer space in near real-time after deployment ; and 
         FIG. 7  is a schematic drawing of a system for detecting the initial orbit of a satellite constructed in accordance with another embodiment of the invention. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     Reference is now made to  FIGS. 2 and 3  in which a system  100  for determining initial orbit, constructed in accordance with the invention, disposed on a launch vehicle  12   a  is shown in detail. System  100  includes a platform  102 . Platform  102  may be a substrate mounted to launch vehicle  12   a,  or may be the structure of launch vehicle  12   a  itself; in other words, system  100  may be integrated into launch vehicle  12   a.    
     System  100  includes a command and data handling subsystem  104  mounted on platform  102 . 
     The command and data handling subsystem  104  receives and processes information from a sensor  106  and a navigation subsystem  108 , each described below, and provides an output to a telecommunications subsystem  140  for reporting results back to earth  10 . 
     As discussed above as known in the art, launch vehicle  12   a  is provided with one or more satellite deployers  204   1 - 204   N . To simplify matters, for ease of explanation, it is assumed in this description that deployers  204  launch payloads P in a direction substantially orthogonal outward from the surface of launch vehicle  12   a  upon which the deployers  204  are disposed. Payloads P are launched with a known velocity in the substantially X P  direction. 
     Each respective sensor  106  is mounted on launch vehicle  12   a  with an orientation facing away from launch vehicle  12   a  to facilitate monitoring payloads P 1 -P N  substantially simultaneously as launched. In other words, as a result of field of view size and orientation selection, positioning of sensors  106  relative to deployers  204 , deployed satellites enter the field of view of a given sensor  104  substantially immediately upon deployment. Operatively, sensors  106  are active sensors positioned near the deployers  204  to assess the relative orbital path of payloads P with respect to sensor  106 . Each sensor  106  emits a signal which is reflected back from each respective payload P within the field of view of the respective sensor  106  to be received by a respective sensor  106  as a reflected signal. Sensors  106  may determine range (distance) and one angle (azimuth) or both angles (azimuth and elevation). As a result, the reflected signal is indicative of position and velocity of the payload P relative to sensor  106 . Preferably sensors  106  are oriented so that the signal is emitted from sensor  106  in a direction substantially parallel with the direction of payload launch; in the X P  direction. This maximizes the period of time within which a specific payload P N  is within the field of view of a respective sensor  106  and the orientation direction of the sensor can be determined prior to launch through simulations. 
     System  100  is primarily concerned with determination and tracking of the initial orbit. Therefore, the field of view of sensor  106  is preferably wide, along an axis Y S , but not necessarily deep along an axis X S  as shown in  FIG. 4 . In preferred non limiting embodiments, the field of view of the sensor is within a range of 6°-160° and preferably 90°. 
     The range of the sensors is preferably between 20 m to 1000 m from sensor  106 , but some contemplated radars and lidars have a range of ranges between 0.05 m up to 200 m. Additionally, launch vehicle  12   a  rotates about its center of mass  202  during the deployment so that as launch vehicle  12   a  travels along its orbital path O R  during a deployment procedure, a specific payload will appear to travel across the field of view of a single sensor  106  as result of the motion of sensor  106  relative to payload P as payload P finds its orbital path as launch vehicle  12   a  rotates. Therefore, it is desirable to have at least a second sensor  106   b  for tracking payloads P. As a specific payload P N  leaves a field of view of a first sensor  106   a  it will come into view of a second sensor  106   b . Most preferably the arrangement of sensors includes as many sensors as needed to obtain a cumulative field of view extending up to a full sphere (4 π steradian) with a range out to about 1 kilometer. As a result, there is a longer tracking time and greater tracking field of view and increased length of the reflected signal; increasing accuracy in determining the current position of the satellite, and the initial orbit of any particular payload P. 
     Sensors  106  emit signals in either the radio or optical frequency range, including visible and near infrared spectra. In a preferred nonlimiting embodiment sensor  106  is a phased array transceiver capable of emitting signals to an object and receiving signals reflected therefrom which are utilized to determine distance and relative position; velocity, azimuth, and elevation. In the preferred non limiting embodiment sensor  106  is a flash lidar sensor, but a radar sensor may be used as well. The received reflected signal is input to the command and data handling subsystem  104  where the distance and velocity of the sensed payload P, relative to sensor  106 , and in turn to system  100 , is determined as a function of the reflected signal. 
     However, determining the position of a particular payload during initial orbit relative to launch vehicle  12   a  is not helpful to determining the orbit relative to earth  10  so that others will know the positioning of the payload P relative to earth and other objects orbiting earth  10 . Therefore command and data handling subsystem  104  also determines the position of the center of mass  202  of launch vehicle  12   a  relative to the frame of reference with the origin at the center of mass  304  of earth  10 . To accomplish this, system  100  also includes a navigation subsystem  108  (also “navigator”) for providing orbit parameters information of launch vehicle  12   a  relative to earth  10  and orientation of launch vehicle  12   a  with respect to the celestial sphere to the command and data handling subsystem  104 . 
     Navigator  108  includes a plurality of navigation sensors for determining the orbital parameters of launch vehicle  12   a,  and in turn of system  100 , relative to earth  10  and its orientation with respect to the celestial sphere. Each of the navigation sensors have a specific role to determine the i) orbital parameters of launch vehicle  12   a,  and in turn of system  100 , relative to earth; ii) the orientation of launch vehicle  12   a,  and in turn of system  100 , relative to earth and iii) the angular speed of launch vehicle  12   a,  and in turn of system  100 , with respect to an inertial (celestial) frame centered at the earth. Each of the navigation sensors have a specific role to determine the i) orbital parameters of system  100  relative to earth; ii) the orientation and iii) the angular speed with respect to an inertial (celestial) frame centered at the Earth 
     A first sensor is one or more sun sensors  110  for determining the orientation of system  100  relative to the sun. A second sensor is a three-axis magnetometer  113  for determining the orientation and strength of the earth magnetic field of the earth at the sensor  113 . The sun sensor and magnetometer measurements are used to determine the orientation of the launch vehicle with respect to an inertial reference frame with origin at the center of mass of the earth. A third sensor is an inertial measurement unit  112 , which much like a gyroscope on a maritime ship, determines the angular rate of the launch vehicle  12   a  relative to a celestial reference frame with origin at the center of mass of earth. A fourth type of sensor is the Global Positioning System (GPS) receiver  114  which receives signals from the GPS satellite network orbiting earth  10  through the GPS antenna  118  to determine the position of launch vehicle  12   a,  and in turn sensor  106 , relative to the earth. Lastly, a star tracker  116  may be used which determines the orientation of system  100  relative to known constellations. 
     It should be noted, that one or more of each of these types of sensors, or none of these types of sensors may be used. It is possible to utilize only a single such sensor, but to increase accuracy, so that in a preferred non limiting embodiment, the above enumerated sensors may be used in combination and in a preferred nonlimiting embodiment; at least one of each sensor is used in combination with all three of the other sensors in orbit determiner  108 . Other such orbital and orientation determination sensors may be used in place of any of the above as is known in the art. 
     Command and data handling subsystem  104  receives the output of navigator 108  through digital input/output module  120  and, utilizing an on board computer  119 , determines the orbit parameters of system  100  relative to a center of mass  300  of the earth  10  (“earth frame”) and the orientation of the launch vehicle with respect to an inertial celestial frame with the origin at the center of mass of earth. Utilizing frame transformation processes, command and data handling subsystem  104  transforms the relative position and velocity vectors of the payload P relative to the launch vehicle  12   a  as determined by sensor  106 , to the earth frame. The result is output to a ground station utilizing telecommunication subsystem  140  having a transceiver  144  and an antenna 142 . In a preferred non limiting embodiment system  100  broadcasts over the S-band. In another non limiting embodiment system  100  broadcasts results to a ground station or to payloads P themselves through a satellite communications system such as Globalstar or Iridium. 
     In a preferred nonlimiting embodiment, system  100  includes an electrical power subsystem  130 . System  100  may be powered by onboard batteries  132  and/or solar panels  134 . A power management and distribution subsystem  136  controls the output of energy from either batteries  132 , solar panels  134  or both, to sensor  106 , sensor  108  and command and data handling subsystem  104  in response to control signals from command and data handling subsystem  104 . In this way, batteries  132  can be conserved as a function of the availability of solar power, and there is a backup power supply to prevent disruption of this functionality. 
     The operation of the electronic components is affected by temperature. As a result, system  100  includes a thermal control subsystem  122  having temperature sensors  124   1 - 124   N  monitoring temperatures at various positions along system  100  and provide an input through analog-to-digital converter  122  commanding data handling subsystem  104 . In a preferred nonlimiting embodiment the thermal control subsystem  122  operates passively and includes insulation  126  and one or more heat conduction components  128  to radiate heat away from the system components that require it. In yet another nonlimiting embodiment the thermal control subsystem includes active thermal control components such as heaters and coolers that are controlled either thermostatically, by a bimetal switch for example, or by the command and data handling subsystem  104 . 
     Reference is now made to  FIGS. 4 and 5  in which a graphical representation of the operation of system  100  is provided. As discussed above, launch vehicle  12   a  can rotate as the payloads P are deployed. Sensors  106  which are fixed to the body of launch vehicle  12   a  rotate with launch vehicle  12   a.  Therefore as discussed above, payloads P may only appear within the field of view of a respective sensor  106  periodically. As shown in one extreme example in  FIG. 4  sensor  106  is attached to launch vehicle  12   a  at a location away from the center of mass of the launch vehicle. Sensor  106  in this example has a field of view that extends 6°, full angle, in elevation and 36° in azimuth, full angle, relative to the boresight axis of the sensor. Because sensor  106  is rigidly attached to the launch vehicle  12   a  that rotates about its axis and the ejection force and environmental forces, such as drag, separate payload P N  from the launch vehicle, in this embodiment, a path indicated by the growth spiral extending from local ejection point (at time t 0 ) of launch vehicle  12   a  is the relative path of motion of the payload P N , in the reference frame of sensor  106 , as it reaches its own particular orbit O N . In the embodiment shown, the relative path is that shown after 60 seconds from separation (at time t f ). The relative positions in which the payload P N  is captured within the field of view of sensor  106  are shown by the relatively thickened portions of the line C 1 -C N , and as expected increases as the payload P N  moves farther away from the initial ejection position. 
     Given the reflected signal received by the system  100  during each instance when the payload P N  is within the field of view of sensor  106 , shown in  FIG. 4 , command and data handling subsystem  104  can operate on this information. In the nonlimiting embodiment shown in  FIG. 4  sensor  106  determines the range (distance) and azimuth angle of payload P N  when the payload P N  is in the field of sensor  106  with a certain cadence. The set of range and azimuth angle pairs is used in an Unscented Kalman Filter (UKF) algorithm to determine the relative position and velocity vectors of payload P N  with respect sensor  106  reference frame. The command and data handling subsystem uses the (known before launch) position of sensor  106  in the reference frame of the launch vehicle  12   a,  the angular rate of the launch vehicle with respect to the celestial sphere determined with the gyroscope of the navigator, and the orbit of the launch vehicle with respect to earth, determined with the GPS receiver of the navigator, together with the relative position and velocity vector of the payload P N  to calculate the orbit of payload P N  with respect to earth. 
     In one nonlimiting embodiment sensor  106  measures range (distance), azimuth, and elevation. In yet another nonlimiting embodiment sensor  106  only measures range. 
     Reference is now made to  FIG. 6  in which the method of operation of initial orbit determination system  100  is shown. In a step  200 , sensor  106  determines information about a sensed payload P N . Sensor  106  is continuously outputting sensor data in step  200 . Sensor  106  outputs data is indicative of either i) range (distance); ii) range or azimuth; and iii) range, azimuth and elevation of P N  relative to the sensor and in turn the platform. 
     At the same time, navigation subsystem  108  is continuously receiving, from a plurality of sensors, data that is used in the initial orbit determination of the payload P N  that is in the field of view of the sensor. In a step  202  navigation subsystem  108  utilizes magnetometer measurements input from magnetometer  113  to determine the orientation of the platform  100  with respect to an inertial reference frame with origin at the center of mass of the earth  10 . In a step  204  the attitude (azimuth , elevation) of launch platform  100  is determined with respect to the celestial sphere by utilizing star tracker  116  or sun sensor  110 . In a step  206  star tracker  116  determines components of the orientation of the quaternion with respect to an inertial reference frame . Simultaneously, in a step  208  navigation subsystem  108 , utilizing inertial measurement unit  112  determines the (x,y,z) components of angular rate of launch vehicle  12   a  measured at the platform  100  with respect to an inertial reference frame. utilizing the output of the onboard gyroscope of inertial measurement unit  112 . Additionally, platform position and velocity with respect to earth  10  are determined in a step  210  either by GPS receiver  114  or by command and data handling subsystem  104  utilizing other inputs. 
     In a step  212 , command and data handling subsystem  104  receives the outputs of magnetometer  113  and sun sensor  110  and determines the components of the orientation quaternion with respect to an inertial reference frame. Simultaneously, in a step  214  command and data handling subsystem  104  estimates the position and velocity of payload P N  relative to at least one sensor  106 . 
     In a step  216 , command and data handling subsystem  104  receives the output of star tracker  116  and inertial measurement unit,  112  determined in step  208 , utilizes the determined components of position vector and velocity vector of the payload P N  and the determined components of the orientation quaternion as determined in step  212  and  214  and transforms the (x,y,z) components of the position vector and velocity vector of payload P N  from the sensor  106  reference frame to an inertial reference frame  2   zz.  In a step  218 , command and data handling subsystem  104  utilizes this transformed inertial reference frame to transform the (x,y,z) components of the position vector and velocity vector of payload P N  from the inertial reference frame  2   zz  to an earth reference frame  2   ww.    
     Command and data handling subsystem  104 , in response to the determined transformed inertial reference frame from step  218 , determines the position and velocity of payload P N  relative to earth  10  in a step  220 . Then in a step  230 , the position of payload P N  and velocity relative to earth is transmitted to earth utilizing telecommunication subsystem  140 . 
     Once system  100  determines an initial orbit of payload P N  and communication subsystem  140  establishes a link the initial orbit of payload P N  is transmitted. In a nonlimiting embodiment communication subsystem  140  transmits the initial orbit data to a ground station directly. In another nonlimiting embodiment communication subsystem  140  transmits the initial orbit to the ground station through a satellite communication system such as Globalstar or Iridium. 
     Reference is now made to  FIG. 7  in which system  100  is deployed on a launch vehicle  12   b;  like numerals are used for like structure for ease of explanation, the primary difference being the orientation of sensor  106  relative to deployers  204 . The field of view of sensor  106  is substantially orthogonal to the direction of deployment of payloads P. In this situation sensor  106  is provided with wide field of view to capture payloads as they leave launch vehicle  12   b.    
     With the above invention, determination of the initial orbit of payloads, space objects, is achievable soon, tens of seconds to minutes, after their deployment from a launch vehicle is achievable. Furthermore, while the above example is provided in connection with initial orbit determination of satellites launched from a launch vehicle, the system can also determine the density of atmosphere, between the launch vehicle and the payload space objects, after deployment. Furthermore, as can be seen above, it can determine both the motion of spacecraft fragments (debris) that result either from impact with an external object or from a spacecraft-internal event that generates debris; including the determination of the direction, size, and speed of the impacting object. Because of this the method and system are easily adaptable to determine the possibility of collision with an object upon which the system resides with another space object. 
     In another nonlimiting embodiment, the active sensor uses the transmit signal to broadcast the initial orbit data to the payload P N  that is equipped with a receiver and command and data handling subsystem capable of receiving and interpreting the data. 
     In other nonlimiting embodiments sensor  106  can have its own microcontroller. The user can set various parameters such as measurement cadence, intensity of the emitted laser beam, etc. through the command and data handling subsystem  104 . The user can also read housekeeping data such as voltages and temperatures that can be transmitted to earth and used for improvements of the design. 
     Additionally, components of the navigation subsystem  108  such as the star trackers, GPS receiver, and Inertial Measurement Unit (IMU) may have their own microcontrollers as well that interface with the command and data handling subsystem  104  with a two-way interface. The user can set update rates, and read housekeeping data such as voltages and temperatures. 
     Because sensor  106  is near the payloads P (on board within meters or less, not earthbound) sensor  106  can be small and use little electric power. Sensor  106  is not overwhelmed by the multitudes of the payloads deployed because only a few payloads P will be in its field of view at the same time. Again, this is due to the proximity to the payloads P of sensor  106 . To gather all the data needed for initial orbit determination the system  100  has components commonly used in satellites. However, in the inventive system they are configured to perform initial orbit determination instead of the functions of a satellite. 
     While this invention has been particularly shown and described to reference the preferred embodiments thereof, it would be understood by those skilled in the art that various changes in form and detail may be made therein without departing from the scope of the invention encompassed by the appended claims.