Abstract:
A system and method for deploying a satellite having an integrated bus and an aperture, wherein the integrated bus extends along its longitudinal axis and wherein the aperture intersects the longitudinal axis. The satellite is deployed into an orbit, wherein deploying includes orienting the satellite so that the aperture points at a desired location. The satellite is spun so that the logitudinal axis and the aperture remain pointing at the desired location and data regarding the desired location is captured via the aperture.

Description:
BACKGROUND 
       [0001]    Micro-satellites or small-satellites have the capability to perform a variety of missions to meet reconnaissance and surveillance needs. They require, however, dedicated launch vehicles to meet the need on demand. A typical approach for launching the micro-satellites has been to rideshare with a larger primary payload. Alternatively, airborne launch vehicles have been proposed for launch on demand. 
         [0002]    What is needed is an integrated micro-satellite system capable of being deployed via rideshare or airborne launch vehicle while still delivering high quality surveillance directly to the user. 
     
    
     
       BRIEF DESCRIPTION OF THE FIGURES 
         [0003]    In the drawings, which are not necessarily drawn to scale, like numerals may describe similar components in different views. Like numerals having different letter suffixes may represent different instances of similar components. The drawings illustrate generally, by way of example, but not by way of limitation, various embodiments discussed in the present document. 
           [0004]      FIG. 1  illustrates one example embodiment of a micro-satellite; 
           [0005]      FIG. 2  illustrates another example embodiment of a micro-satellite; 
           [0006]      FIG. 3  illustrates one example embodiment of an air launch vehicle carrying the micro-satellite of  FIG. 2 ; 
           [0007]      FIG. 4  illustrates another example embodiment of a micro-satellite; 
           [0008]      FIG. 5  illustrates a method of deploying the micro-satellite of  FIGS. 1 ,  2  and  4 . 
       
    
    
     DETAILED DESCRIPTION 
       [0009]    In the following detailed description of example embodiments of the invention, reference is made to specific examples by way of drawings and illustrations. These examples are described in sufficient detail to enable those skilled in the art to practice the invention, and serve to illustrate how the invention may be applied to various purposes or embodiments. Other embodiments of the invention exist and are within the scope of the invention, and logical, mechanical, electrical, and other changes may be made without departing from the subject or scope of the present invention. Features or limitations of various embodiments of the invention described herein, however essential to the example embodiments in which they are incorporated, do not limit the invention as a whole, and any reference to the invention, its elements, operation, and application do not limit the invention as a whole but serve only to define these example embodiments. The following detailed description does not, therefore, limit the scope of the invention, which is defined only by the appended claims. 
         [0010]    As noted above, micro-satellites are used for reconnaissance and surveillance. It can be necessary to launch such satellites with little or no warning. 
         [0011]    The Airborne Launch Assist Space Access (ALASA) program launched in 2012 by the Defense Advanced Research Projects Agency (DARPA) is aimed at getting satellites in the air quickly, cheaply, and from anywhere rather than from a limited number of launch sites. It is anticipated that airplane-based launch systems could get satellites into space on a 24-hour turnaround. 
         [0012]    A dedicated launch of small satellites on airborne launch vehicles creates a paradigm shift in space utilization. Airborne launching of satellites has several advantages, including the ability to launch within hours of call-up, an increase in orbit accessibility (via the availability of multiple launch sites), and the ability to adapt the launch parameters to achieve the best orbit utilization for the given mission. However, in order to effectively utilize the airborne launch vehicle the micro-satellite must meet size, weight and power constraints. These constraints can be met via an integrated payload and bus configuration that fits within the launch vehicle while also meeting mission performance requirements at the lowest possible cost. 
         [0013]    An integrated micro-satellite design can also be launched effectively in a rideshare mode on currently available launch vehicles. In one example embodiment, the micro-satellites are scalable; the satellite selected for launch is a function of the quality of surveillance imagery desired or expected. 
         [0014]    Some such micro-satellite systems are shown in  FIGS. 1-5 . Most air launch vehicles have an ogive fairing that restricts accommodation of the desired payload.  FIGS. 1-5  illustrate scalable means of integrating the payload and the bus while effectively utilizing the shape and volume of the launch vehicle and at the same time maximizing the overall mission imaging capability. 
         [0015]    One embodiment of a micro-satellite system capable of meeting the aforementioned requirements is shown in  FIG. 1 . In the embodiment shown in  FIG. 1 , micro-satellite system  100  includes a parabolic aperture  102 , a focal plane array  104 , and a bus  106 , all located along a longitudinal axis  101 . In one example embodiment, aperture  102  is a thermally stable composite dish separated from bus  106  via multifunctional struts  120 . In the example embodiment shown, focal plane array  104  is an EO/IR focal plane array. 
         [0016]    In one example embodiment, incoming radiation arrives approximately parallel to the longitudinal axis. Parabolic aperture  102  receives the incoming radiation and focuses it on focal plane array  104 . 
         [0017]    In the example shown, bus  106  is located in front of aperture  102 . In one example embodiment, bus  106  includes an attitude control subsystem (ACS)  108 , a propulsion subsystem  110 , a command and data handling subsystem  112 , a data processing and storage subsystem  114  and a power subsystem  116 . 
         [0018]    In the example embodiment shown in  FIG. 1 , a parabolic communication antenna  115  is mounted at the front end of bus  106 . In one such embodiment, the communication antenna is sized to fit at the end of the bus. In one such embodiment, the cylindrical configuration of bus  106  and the placement of parabolic aperture  102  provide rotational symmetry so that the spacecraft  100  can be stabilized by spinning. In one such embodiment, the spinning of satellite  100  also tends to foster a uniform temperature on the parabolic aperture for reduced thermal distortion. In another example embodiment, bus  106  has a shape that approximates a slender rectangular prism, having a small cross-section along the longitudinal axis. 
         [0019]    In one embodiment, bus  106  is miniaturized to impart minimum obstruction to radiation collection. As can be seen in  FIG. 1 , in one example embodiment bus  106  is approximately 0.1 meters in diameter 
         [0020]    In one such embodiment, the cylindrical structure of the bus  106  as well as the back side of the parabolic aperture  102  allow direct mounting of solar arrays ( 122  and  124 ) such that a portion of the solar array is always pointed towards the Sun. In the example embodiment shown in  FIG. 1 , a GPS receiver  118  is mounted over solar array  124  at the back of aperture  102 . 
         [0021]    In one example embodiment, communications antenna  115  .of  FIG. 1  has a diameter of approximately 10 cm, aperture  102  has a diameter of approximately 90 cm and the length of micro-satellite  100  is approximately 2 m, Micro-satellite system  100  can be scaled up or down as necessary to meet the mission parameters. 
         [0022]    Another embodiment of a micro-satellite system  100  capable of air launch is shown in  FIG. 2 . In the Cassegrainian configuration shown in  FIG. 2 , the Focal Plane Array  104  is mounted on the same side as the parabolic aperture  102  (also refereed to as “primary reflector”). This results in much longer focal length, which in turn results in better ground separation distance (GSD) for imaging. In the example embodiment shown in  FIG. 2 , bus  106  is mounted on the longitudinal axis behind aperture  102  while the back side of the parabolic communication antenna  115  includes an optical reflector  126  (also known as the “secondary reflector”). As seen in  FIG. 3 , the example embodiment of satellite system  100  shown in  FIG. 2  can be sized to fit inside the ogive fairing of a typical airborne launch vehicle  200 . In the example embodiment shown in  FIG. 2 , bus  106  is approximately 0.1 meters in diameter. 
         [0023]    In one example embodiment, communications antenna  115  of  FIG. 2  has a diameter of approximately 10 cm, aperture  102  has a diameter of approximately 90 cm and the length of micro-satellite  100  is approximately 2 m, Micro-satellite system  100  can be scaled up or down as necessary to meet the mission parameters. In one such example embodiment, launch vehicle  200  is 264 inches in length and 36 inches in diameter along its main body but flares to 56 inches in diameter before reaching the fins. 
         [0024]    In yet another embodiment of micro-satellite system  100 , as shown in  FIG. 4 , the propulsion and the attitude control subsystems ( 108 ,  110 ) are separated from the main bus structure by means of a deployable boom  128  (shown extended in  FIG. 4 ). Lower system mass is achieved in this approach by making use of the gravity gradient torque to further stabilize and point spacecraft  100 . Nadir pointing of spacecraft  100  results in lower drag for extended missions, requiring less fuel: By separating the propulsion system from the bus, the change in the size of the propulsion tank for longer and/or manevering missions is accomplished without affecting the size of the bus. Further, the amount of fuel required for attitude control is substantially reduced due to the advantage gained from the moment arm. 
         [0025]    In one embodiment, micro-satellite is scaled to fit into the ESPA ring for a rideshare launch. Such an embodiment is shown in  FIG. 5 , where one or more satellites  100  are mounted on a standard ESPA ring  302  as part of a payload on a launch vehicle  300 . Each ring  302  includes one or more clamp mechanisms  304  for holding the satellite  100  in place in ring  302 . 
         [0026]    In cone example embodiment, ring  302  is 1.5 meters in diameter while clamp mechanisms  304  have an internal diameter of approximately 38 cm. In such one embodiment bus  106  is designed to fit within clamp mechanism  304 . In the embodiment shown in  FIG. 5 , aperture  102  is 90 cm in diameter. In operation, ring  302  is mounted on an LV adapter  306  within launch vehicle  300 . 
         [0027]    In one embodiment, the primary reflector (or the parabolic aperture) is designed to be deployable in space such that very large aperture may be employed to achieve even better imaging quality while being able to stow the micro-satellite in the launch vehicle occupying much smaller volume. 
         [0028]    The micro-satellites described above provide a totally integrated solution of the payload and the bus functions that is scalable to fit in different air launch vehicles. It is capable of accommodating much larger apertures and, therefore, delivers superior imaging capability with better resolution. As is illustrated in  FIG. 3 , each micro-satellite can be tailored to fit its particular launch vehicle fairing in order to deliver the maximum possible capability in terms of resolution and data transfer. Similarly, as is illustrated in  FIG. 5 , each micro-satellite can be tailored to fit in an ESPA ring for lauch with a primary payload. In both case, the axially symmetric approach allows high degree of stability by spinning the satellite; stability can be further improved with gravity gradient assist. The overall impact is to deliver superior earth imagery anywhere on earth at a much lower cost and on demand. 
         [0029]    The nadir-pointed, axially symmetric profile of the satellite results in. lower drag and better stability for extended life. In addition, mounting a solar array on an axially symmetric surface removes the requirement for articulation for Sun pointing of the solar array. Furthermore, the micro-satellites described above demonstrate better imaging resolution with lower system mass, with imaging in some example embodiments to NIIRS 5 level or above. 
         [0030]    To date, there is no scalable means of integrating the payload and the bus into an airborne launch vehicle while effectively utilizing the shape and volume of the launch vehicle and, at the same time maximizing the overall mission imaging capability. The present system and method provide such a scalable means. Also, to date, there is no scalable means of integrating the payload and the bus into a micro-satellite which can be scaled as needed for a rideshare launch. The present system and method provide such scalable means. 
         [0031]    The micro-satellites described above are scalable to fit inside a variety of air launch vehicles and tailorable to meet specific need. The integrated approach makes it so that a satellite can be launched at any time, and in some cases, within an hour of call up. They also can be launched from a variety of launch sites via a variety of airborne launch vehicles, which further increases flexibility of launch angle and launch altitude. 
         [0032]    Although specific embodiments have been illustrated and described herein, it will be appreciated by those of ordinary skill in the art that any arrangement which is calculated to achieve the same purpose may be substituted for the specific embodiments shown. The invention may be implemented in various modules and in hardware, software, and various combinations thereof, and any combination of the features described in the examples presented herein is explicitly contemplated as an additional example embodiment. This application is intended to cover any adaptations or variations of the example embodiments of the invention described herein. It is intended that this invention be limited only by the claims, and the full scope of equivalents thereof.