Small satellite communications antenna and container deployment mechanism

Abstract

A dual-use spring attached to a small satellite that performs two common functions for small satellites including operating as a communications antenna for the small satellite, which eliminates the need for a separate antenna deployment step, and ejecting the small satellite from a modified deployment container mounted on a launch vehicle. The deployment container is modified by removing a conventional deployment spring and pusher plate, which increases available container space.

Claims

1. A satellite dispensing system, comprising: a modified satellite deployment container to accommodate a satellite, the modified satellite deployment container structured without a conventional satellite ejection mechanism and having a door at one end and a back panel at an end opposite the door; and a dual-use spring attached externally to an end panel of the satellite to function as both a communications antenna and as a satellite deployment mechanism, wherein the satellite is positioned inside the modified deployment container with the modified deployment container door closed so that the dual-use spring is compressed against the back panel of the modified deployment container, and at a desired orbit, the dual-use spring applies a force to the satellite to eject the satellite from the modified deployment container when the modified deployment container door opens, the dual-use spring being carried with the satellite and uncompressing to an antenna operational length as it pushes the small satellite out the modified deployment container to function as the satellite's communications antenna.

2. The satellite dispensing system of claim 1, wherein the end panel of the satellite functions as an antenna reflector plate.

3. The satellite dispensing system of claim 1, wherein the dual-use spring has mechanical properties sufficient to provide a force and a velocity needed for deployment of the satellite and electromagnetic properties needed to function as the communications antenna.

4. The satellite dispensing system of claim 1, wherein the dual-use spring has a helical shape and is made of a material that provides electrical conductivity and electromagnetic radiation properties sufficient to function as the satellite communications antenna.

5. The satellite dispensing system of claim 1, wherein the dual-use spring is made from a material that provides desired flexibility, stiffness, and spring energy needed for compression and expansion to a desired deployment length and shape to function as the satellite deployment mechanism, and is covered with an electrically conductive material to function as the satellite communications antenna.

6. The satellite dispensing system of claim 1, further comprising an antenna connector affixed to the end panel of the satellite to which the dual-use spring is attached.

7. The satellite dispensing system of claim 1, wherein the deployment container further comprises a signal-controlled door release mechanism that secures the door in a closed position and opens the door once the desired orbit is reached.

8. The satellite dispensing system of claim 1, wherein the satellite is a CubeSat.

9. The satellite dispensing system of claim 8, wherein the CubeSat and the modified deployment container both have a size of 3 U.

10. The satellite dispensing system of claim 1, wherein the dual-use spring is a right-hand wound conical spring for wideband communications.

11. The satellite dispensing system of claim 1, wherein the dual-use spring eliminates a need for an additional antenna deploying mechanism that unfurls a stowed antenna, along with associated hardware and logic, to provide additional volume within the deployment container.

12. A satellite dispensing system for deploying one or more small satellites, comprising: one or more small satellites, an end one of the satellites being a dispensing satellite; a modified deployment container to accommodate the one or more satellites, the modified deployment container structured without a conventional satellite ejection mechanism and having a door at one end through which the one or more small satellites are deployed and a back panel at an end opposite the door; and a dual-use spring coupled externally to the dispensing satellite at an end panel of the dispensing satellite, wherein the one or more small satellites are pushed into the modified deployment container with the dispensing satellite being positioned closest to the back panel of the modified deployment container so that the dual-use spring is adjacent to the back panel, and the door is latched so that the dual-use spring is in a compressed state, the dual-use spring provides a force to eject the one or more small satellites from the modified deployment container once a desired orbit is reached and the modified deployment container door is opened, and the dual-use spring functions as a communications antenna for the dispensing satellite after ejecting the one or more small satellites from the modified deployment container.

13. The satellite dispensing system of claim 12, wherein the end panel of the dispensing satellite functions as an antenna reflector plate.

14. The satellite dispensing system of claim 12, wherein the dual-use spring is made of a material that provides mechanical properties sufficient to provide a force and a deployment velocity needed for deployment of the one or more satellites and electrical conductivity and electromagnetic radiation properties needed to function as the communications antenna for the dispensing satellite.

15. The satellite dispensing system of claim 12, wherein the dual-use spring is made from a material that provides desired flexibility, stiffness, and spring energy needed for compression and expansion to a desired deployment length and shape to function as the deployment mechanism for the one or more satellites, and is covered with an electrically conductive material to function as the communications antenna for the dispensing satellite.

16. The satellite dispensing system of claim 12, wherein the modified deployment container further comprises a signal-controlled door release mechanism that secures the door in a closed position and opens the door once the desired orbit is reached.

17. The satellite dispensing system of claim 12, wherein the one or more satellites are CubeSats.

18. A method of deploying a small satellite, comprising: placing a small satellite in a stowed configuration; attaching a dual-use spring to an end panel of the small satellite; providing a modified deployment container with a door at one end and a back panel at an end opposite the door, and structuring the modified deployment container without a conventional satellite ejection mechanism; placing the stowed small satellite with the dual-use spring into the modified deployment container so that the dual-use spring is adjacent to the back panel of the modified deployment container; pushing the stowed small satellite into the modified deployment container until the door of the modified deployment container can be closed to secure the small satellite inside the modified deployment container and compress the dual-use spring against the back panel; installing the modified deployment container on board a launch vehicle and launching the launch vehicle into space; and opening the door of the modified deployment container once a desired orbit is reached and ejecting, by force of the compressed dual-use spring, the stowed small satellite from the modified deployment container, the dual-use spring remaining attached to the small satellite, uncompressing into an operating length as it pushes the small satellite out the modified deployment container, and providing antenna communications, wherein the dual-use spring eliminates an additional deployment stage of unfurling a stowed antenna, along with associated hardware and logic, providing additional volume within the modified deployment container.

19. The method of deploying a small satellite of claim 18, further comprising fabricating the dual-use spring from a material that provides mechanical properties sufficient to provide a force and a deployment velocity needed for deployment of the more satellite and electrical conductivity and electromagnetic radiation properties needed to function as the communications antenna for the small satellite.

20. The method of deploying a small satellite of claim 18, further comprising fabricating the dual-use spring from a material that provides desired flexibility, stiffness, and spring energy needed for compression and expansion to a desired deployment length and shape to function as the deployment mechanism for the small satellite, and covering the dual-use spring with an electrically conductive material to function as the communications antenna for the small satellite.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Various aspects and advantages of the invention will become apparent from the following detailed description, taken in conjunction with the accompanying drawings. The drawings are not necessarily drawn to scale. In the drawings:

(2) FIG. 1 illustrates a side view of a conventional deployment container with a side panel removed showing a small satellite, a pusher plate, and a deployment spring;

(3) FIG. 2 illustrates a conventional launch vehicle with an upper stage and a fairing;

(4) FIG. 3 illustrates a small satellite with a dual-use spring according to an embodiment of the present invention;

(5) FIG. 4A illustrates a side view of a closed deployment container with a side panel removed showing the small satellite and the dual-use spring of FIG. 3, with the dual-use spring fully compressed, according to an embodiment of the present invention;

(6) FIG. 4B illustrates a side view of the deployment container, the small satellite, and the dual-use spring of FIG. 4A, with a door of the deployment container open and the small satellite beginning to be ejected from the deployment container by the expanding dual-use spring, according to an embodiment of the present invention;

(7) FIG. 4C illustrates a side view of the deployment container, the small satellite, and the dual-use spring of FIG. 4B, with the small satellite and the dual-use spring being further ejected from the deployment container by the dual-use spring, according to an embodiment of the present invention;

(8) FIG. 4D illustrates a side view of the deployment container, the small satellite, and the dual-use spring of FIG. 4C, with the small satellite and the dual-use spring being completely ejected from the deployment container, according to an embodiment of the present invention;

(9) FIG. 5A illustrates a side view of the deployment container, two small satellites, and the dual-use spring, with a door of the deployment container open and the small satellites beginning to be ejected from the deployment container by the expanding dual-use spring, according to an embodiment of the present invention;

(10) FIG. 5B illustrates a side view of the deployment container, the two small satellites, and the dual-use spring of FIG. 5A, with the small satellites and the dual-use spring being further ejected from the deployment container by the dual-use spring, according to an embodiment of the present invention; and

(11) FIG. 6 illustrates dimensions, pitch, and pitch angle of the dual-use spring, according to an embodiment of the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

(12) The invention is a dual-use spring, functioning as a communications antenna and container deployment mechanism that satisfies container deployment requirements and small satellite radio frequency communications requirements while repurposing approximately 12% of existing container volume that is unused after small satellite deployment.

(13) FIG. 1 illustrates a side view of a conventional deployment container 100 with a front side panel removed so that a small satellite 110, which has been loaded through a door 120, is visible. A typical small satellite mission begins with the completed satellite 110 put into a “stowed” configuration, meaning hardware such as solar panels and antennas (not shown in FIG. 1) are compacted within or around the small satellite 110 to fit inside the deployment container 100. In deployment containers such as the P-POD, a certain amount of volume 130 at an end 140 of the deployment container 100 opposite the door 120 is reserved for an ejection mechanism that includes a compressed deployment spring 150 and a pusher plate 160.

(14) The spring 150 is attached internally to the end 140 of the deployment container 100 and to the pusher plate 160. The small satellite 110 is integrated to the container 100 by pushing the satellite 110 into the deployment container 100 and against the pusher plate 160 to compress the spring 150, and then closing the deployment door 120. The container 100 is integrated into a rideshare satellite launcher 200 in an upper stage 210 of a launch vehicle (lower stages not illustrated) or a launch vehicle fairing 220, as illustrated in FIG. 2, and the ensemble is launched. Once the desired orbit is reached, the door 120 opens and the small satellite 110 is ejected via the spring 150 and the pusher plate 160 providing deployment velocity and force on orbit. The spring 150 and the pusher plate 160 are now useless weight on the launch vehicle upper stage.

(15) After the small satellite 110 is ejected from the deployment container 100, a dormant state clock begins counting down. The small satellite 110 is required to be in a dormant state (i.e., remain in stowed configuration) until a certain amount of time has elapsed. This is typically a requirement to avoid unintended interference due to emissions from the ejected small satellite 110 and to avoid potential deployment failure modes (e.g., a deployable mechanism snagging on the deployment container 100 and becoming wedged). The small satellite 110 then undergoes a secondary deployment sequence in which appendages such as antennas and solar panels are released into an on-orbit configuration.

(16) After the small satellite 110 has separated from the deployment container 100, a small satellite appendage such as an antenna typically needs a separate spring or a motor to supply the motive force needed to extend the antenna from its collapsed, stowed state. Any mechanical failure may result in the antenna not being properly deployed. Other approaches use self-deploying appendages that are compacted (e.g., collapsed or folded) and stowed before launch and deploy after the small satellite 110 has separated from the deployment container 100 using stored strain energy resulting from the compaction. Either approach requires two deployment sequences—one for the satellite and a secondary one for the antenna.

(17) FIG. 3 illustrates an example dual-use spring 300 and a CubeSat 310 in accordance with an embodiment of the invention. In the embodiment shown in FIG. 3, the example CubeSat 310 has a size of 3 U in accordance with the CubeSat standard.

(18) The dual-use spring 300 is affixed externally to an end panel 320 of the CubeSat 310. The end panel 320 of the CubeSat 310 functions as a conductive antenna reflector plate and the interface for the antenna connector 330. Alternatively, radials or netting (not shown) may be used to extend a separate reflector plate (also referred to as a ground plane) to increase directivity of the dual-use spring 300 as an antenna. The CubeSat 310 is stowed as described above except for the dual-use spring 300, which remains connected to, but outside, the CubeSat 310.

(19) FIGS. 4A-4D illustrate an example deployment container 400 in accordance with an embodiment of the invention with a side panel removed to view a payload area within the container 400. The deployment container 400 has been modified by removing the deployment spring 150 and the pusher plate 160 from the conventional deployment container 100. In the embodiment shown in FIGS. 4A-4D, the deployment container 400 has a size 3 U in accordance with the CubeSat standard and accommodates the 3 U CubeSat 310.

(20) Not shown are components such as connectors and sockets that provide communications with the launch vehicle or mechanisms for mounting the deployment container 400 to the launch vehicle.

(21) The CubeSat 310 with the dual-use spring 300 is positioned in the deployment container 400 such that the dual-use spring 300 is between the CubeSat 310 and a back panel 410 of the deployment container 400. FIG. 4A illustrates the CubeSat 310 integrated to the deployment container 400 with the dual-use spring 300 in a compressed state. The CubeSat 310 is pushed into the deployment container 400 on guide rails (not shown) so that the dual-use spring 300 is compressed, and then a spring-loaded deployment door 420 is closed and latched so that the dual-use spring 300 remains compressed, securing the CubeSat 310 within the deployment container 400. The deployment door 420 is coupled to the deployment container 400 by a hinge 430, and may rotate about the hinge 430 to open/close the deployment container 400. A signal controlled door release mechanism 440 controls the opening of the deployment door 420 (electrical connections not shown) once the desired orbit is reached. The deployment container 400 is integrated to the launch vehicle and the ensemble is launched as described above.

(22) FIGS. 4B and 4C illustrate deployment of the CubeSat 310 immediately after the deployment door 420 is opened. Once the desired orbit is reached, the launch vehicle sends an electrical deployment signal to the door release mechanism 440 to release the deployment door 420, which swings open at least 90 degrees, and the CubeSat 310 is ejected via the dual-use spring 300. The dual-use spring 300 operates as a satellite deployment mechanism by supplying the spring force needed to push the CubeSat 310 toward a passageway created by the open door 420, which ejects the CubeSat 310 from the deployment container 400 and deploys the CubeSat 310 in satellite orbit. FIGS. 4B and 4C illustrate the dual-use spring 300 expanding to push the CubeSat 310 out the deployment door 420.

(23) FIG. 4D illustrates the CubeSat 310 after it has completely separated from the deployment container 400 with the dual-use spring 300 deployed to its operational length. The dual-use spring 300 is carried with the CubeSat 310 as the dormant state clock counts down rather than being left behind on the deployment container 400 as useless volume and mass. Once the requisite time has elapsed, the dual-use spring 300 performs the function of the communications antenna for the CubeSat 310.

(24) Thus, the dual-use spring 300 operates as both the satellite deployment mechanism and as the communications antenna, while reducing technical risk and mechanical failure by removing the antenna deployment step, associated hardware, and logic.

(25) When more than one CubeSat is loaded into the deployment container 400, only the CubeSat closest to the back panel 410 of the deployment container 400 would be attached to the dual-use spring 300, which would supply the force needed to eject all the CubeSats from the deployment container 400. FIGS. 5A and 5B illustrate two CubeSats, a 1 U CubeSat 500 and a 2 U CubeSat 510 with separator springs 520, being ejected from the deployment container 400 via the dual-use spring 300. In the embodiment shown in FIGS. 5A and 5B, the dual-use spring 300 is attached to the 1 U CubeSat 500 and functions as the communications antenna for the 1 U CubeSat 500.

(26) The dual-use spring 300 is designed to have the mechanical properties needed to accomplish the force and deployment velocity for the deployment container 400, as well as the electromagnetic properties required for the desired small satellite radio frequency communications antenna that can be compressed and stowed to occupy a very small, compact volume within the deployment container 400.

(27) The dual-use spring 300 is fabricated from a material such as aluminum or a copper alloy to provide good electromagnetic radiator and spring characteristics. However, as will be appreciated by those skilled in the art, other materials may also be applicable to provide these features. Common steel springs can deliver the required mechanical properties but are typically not desirable due to low electrical conductivity.

(28) Alternatively, the dual-use spring 300 may be made from a material that gives the desired flexibility and stiffness to operate as the deployment mechanism, and spring energy necessary for compression and for returning to the desired deployed length and shape, but that is covered with an electrically conductive material, such as copper tape, to provide both electrical conductivity and spring characteristics.

(29) The electrical performance of the axial mode, dual-use helical spring 300 is influenced by its geometric parameter values and can be tuned as required. Referring to FIG. 6, the wire diameter 600 and the spring diameter 610 set the center frequency and bandwidth. The spring diameter 610 is bounded by the diameter of the deployment container 400 and must be less than 10 cm. The spring length 620, number of turns, vertical separation 630, and pitch angle 640 between turns are also important for beam shaping and spring force. The dual-use spring 300 provides a communications frequency of approximately 3 GHz or less and a higher antenna gain than traditional CubeSat antennas while avoiding the complexity, mass, and volume problems associated with an extending, expanding, or unfurling antenna. For example, an antenna designed with a conductor diameter 600 of 6 mm to operate at 1 GHz with 10 turns and a spacing 630 of 0.23 wavelengths and meeting the diameter 610 constraint of 10 cm results in a gain of 13.2 dBi, a half power beam width of 34.2 degrees, and a bandwidth of 136 MHz. The ejection speed of the CubeSat 310 from the deployment container 400 is dependent on the stiffness and size of the coils and should be fixed according to the deployment speed specification. The deployment speed is on the order of 1.5 m/s for standard mass CubeSats.

(30) The dual-use spring 300 can be fabricated in constant pitch, conical, barrel, hourglass, or variable pitch configurations to achieve the desired force, velocity, and radiation pattern. Any of these configurations may be preferred based on the communications application. Additionally, the dual-use spring 300 can be left-handed or right-handed depending on preferred polarization. In an alternative embodiment, the dual-use spring 300 is a right-hand wound conical spring for a wideband communications application.

(31) The dual-use spring 300 and corresponding deployment container 400 of the invention provide several advantages, including the following:

(32) (a) The dual-use spring 300 functions as both a helical communications antenna and as the satellite ejection mechanism, replacing the conventional container deployment spring and pusher plate.

(33) (b) The invention better utilizes limited deployment container volume which normally forces the CubeSat-Class spacecraft to achieve high mass and volume efficiency. The invention increases volume utilization in containerized satellites by up to approximately 12%.

(34) (c) The invention obviates the requirement for a second antenna deployment stage, which reduces technical and mechanical risk inherent to a second deployment by removing the antenna deployment step, associated hardware, and logic. This eliminates additional mass and volume of a secondary deployment system.

(35) (d) The invention simplifies small satellite design because a separate antenna deploying mechanism is not required, enabling simple and reliable antenna deployment from a compressed state.

(36) While the foregoing written description of the invention enables one of ordinary skill to make and use what is described herein, those skilled in the art will understand and appreciate the existence of variations, combinations, and equivalents of the disclosed embodiments and methods. The invention should therefore not be limited by the above description, but by all embodiments and methods within the scope and spirit of the invention as disclosed.