Satellite apparatus, satellite assembly, and method of deploying a satellite from a launch vehicle

Abstract

A satellite apparatus, a satellite assembly, and a method of deploying a satellite from a launch vehicle are disclosed. The satellite apparatus includes a housing having first and second walls opposite each other and a support structure spanning the first and second walls within the housing. The support structure is structurally connected to the housing only at the first and second walls, and end portions of the support structure are configured to be connected to the launch vehicle by a separation system.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to a satellite apparatus, and more particularly, to a satellite apparatus including a housing and a support structure and a method of stacking the satellite apparatus. BACKGROUND

[0002] To offset the high launch costs, an increasing number of satellites are being deployed from a single launch vehicle. Smaller satellites, such as microsats and CubeSats, are particularly good candidates for multi-payload single launch. However, the intense and highly directional forces of launch require such mounting structures and satellite designs that are specifically designed to efficiently transfer loads to the launch vehicle structure and withstand conditions such as high bending or torsional loads resulting from the chosen mounting arrangement and orientation. Efficient structural designs and arrangements are needed that maximize the use of the limited launch vehicle space and allow for efficient deployment in space while minimizing weight, production cost, and launch preparation time. SUMMARY

[0003] The present disclosure provides systems, apparatuses, and methods relating to satellite support structures and assemblies. In some examples, a satellite apparatus can include a housing having opposing first and second walls, and a support structure within the housing spanning the first and second walls. The support structure can be structurally connected to the housing only at the first and second walls, and a proximal portion of the support structure can be configured to be connected to a launch vehicle by a separation system.

[0004] In some examples, a satellite assembly can include a launch vehicle having a launch axis and a plurality of satellites inside the launch vehicle. The plurality of satellites can be stacked horizontally relative to the launch axis.

[0005] In some examples, a method of deploying a satellite from a launch vehicle can include loading a plurality of satellites inside a launch vehicle by stacking the satellites horizontally relative to a vertical launch axis. The method can also include transporting the satellites in the launch vehicle to space, and separating the satellites from the launch vehicle horizontally relative to the vertical launch axis.

[0006] The features, functions, and advantages of the present disclosure can be independently implemented in various examples of the present disclosure, or can be combined with other examples, and further details of the present disclosure can be seen with reference to the following description and drawings. BRIEF DESCRIPTION OF DRAWINGS

[0007] Figure 1 is a schematic illustration of an illustrative satellite launched and deployed from a launch vehicle.

[0008] Figure 2 is a block diagram of a satellite of Figure 1 is a block diagram of a satellite of

[0009] Figure 3 Isometric view of a plurality of illustrative satellite stacks according to aspects of the present disclosure mounted to a launch vehicle payload adapter according to aspects of the present disclosure.

[0010] Figure 4 Is an isometric front view of a proximal satellite of a satellite assembly of Figure 3

[0011] Figure 5 Is an isometric rear view of a satellite of Figure 4

[0012] Figure 6 Is an isometric front view of a core structure, fuel tank, and shell panels of a satellite of Figure 4

[0013] Figure 7 Is an isometric view of a shell of a satellite of Figure 4

[0014] Figure 8 Is an isometric mapping of an illustrative satellite stack including a core structure of a proximal satellite of Figure 4

[0015] Figure 9 Is a cross-sectional view of a mounting plate and core structure of Figure 8

[0016] Figure 10 Is a cross-sectional view of a satellite stack of Figure 8

[0017] Figure 11 Is a flowchart depicting steps of an illustrative method of deploying a satellite from a launch vehicle according to the present teachings. DETAILED DESCRIPTION

[0018] ​​​​​​​Various aspects and examples of satellites having cylindrical support structures and related stacked satellite assemblies and methods are described below and illustrated in the associated drawings. Unless otherwise indicated, satellites and / or satellite assemblies according to the present teachings, and / or various components thereof, can but need not include at least one of the structures, components, functions, and / or variations described, shown, and / or combined herein. Moreover, unless specifically excluded, process steps, structures, components, functions, and / or variations described, shown, and / or combined herein can include other similar devices and methods, including those interchangeable between examples disclosed. The following description of various examples is merely illustrative in nature and is in no way intended to limit the disclosure, its application or uses. Additionally, the advantages provided by the examples described below are illustrative only and are not all-inclusive.

[0019] This detailed description includes the following sections, which immediately follow the summary below as described hereinafter: (1) Overview; (2) Examples, Components, and Alternatives; (3) Illustrative Combinations and Additional Examples; (4) Advantages, Features, and Benefits; and (5) Conclusion. The Examples, Components, and Alternatives section is further divided into sub-sections A through C, each of which is labeled accordingly.

[0020] Overview

[0021] Generally, a satellite according to the present teachings can include a hollow central support structure and include a shell. The central support structure can be a primary structure of the satellite that supports the shell and is connected to a launch vehicle. The central support structure can span between first and second panels of the shell and can be connected to the shell only through the first and second panels. A payload and operational equipment of the satellite can be supported by the shell. The central support structure can be additively manufactured and include a cylindrical wall having an array of diamond-shaped apertures. For launch, the satellite can form part of a structural satellite launch configuration.

[0022] Generally, a structural satellite launch configuration according to the present teachings can include two satellites, each satellite having a central support structure that is a primary structure of that satellite. The central support structures of the two satellites can be connected to form a single beam structure that can be installed to a payload adapter of a launch vehicle as a cantilever beam. In other words, the central support structures of the two satellites can define a core axis, and the satellites can be installed in the launch vehicle such that the core axis is perpendicular to a launch axis of the launch vehicle.

[0023] In some examples, the structural satellite launch configuration can include multiple pairs of connected satellites. Each pair of connected satellites can be connected to a central ring payload adapter of a launch vehicle, the connected satellites extending radially outward from the ring adapter. In some examples, the structural satellite launch configuration can include three or more stacks of one or more pairs of connected satellites with a central support structure, the one or more stacks mounted to a launch vehicle such that a core axis defined by the central support structure is perpendicular to a launch axis of the launch vehicle.

[0024] Examples, Components, and Alternatives

[0025] The following sections describe selected aspects of illustrative satellites and related components and / or methods. The examples in these sections are intended to be illustrative and should not be construed as limiting the overall scope of the present disclosure. Each section can include one or more distinct examples, and / or contextual or related information, functionality, and / or structure.

[0026] A. Illustrative Satellites and Related Methods

[0027] The examples disclosed herein can be described in the context of an illustrative satellite launch method 80 (see Figure 1 ) and an illustrative satellite 100 (see Figure 2 ). In the present example, the method 80 includes three phases: a launch phase 20, a separation or deployment phase 40, and an operations phase 60. The launch phase 20 can include transporting the satellite 100 (or spacecraft) from a planetary body 120, such as the Earth, to outer space 122 using a launch vehicle 124. In the context of the Earth, outer space can include regions beyond the Karman line. The deployment phase 40 can include separating the satellite 100 from the launch vehicle 124 once a desired position, trajectory, and / or orbit has been achieved. The operations phase 60 can include preparing the satellite 100 for operations, such as establishing communication with a controller on the planetary body 120, extending solar panels or instrument arms, and / or maneuvering to a desired orientation relative to the planetary body. In some examples, the method can also include design, production, and / or service phases.

[0028] Each process of the method 80 can be performed or implemented by a system integrator, third parties, and / or an operator (e.g., a customer). For the purposes of the present specification, a system integrator can include, without limitation, any number of aerospace manufacturers and prime system subcontractors; third parties can include, without limitation, any number of resellers, subcontractors, and suppliers; and an operator can be a communications company, a leasing company, a military entity, a service organization, etc.

[0029] As Figure 2As shown, satellite 100 can include a bus 102 having a plurality of satellite systems and include a payload 104 and a separation system 106. Examples of the plurality of systems include one or more of a primary structure 108, a propulsion system 110, a power system 112, a thermal management system 114, a radiation shielding system 116, and a communication system 118. Each system can include various subsystems, e.g., controllers, processors, actuators, effectors, motors, generators, etc., depending on the functionality involved. Any number of other systems can be included. Although an example of an unmanned artificial satellite is shown, the principles disclosed herein can be applied to other aerospace vehicles and technologies, e.g., launch vehicles, space stations, manned spacecraft, and / or interstellar probes.

[0030] The apparatus and methods shown or described herein can be employed during any one or more stages of satellite launch method 80. For example, during launch stage 20, two or more satellites can be stacked perpendicular to the launch axis of launch vehicle 124. Similarly, one or more instances of the apparatus or methods implemented or combinations thereof can also be utilized, for example but not limited to, when satellite 100 and / or launch vehicle 124 are in preparation prior to performing launch method 80. Likewise, during deployment stage 40, the apparatus, one or more instances of the methods, or combinations thereof can be utilized, for example, by deploying satellites radially outward from launch vehicle 124 perpendicular to the launch axis of launch vehicle 124.

[0031] B. Illustrative Satellite Assembly

[0032] As Figures 3 to 6 shown, this section describes an illustrative satellite assembly 200. Satellite assembly 200 is an example of a structural satellite launch configuration as described above. The assembly includes a plurality of satellite stacks 210. Each stack 210 includes a proximal satellite 212 and a distal satellite 214, and each stack is connected to a payload adapter 216 by a mounting plate 218. In some examples, a stack can include three or more satellites.

[0033] In Figure 3 the example shown, payload adapter 216 includes a ring structure 217, e.g., an Evolved Secondary Payload Adapter (ESPA) manufactured by Moog, Inc. The plurality of satellite stacks 210 are connected to ring structure 217 at six mounting points 220 arranged symmetrically around the ring structure. In Figure 3 one of the satellite stacks 210 is not depicted to show the corresponding mounting point 220. In general, the plurality of satellite stacks can be arranged symmetrically around payload adapter 216 in order to balance the loads imparted to the payload adapter.

[0034] The payload adapter 216 is part of the launch vehicle having a launch axis 222. The launch axis can also be described as the longitudinal axis, z-axis, or vertical axis of the launch vehicle. Directions perpendicular to the launch axis can be described as lateral and / or horizontal.

[0035] Prior to launch, the launch axis can be aligned with the vertical direction as defined by a gravitational reference frame. During launch, the launch axis can rotate relative to the gravitational reference frame as the launch vehicle moves along a non-linear launch trajectory. Accordingly, for clarity in the following description, directional terms such as "up", "down", "top", "bottom", and the like should be understood with respect to the launch axis.

[0036] In the present example, the ring structure 217 of the payload adapter 216 has a central axis 224 that is parallel to the launch axis 222. Each satellite stack 210 has a core axis 226 that can also be described as the longitudinal axis or central axis of the stack. The core axis 226 of each satellite stack 210 extends through a center point 228 of the ring structure 217 that is on the central axis 224 of the ring structure. That is, the core axes of the plurality of satellite stacks intersect at the center point of the ring structure.

[0037] The plurality of satellite stacks 210 can be described as horizontal stacks, branches, protruding assemblies, and / or radially connected satellite groups. Each satellite stack 210 extends radially outward from the ring structure 217 perpendicular to the central axis 224 of the ring structure. That is, the core axis 226 of each satellite stack is perpendicular to the launch axis 222.

[0038] The proximal satellites 212 of each satellite stack are releasably connected to a corresponding mounting plate 218 by a separation system and / or device as discussed further below. The distal satellites 214 of each satellite stack are similarly releasably connected to a corresponding proximal satellite by the separation system and / or device. Each mounting plate 218 is fixedly attached to one of the mounting points 220 of the ring structure 217. In the present example, the mounting plates are bolted to the ring structure. In some examples, the mounting plates can be an integral part of the payload adapter 216 and / or the proximal satellites can be directly connected to the mounting points. In some examples, the mounting plates 218 can support other additional payloads or launch vehicle components, and / or can form part of another structure.

[0039] Figures 8 to 10 One satellite stack 210 is depicted, and Figures 4 to 7The proximal satellite 212 of the satellite stack is depicted. Except where otherwise noted, it is understood that the description of the satellite stack applies equally to each satellite stack 210. In general, a satellite of a satellite assembly as described herein can include a primary structure as described below, but can vary in the specifications of the payload, the housing design, and the operational systems such as communications, shielding, and thermal regulation.

[0040] Figure 4 and Figure 5 are opposite isometric views of the proximal satellite 212. In the case of the orientation of the satellite relative to the payload adapter, Figure 4 depicts the distal side of the satellite, Figure 5 depicts the proximal side of the satellite. In the present example, the proximal satellite 212 is generally cubical, and includes a housing 234 composed of six planar and square or rectangular wall panels. More specifically, the satellite includes a front panel 236 and an opposite rear panel 238. Four equipment panels 240 span between the front and rear panels.

[0041] The proximal satellite 212 also includes a plurality of patch antennas 242, a cylindrical baffle 244, and a solar panel 246 composed of two deployable panels. Four thrusters 248 are mounted in brackets 250 at the four corners of the rear panel 238. The satellite 212 can also include any appropriate operational or payload equipment, including but not limited to fuel tanks, star trackers, reaction wheels, heat sinks, heat sink panels, and / or avionics. Most of the equipment can be mounted to the interior surfaces of the equipment panels 240.

[0042] Figure 5 A proximal end portion 262 of a cylindrical core structure 260 of the proximal satellite 212 is also shown extending through the rear panel 238. In Figure 6 The core structure 260 can be seen more completely in FIG. 2B, in which the housing 234 is depicted as transparent. The core structure can also be described as a support structure, a hollow column, and / or a central beam.

[0043] The core structure 260 is a hollow cylinder that spans between the rear panel 238 and the front panel 236 and is enclosed in the housing 234. The core structure can also be described as a hollow column. The core structure defines a central axis 270 and is centered in the proximal satellite 212. A proximal end portion 262 is fixed to the rear panel 238, and a distal end portion 264 of the core structure is fixed to the front panel 236. A wall 266 including a plurality of apertures 268 extends between the two end portions. The wall 266 is relatively thin with respect to the diameter of the core structure, allowing the core structure to be strong and rigid but relatively light. The apertures 268 can further lighten the core structure without sacrificing desirable structural properties.

[0044] The core structure 260 functions at the main structure of the satellite and is configured to structurally connect the satellite to both the launch vehicle and the distant satellite. More specifically, the near-end portion 262 is configured to connect to the launch vehicle's payload adapter via a mounting plate, and the far-end portion 264 is configured to connect to the core structure of another satellite. Each end portion of the core structure 260 is configured to be connected via a separation system or device. In this example, the two end portions are configured to be connected via a similar separation system, as further described below. In some examples, the near-end portion and the far-end portion may be configured to be connected via different separation systems or devices.

[0045] The distal portion 264 includes an interface flange 272 that contacts the inner surface of the front panel 236. In this example, the core structure is secured to the front panel by a plurality of fasteners that extend through orifices in the interface flange 272 and into the front panel. The proximal portion 262 includes a plurality of interface tabs surrounding the circumference of the wall 266, extending from the wall and contacting the outer surface of the rear panel 238. In this example, the core structure is secured to the rear panel by a plurality of fasteners that extend through orifices in the interface tabs 274 and into the rear panel.

[0046] like Figure 6 As shown, the near-side satellite 212 includes a fuel tank 280 mounted inside the core structure 260. This arrangement helps maximize space efficiency within the satellite, allowing the core structure 260 to have a large cross-section for increased strength and stiffness without wasting internal space. An aperture 268 facilitates necessary access and / or connection to the fuel tank, such as fluid connection to a fill valve or electrical connection to sensors.

[0047] The outer diameter of the fuel tank 280 may be close to, but smaller than, the inner diameter of the wall 266 of the core structure 260. To allow for a tight-fitting fuel tank positioning within the core structure 260, the core structure comprises two parts that can be assembled around the fuel tank 280. More specifically, the core structure 260 includes a first segment 276 and a second segment 278.

[0048] The first segment 276 includes a proximal portion 262, and the second segment 278 includes a distal portion 264. The first and second segments are bolted together at an intermediate interface 284. The intermediate interface 284 can be described as being located along a portion of the wall 266 in the intermediate section (part-way). Each of the first and second segments is thickened near the intermediate interface 284 to strengthen and reinforce the connection and to allow the two segments to function as a single effective structural support for the satellite 212.

[0049] In the present example, the wall 266 also includes apertures for a plurality of shim bolts 282 to center and precisely position the fuel tank 280 in the core structure. In general, the core structure 260 can include any customizations or modifications suitable for the mounting, support, or integration of the operational equipment of the proximal satellite 212.

[0050] Figure 7 is an isometric view of the housing 234 of the proximal satellite 212, with the panels 236, 238, 240 depicted as transparent. In the present example, the panels of the housing are connected by eight corner brackets 286 and four corner clips 288. The corner brackets 286 include four thruster mounts 250, and each corner bracket is positioned within a corner of the housing 234 at a location where a corner of the front panel 236 or the rear panel 238 intersects with a corner of two equipment panels 240. Each corner clip 288 extends along the inside of an edge of the housing 234 where the edges of two equipment panels 240 meet.

[0051] The body 287 of each corner bracket 286 sits over a corresponding corner cutout of the front panel 236 or the rear panel 238. The sides of the body 287 of the corner bracket contact the interior surfaces of three adjacent panels, and can be bonded or otherwise secured to these panels. The corner brackets 286 also include corner tabs 289 that are configured to contact the interior surfaces of the corner clips 288. For each corner bracket 286, one corner tab 289 can be bonded or otherwise secured to an adjacent corner clip 288. Thus, each corner clip 288 can be secured to a first corner bracket at the rear panel 238 and a second corner bracket at the front panel 236, and extends between the first and second corner brackets.

[0052] The corner clips 288 and the corner brackets 286 can structurally connect the equipment panels 240 to the front panel 236 and the rear panel 238, which in turn are structurally connected to the core structure of the satellite. The equipment panels 240 are directly connected to the core structure. In other words, the core structure is connected to the equipment panels 240 only through the front and rear panels. The load from the equipment mounted to the equipment panels 240 can be transferred through the front and rear panels 236, 238 to the core structure.

[0053] The housing 234 does not form part of the primary structure of the proximal satellite 212. As Figures 8 to 10 shown, the core structure 260 of the proximal satellite 212 is directly connected to the mounting plate 218 and the core structure of the distal satellite. The core structure 260 is the primary structure of the proximal satellite 212, as well as the primary load path to the launch vehicle. The housing 234 encloses and encases the core structure and the satellite equipment. The housing is supported by the core structure 260, and the housing is not directly connected to the launch vehicle. Thus, the freedom of material selection and design of the housing 234 is significantly increased.

[0054] As shown in Figure 7 and described above, in the present example, the shell 234 is generally cuboid and composed of planar panels. In general, the shell can have any shape suitable for housing the satellite system and equipment and efficiently stow within the launch vehicle. For example, the shell can be a polyhedron, can include curved panels, and / or can have an irregular shape. The shell 234 can be generally symmetrical and / or balanced about the central axis 270 to separate directly and without tumbling during deployment.

[0055] In the present example, the panels 236, 238, 240 include composite honeycomb sandwich material. In general, the panels can include any lightweight material or material of sufficient strength to support the installed equipment. For example, the panels can be additively manufactured and / or include additively manufactured portions that can be formed by three-dimensional (3D) printing, laser sintering of metal alloys, or other methods. The panels need not be designed for the strength or stiffness required for the primary structure.

[0056] The shell 234 can be highly customized for the selected payload and operational satellite equipment. Observation ports, supports, shields, access holes, or other modifications can be formed on the shell without affecting the primary structure of the satellite. In particular, in combination with the rapid prototyping and design implementation capabilities of additively manufactured components, this degree of freedom can significantly simplify the design and reduce testing and certification time.

[0057] In the shell 234, the front panel 236 and the rear panel 238 can have the most design constraints. That is, the two panels or structures of the shell 234 that connect to the proximal and distal ends of the core structure can need to be configured to interface with the core structure. In the present example, the core structure interfaces with the front and rear panels as defined by the location and orientation of the satellite thrusters. In some examples, the core structure can connect to the side of the satellite such that the satellite can be described as mounted sideways to the launch vehicle, or can be mounted to any two opposing walls or wall portions of the shell.

[0058] Referring again to Figure 7 , the rear panel 238 includes a circular aperture 290 with six circumferential cuts or grooves 292. The circular aperture 290 and the grooves 292 can allow the core structure 260 to protrude through the rear panel 238 such that the proximal portion 262 is outside of the rear panel, as shown in Figure 5 More specifically, the wall 266 can extend through the circular aperture 290, and the six compartments 294 protruding from the wall 266 can extend through the grooves 292. In other words, the rear panel 238 can include an aperture shaped to correspond with the core structure 260 such that the proximal portion 262 is extendable through the rear panel. The specific shape of the proximal portion 262 and the aperture in the rear panel 238 can depend on the selected separation system, as further described below.

[0059] As shown in FIG. 2, the proximal satellite 212 includes a generally cuboid housing 234 and a cylindrical core structure 260 having a proximal end portion 262 and a distal end portion 264. The core structure 260 spans between a front panel 236 and a rear panel 238, with the proximal end portion 262 extending through the rear panel to connect to the distal end portion 264 of the distal satellite 214. Figure 4 And Figure 7 As shown in FIG. 2, the proximal satellite 212 includes a generally cuboid housing 234 and a cylindrical core structure 260 having a proximal end portion 262 and a distal end portion 264. The core structure 260 spans between a front panel 236 and a rear panel 238, with the proximal end portion 262 extending through the rear panel to connect to the distal end portion 264 of the distal satellite 214.

[0060] Figure 8 The proximal satellite 212 is depicted as part of the satellite stack 210, connected to the distal satellite 214 and the mounting plate 218. Similar to the proximal satellite 212, the distal satellite 214 includes a generally cuboid housing 334 and a cylindrical core structure 360 having a proximal end portion 362 and a distal end portion 364. The core structure 360 spans between a front panel 336 and a rear panel 338, with the proximal end portion 362 extending through the rear panel to connect to the distal end portion 264 of the proximal satellite 214.

[0061] In this example, the distal satellite 214 is generally identical to the proximal satellite 212, except for the construction of the distal end portion 364 of the core structure 360. Thus, the reference numerals for the components of the distal satellite 214 match those of the corresponding components of the proximal satellite 212. In general, the distal satellite 214 can include a core structure 360 that generally matches and is configured to connect to the core structure 260 of the proximal satellite 212, but can differ in design from the proximal satellite 212. For example, the payloads, operational equipment, and / or housings of the two satellites can differ.

[0062] The distal end portion 364 of the distal satellite 214 is constructed more simply than the distal end portion 264 of the proximal satellite 212, as shown more clearly in Figure 9 Unlike the distal end portion 264, in this example, the distal end portion 364 need not be configured to connect to another core structure. Thus, the distal end portion 364 includes a cylindrical wall 366, up to a circular annular interface flange 372, to connect to the front panel of the distal satellite 214. This more simple shape can desirably be lighter. In some examples, to simplify manufacturing and / or satellite design, the distal end portion 364 of the distal satellite 214 can match the distal end portion 264 of the proximal satellite 212. In some examples, the satellite stack 210 can include three or more satellites, with at least one satellite including a core configured to connect to an adjacent satellite at both a proximal end and a distal end.

[0063] In the present example, the distal satellite 214 is connected to the proximal satellite 212 in the same manner as the proximal satellite is connected to the mounting plate 218. Thus, the proximal end portion 362 of the core structure 360 of the distal satellite 214 mates with the proximal end portion 262 of the core structure 260 of the proximal satellite 212. The mounting plate 218 also includes a distal end portion 464 that mates with the distal end portion 264 of the proximal satellite 212.

[0064] The mounting plate 218 can serve as an adapter to facilitate the structural connection between the core structure 260 of the proximal satellite 212 and the launch vehicle payload adapter. The mounting plate includes a proximal end portion 462 configured for connection to mounting points of the payload adapter. In the present example, the proximal end portion 462 includes a square planar surface with bolt holes at each corner. The proximal end portion 462 and the distal end portion 464 are joined by a cylindrical central wall with support struts. In general, the mounting plate 218 can have any geometry or configuration suitable for providing a strong connection and payload path between the core structure of the satellite and the launch vehicle.

[0065] The mounting plate 218, the core structure 260, and the core structure 360 together can serve as a cantilever beam extending horizontally outward from the launch vehicle payload adapter. The combined core structures are sufficiently rigid to support both the proximal satellite 212 and the distal satellite 214 to withstand the bending moments and vibration loads associated with launch. The joined core structures also provide a solid and simple load path to the launch vehicle. As described further below, the dimensions and design of the core structures provide the required rigidity while minimizing weight.

[0066] Also important to the stiffness of the beam affected by the combined core structures is the stiffness of the connections between the core structure 260 and the core structure 360 and between the core structure 260 and the mounting plate 218. Any effective decoupling system or device can be used to connect the satellites. However, a system that provides a direct connection between the core structures, such as depicted in the present example, can provide a sufficiently stiff connection.

[0067] The satellite stack 210 includes a proximal decoupling system 410 connecting the proximal satellite 212 and the mounting plate 218, and a distal decoupling system 412 connecting the distal satellite 214 and the proximal satellite 212. As Figures 4 to 6 As most clearly shown in FIG. 4, each decoupling system 410, 412 includes six separable connectors 414 and three push-off pins 416. Each separable connector 414 includes a male portion 415 and a female portion 417.

[0068] On the proximal satellite 212, the male portions 415 of the separable connectors 414 of the separation system 410 are housed in the compartments 294 of the proximal end portion 262 and protrude through the interface flange 258 of the proximal end portion to engage corresponding female portions on the mounting plate. On the side of the interface flange 272 opposite the front panel 236, the female portions 417 of the separable connectors 414 of the separation system 412 are mounted in recesses in the wall 266 of the distal end portion 264. Corresponding apertures in the interface flange 272 allow the corresponding male portions on the distal satellite to pass through the flange to the female portions 417.

[0069] The push-off pins 416 of the separation system 412 are similarly mounted. That is, the push-off pins are mounted on the side of the interface flange 272 opposite the front panel 236 and these push-off pins extend through corresponding apertures in the interface flange and the front panel to contact the interface flange of the proximal end portion of the distal satellite 214. The interface flange 258 of the proximal end portion 262 of the proximal satellite 212 includes three scallops 420 to engage the push-off pins on the mounting plate 218.

[0070] The separable connectors 414 and the push-off pins 416 of the separation systems 410, 412 are evenly spaced around the circumference of the core structure 260, 360. Each device 414, 416 of the separation system is spring-actuated for smooth and reliable separation and these devices are connected to a control system for coordinated triggering.

[0071] As shown in Figures 9 to 10 , the devices 414, 416 of the separation system 410 are not interposed between the mounting plate 218 and the core structure 260. Similarly, the devices of the separation system 412 are not interposed between the core structure 260 and the core structure 360. The interface flange 258 of the proximal end portion 262 of the core structure 260 directly contacts the mounting plate 218. The interface flange 272 of the distal end portion 264 of the core structure 260 and the interface flange 358 of the proximal end portion 362 of the core structure 360 contact the inner and outer surfaces, respectively, of the front panel 236 of the proximal satellite, with only the front panel between the two flanges. The direct connection between the core structures can result in the desired stiffness. The devices of the separation system can also be individually configured to facilitate the rigid connection.

[0072] In this example, the separation system 412 also includes two separable housing connectors 430, as shown in Figure 8 and Figure 10The housing connectors 430 each include a first bracket 431 mounted to the front panel 236 of the proximal satellite 212 and a second bracket mounted to the rear panel 338 of the distal satellite 214. The first and second brackets 431, 432 are connected by a spring-actuated releasable mechanism similar to the separable connector 414. Together, the two connected brackets 431, 432 have an axial extent that matches the proximal end portion 362 of the distal satellite, allowing the housing connector 430 to bridge between the two satellites when the core structure 260 is connected to the core structure 360.

[0073] The housing connectors 430 can be configured and / or positioned according to the geometry or other characteristics of the housings 234 and 334. In the present example, two housing connectors are positioned at opposite lateral edges of the satellites to provide additional lateral stability to the connection between the satellites and to assist in tumble-free separation.

[0074] Figure 9 is a cross-sectional view of the main structure of the satellite stack 210, including the mounting plate 218, the separation system 410, the core structure 260, the separation system 412, and the core structure 360. Each core structure has an inner diameter 424 defined by the inner surface of the wall 266 or 366. Each core structure 260, 360 also has a length 426 from the interface flange at the proximal end to the interface flange at the distal end. The inner diameter 424 and the length 426 of the core structure 260 and the core structure 360 are each the same.

[0075] In the present example, the core structures have an inner diameter 424 of about 12 inches and a length of about 20 inches. The thickness of the wall 266, 366 of the core structures is between about one one-hundredth and two one-hundredths of an inch. The core structures and the cantilever beam structure formed by connecting the core structures are strong and rigid enough to support each of the proximal and distal satellites at about 100 kilograms at a vibration frequency of about 30 hertz. In general, the core structures 260, 360 can be any size appropriate for the size and weight of the satellites. That is, the core structure design can be adapted for use with micro-satellites to full-size satellites.

[0076] The apertures 268 in the wall 266 and the apertures 368 in the wall 366 can help reduce the weight of the core structures without sacrificing structural strength. In the present example, each core structure includes two arrays of apertures 428, with a first array in the first segment and a second array in the second segment. In some examples, the apertures can be arranged in additional arrays, and in examples such as the monolithic core structure, the apertures can form a single array.

[0077] Each of the orifices 268, 368 is diamond shaped. In the present example, each orifice is approximately two inches in length, and is spaced approximately one quarter inch from an adjacent orifice. The array of orifices 428 can also be described as a grid and / or a diamond lattice. Diamond shapes can be particularly suitable for additive manufacturing. Any desired orifice shape can be used, and the appropriate orifice shape can depend on the manufacturing method selected. The orifice size and spacing can be selected according to the desired structural and / or electromagnetic properties of the core structure.

[0078] In the present example, the core structure 260, 360 is additively manufactured from metal. More specifically, the core structure can include a laser sintered metal alloy that is manufactured using direct metal laser sintering (DMLS) of an aluminum alloy. In general, the core structure can be manufactured according to any effective method and any sufficiently strong and light material. Additive manufacturing of the core structure can be particularly suitable for the formation of thin walls, orifices, and custom interface features.

[0079] Other components of the satellites 212, 214 can advantageously be manufactured using additive manufacturing methods such as DMLS or electron beam melting (EBM). For example, the fuel tanks 280, gimbals 288, gimbled brackets 286, and / or faceplates 236, 238, 240 can be additively manufactured.

[0080] C. Illustrative Methods

[0081] This section describes steps of an illustrative method of deploying a satellite from a launch vehicle; see Figure 11 The above-described aspects of the satellite, the structural satellite launch configuration, and / or the launch vehicle payload adapter and mounting plate can be used in the method steps described below. Where appropriate, reference can be made to components and systems that can be used to perform each step. These references are for illustration, and are not intended to limit the possible ways in which any particular step of the method can be performed.

[0082] Figure 11 is a flow diagram that illustrates the steps executed in an illustrative method, and can not recite the complete process or all of the steps of the method. Although various steps of the method 500 are described below and depicted in Figure 11 , these steps need not necessarily all be performed, and in some cases can be performed simultaneously or in a different order than shown.

[0083] At step 510, the method includes loading a plurality of satellites in a launch vehicle. The launch vehicle can include any launch vehicle suitable for transporting a payload into space. For example, the launch vehicle can be a consumable autonomous vehicle, or can be a manned spacecraft. Step 510 can be performed as part of preparing the launch vehicle for launch, and the plurality of satellites can be configured to be connected to and launched in the launch vehicle. Loading the satellites can include attaching the satellites to a payload adapter of the launch vehicle and / or to each other using one or more separation systems and / or devices. The satellites can be loaded according to sub-steps 512-516 of step 510.

[0084] Sub-step 512 includes horizontally stacking the plurality of satellites relative to a vertical axis of the launch vehicle. In other words, two or more satellites can be positioned adjacent to each other along a horizontal axis. The vertical axis can correspond to an orientation of the launch vehicle during preparation for launch and / or can correspond to a launch direction or launch axis. The vertical axis can also be referred to as a primary axis of the launch vehicle. The two or more satellites can be referred to as a horizontal stack and / or a lateral assembly. Only one of the satellites in the stack can be directly connected to the launch vehicle.

[0085] Sub-step 514 of sub-step 512 includes connecting cylindrical core structures of adjacent satellites of the stacked satellites. A primary structure of each satellite of the plurality of satellites can include a cylindrical core structure. Each core structure can have the same diameter, and the core structures can be configured to be connected to another core structure by a separation system. Within the horizontal stack of satellites, each satellite can be connected to an adjacent satellite by a core structure. The connected core structures of the satellites of the stack can form a crossbeam extending horizontally out from a payload adapter of the launch vehicle.

[0086] Sub-step 516 of step 510 includes assembling and attaching a plurality of stacks around a ring structure of the launch vehicle. The ring structure can be a payload adapter of the launch vehicle, and can include a plurality of attachment points or mounting points. The plurality of horizontal stacks can be assembled according to sub-step 512, and one satellite of each stack can be connected to a mounting point of the ring structure. The horizontal stack of satellites and / or a horizontal axis of each stack can extend radially outward from the ring.

[0087] Step 518 includes transporting the plurality of satellites into space in the launch vehicle. Step 518 and / or the method 500 can include completing launch preparations for the launch vehicle and / or the satellites. For example, the method can include connecting control systems and separation systems, enclosing the satellites in thermal protection, and / or loading other payloads. Step 518 can include launching the launch vehicle, and propelling the launch vehicle into space by a rocket.

[0088] Space can be understood to include any desired region or location for deployment of one or more of the plurality of satellites being transported. For example, space can include, but is not limited to, a region beyond the Karman line of the Earth, a region beyond the atmosphere of a planetary body, or an orbit around a non-planetary body.

[0089] Step 520 includes separating the satellite from the launch vehicle perpendicular to the launch vehicle axis. Separating the satellites can be performed sequentially and can be performed by actuating the separation systems that connect adjacent satellites and the separation systems that connect the satellite stack to the ring structure of the launch vehicle in turn.

[0090] The separation systems can be disposed between adjacent satellites and / or otherwise configured to provide a separation impulse in a direction parallel to the axis along which the satellites are stacked. In other words, each separation system can be actuated to push the satellite away from the launch vehicle in a direction perpendicular to the main axis of the launch vehicle and / or in a direction radially outward from the ring structure.

[0091] Further illustrative and non-exclusive examples in accordance with the present disclosure as described in the following paragraphs include implementations in accordance with the following clauses:

[0092] In examples in accordance with the present disclosure, the first satellite apparatus (100, 212) includes a housing (234) including first and second walls (236, 238) opposite one another and a support structure (260) spanning the first and second walls opposite one another and enclosed by the housing, wherein a proximal portion (262) of the support structure is configured to be connected to the launch vehicle (124) by the separation system (106, 410).

[0093] Optionally, in the apparatus of the preceding paragraph, the support structure (260) is a hollow column.

[0094] Optionally, in the apparatus of one of the preceding paragraphs, the support structure (260) is cylindrical.

[0095] Optionally, in the apparatus of one of the preceding paragraphs, the support structure (260) is composed of a laser sintered metal alloy.

[0096] Optionally, in the apparatus of one of the preceding paragraphs, the support structure (260) includes a wall portion (266) having diamond-shaped apertures (268).

[0097] Optionally, in the apparatus of one of the preceding paragraphs, a distal portion (264) of the support structure (260) is configured to be connected to another satellite (214) by the separation system (106, 412).

[0098] Optionally, the apparatus of one of the preceding paragraphs further comprises a second satellite apparatus (214) comprising a housing (334) comprising first and second walls (336, 338) opposite one another and a support structure (360) spanning the first and second walls opposite one another and enclosed by the housing, wherein the distal portion (264) of the support structure in the first satellite (212) is connected to the proximal portion (362) of the support structure in the second satellite.

[0099] Optionally, in the apparatus of one of the preceding paragraphs, the load is transferred from the second satellite (214) through the support structure (260) of the first satellite (212) to the launch vehicle (124).

[0100] Optionally, in the apparatus of one of the preceding paragraphs, the support structures (260, 360) of the first and second satellite apparatuses (212, 214) share a common central axis (226) when connected.

[0101] Optionally, in the apparatus of one of the preceding paragraphs, the support structure (260) of the first satellite (212) is connected to the support structure (360) of the second satellite (214) by a separation system (106, 412).

[0102] In another example in accordance with the present disclosure, a satellite assembly (200) comprises a launch vehicle (124) having a launch axis (222) and a plurality of satellites (212, 214) stacked inside the launch vehicle perpendicular to the launch axis.

[0103] Optionally, in the apparatus of the preceding paragraph, each of the plurality of satellites (212, 214) has an internal cylindrical core structure (260, 360) aligned along a common core axis (226).

[0104] Optionally, in the apparatus of one of the preceding paragraphs, the cylindrical core structure (260, 360) is additively manufactured.

[0105] Optionally, in the apparatus of one of the preceding paragraphs, the cylindrical core structure (260, 360) has a wall portion (266, 366) comprising diamond-shaped apertures (268, 368).

[0106] Optionally, in the apparatus of one of the preceding paragraphs, the plurality of satellites (212, 214) are fixed in a plurality of satellite stacks (210) equally distributed around a circumference of a ring structure (217).

[0107] Optionally, in the apparatus of one of the preceding paragraphs, the ring structure (217) has a center point (228), each of the satellites (212, 214) includes a cylindrical core structure (260, 360) aligned along a core axis (226) that passes through the center point of the ring structure.

[0108] Optionally, in the apparatus of one of the preceding paragraphs, each satellite (212, 214) is connected to an adjacent stacked satellite (212, 214) by a separation system (106, 412).

[0109] In another example in accordance with the present disclosure, a method (500) of deploying satellites from a launch vehicle includes a loading step (510) of loading a plurality of satellites inside a launch vehicle by a stacking step (512) of horizontally stacking the plurality of satellites relative to a vertical launch axis, a transporting step (518) of transporting the satellites to space on the launch vehicle, and a separation step (520) of separating the satellites from the launch vehicle relative to the vertical launch axis.

[0110] Optionally, in the method of the preceding paragraph, each satellite has a cylindrical core structure, the loading step (510) includes a connecting step (514) of connecting the cylindrical core structures of adjacent satellites.

[0111] Optionally, in the method of one of the preceding paragraphs, the stacking step (512) includes:

[0112] a connecting step (514) of connecting the cylindrical core structures of adjacent satellites.

[0113] Different examples of satellites and satellite assemblies described herein provide several advantages over known solutions for designing and installing satellites for launch. For example, illustrative examples described herein allow for a robust and simple satellite structure design.

[0114] In addition, illustrative examples described herein provide a rigid and lightweight main structure, among other benefits.

[0115] In addition, illustrative examples described herein allow for rapid and inexpensive production of satellite main structures by additive manufacturing, among other benefits.

[0116] In addition, illustrative examples described herein allow for a robust and simple load path for stacking satellites, among other benefits.

[0117] In addition, illustrative examples described herein remove main structure functional limitations from the satellite housing, among other benefits.

[0118] No known system or device can perform these functions, particularly in a horizontal configuration. Thus, the illustrative examples described herein are particularly useful for efficient use of secondary payload space in launch vehicles. However, not all of the examples described herein provide the same advantages or the same degree of advantages.

[0119] The disclosure set forth above can encompass a variety of different examples with independent utility. While each of these has been disclosed in its preferred form, the particular examples disclosed herein are not to be construed as limiting as many variations are possible. With respect to section headings used in the present disclosure, these have been included for organizational purposes only. The subject matter of the present disclosure includes all novel and nonobvious combinations and subcombinations of the various elements, features, functions and / or properties disclosed herein. The following claims particularly point out certain combinations and subcombinations that are regarded as novel and nonobvious. Other combinations and subcombinations can be claimed by amending the following claims accordingly. Such amended claims, whether they designate a single combination as measured by the number of dependencies, or a multiple combination as measured by the number of dependencies in a single claim (e.g., a dependent claim referencing a multiple dependent claim; and claims containing expressing an amount of support for a claim), are to be construed in accordance with 35 U.S.C. 112, Sixth Paragraph.

Claims

1. A satellite assembly (200) comprising a launch vehicle (124) having a launch axis (222); a separation system; and first satellite equipment (100, 212); And the second satellite equipment (214), of which: The first satellite device (100, 212) includes: The first housing (234) includes a first wall (236) and a second wall (238) opposite to each other, and A first support structure (260) spans across the first and second walls, which are opposite to each other, and is surrounded by the first shell. The proximal portion (262) of the first support structure is configured to be connected to the launch vehicle (124) via the separation system (106, 410). The second satellite device (214) includes: The second housing (334) includes a third wall (336) and a fourth wall (338) opposite to each other, and The second support structure (360) spans the opposing third and fourth walls of the second housing (334) and is surrounded by the second housing (334). The distal portion (264) of the first support structure (260) in the first satellite device (212) is connected to the proximal portion (362) of the second support structure (360) in the second satellite device (214). The first satellite device (212) and the second satellite device (214) are located inside the launch vehicle and stacked perpendicular to the launch axis.

2. The satellite component (200) according to claim 1, wherein, The first support structure (260) includes a wall portion (266) having a rhomboid opening (268).

3. The satellite assembly (200) according to claim 1 or 2, wherein, The payload is transferred from the second satellite device (214) to the launch vehicle (124) through the first support structure (260) of the first satellite device (212).

4. The satellite assembly (200) according to claim 1 or 2, wherein, The first support structure (260) of the first satellite device (212) and the second support structure (360) of the second satellite device (214) share a common central axis (226) when connected.

5. The satellite assembly (200) according to claim 1 or 2, wherein, The first support structure (260) of the first satellite device (212) is connected to the second support structure (360) of the second satellite device (214) via the separation system.

6. The satellite assembly (200) according to claim 1, wherein, The first support structure (260) is a hollow column.

7. The satellite assembly (200) according to claim 6, wherein, The first support structure (260) is cylindrical.

8. The satellite assembly (200) according to claim 1, wherein, The first support structure (260) is made of laser-sintered metal alloy.

9. The satellite assembly (200) according to claim 1, wherein, The second support structure (360) is cylindrical.

10. The satellite assembly (200) according to claim 9, wherein, The second support structure (360) has a wall (366) including a diamond-shaped opening (368).

11. The satellite assembly (200) according to claim 1, further comprising one or more other satellite devices forming a plurality of satellite devices with the first satellite device (212) and the second satellite device (214), wherein, The plurality of satellite devices are fixed in a plurality of satellite stacks (210) that are equally spaced around a ring structure (217), wherein the ring structure (217) has a center point (228), and each of the plurality of satellite devices includes a cylindrical core structure aligned along a core axis (226) passing through the center point of the ring structure.

12. The satellite assembly (200) according to claim 11, wherein, Each of the plurality of satellite devices is connected to adjacent stacked satellite devices via the separation system.

13. The satellite assembly (200) according to claim 11, wherein, The cylindrical core structure is manufactured using additive manufacturing.

14. A method (500) for deploying satellite equipment from a launch vehicle in a satellite assembly (200) as described in any one of claims 1-13, comprising: The loading step (510) involves loading the first satellite device and the second satellite device inside the launch vehicle by stacking the first satellite device and the second satellite device horizontally relative to the vertical launch axis in a stacking step (512). The transport step (518) of transporting the first satellite equipment and the second satellite equipment into space in the launch vehicle, and Separation step (520) in which the first satellite device and the second satellite device are horizontally separated from the launch vehicle relative to the vertical launch axis.

15. The method according to claim 14, wherein, The loading step (510) includes: The connection step (514) of the first support structure and the second support structure connecting the adjacent first satellite device and the second satellite device.

16. The method according to claim 14 or 15, wherein, The stacking step (512) includes: Assembly step (516) of assembling a stack of multiple satellite devices evenly distributed around a ring structure (217).

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