Patent Description:
<CIT>, in accordance with its abstract, states a three-axis spacecraft including a spacecraft body including first and second opposing radiator/equipment panels, first and second opposing mounting panels, an earth deck and a zenith deck. The zenith deck faces the Earth when the spacecraft is on orbit and the first and second mounting panels face an east and west direction relative to the Earth when the spacecraft is on orbit. The spacecraft further includes a mounting cylinder extending through the spacecraft body and out of the first and second mounting panels.

<CIT>, in accordance with its abstract, states a spacecraft may include an upper core structure or a lower core structure. The upper core structure may include an upper cylinder for supporting an upper spacecraft of a dual-manifest launch configuration. The lower core structure may include a lower cylinder for supporting a lower spacecraft with the upper cylinder mounted on top of the lower cylinder. The upper cylinder may have an upper cylinder inner diameter that may be substantially similar to the lower cylinder inner diameter.

<CIT>, in accordance with its abstract, states a multi-star distributor force bearing structure which comprises a force bearing cylinder with a grid structure. The grid structure is a skin-free grid structure. A first inner reinforced n-shaped frame is arranged at a first end of the force bearing cylinder, and a second inner reinforced n-shaped frame is arranged at a second end of the force bearing cylinder; and at least one first satellite supporting table is arranged on an outer side arc surface of the first end of the force bearing cylinder, and at least one second satellite supporting table is arranged on the outer side arc surface of the second end of the force bearing cylinder.

<CIT>, in accordance with its abstract, states techniques for deploying a plurality of smallsats from a common launch vehicle where a structural arrangement provides a load path between an upper stage of the launch and the plurality of spacecraft. Each spacecraft is mechanically coupled with the launch vehicle upper stage only by the structural arrangement. The structural arrangement includes at least one trunk member that is approximately aligned with the longitudinal axis of the launch vehicle upper stage, a plurality of branch members, each branch member being attached to the trunk member and having at least a first end portion that is substantially outboard from the longitudinal axis; and a plurality of mechanical linkages, each linkage coupled at a first end with a first respective spacecraft and coupled at a second end with one of the plurality of branch members, the trunk member or a second respective spacecraft.

<CIT>, in accordance with its abstract, states a modular device for a spacecraft includes a propulsion system having a tank, a plenum, and a manifold, wherein the propulsion system is integrally formed with a structural frame of the spacecraft. A method of manufacturing the modular device is also discussed, the method including utilizing an additive manufacturing process to construct the propulsion system.

The present disclosure provides systems, apparatus, and methods relating to satellite support structures and assemblies.

A satellite assembly comprises: a launch vehicle having a launch axis; a first separation system; a first satellite apparatus; and a second satellite apparatus, wherein: the first satellite apparatus, comprises: a first housing including first and second opposing walls; and a first support structure (<NUM>) spanning the first and second opposing walls and enclosed by the first housing; an end portion of the first support structure is configured for connection to the launch vehicle by the first separation system; the second satellite apparatus comprises: a second housing including first and second opposing walls, and a second support structure spanning the first and second opposing walls of the second housing and enclosed by the second housing; a distal end portion of the first support structure in the first satellite apparatus is connected to a proximal end portion of the second support structure in the second satellite apparatus; wherein the first and second satellite apparatus are stacked perpendicular to the launch axis inside the launch vehicle.

In some examples, a method of deploying satellites from a launch vehicle may include stowing a plurality of satellites inside a launch vehicle by stacking the satellites horizontally relative to a vertical launch axis. The method may further include carrying the satellites to space in the launch vehicle, and separating the satellites from the launch vehicle horizontally relative to the vertical launch axis.

Features, functions, and advantages may be achieved independently in various examples of the present disclosure, or may be combined in yet other examples, further details of which can be seen with reference to the following description and drawings. 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 in nature and not all examples provide the same advantages or the same degree of advantages.

This Detailed Description includes the following sections, which follow immediately below: (<NUM>) Overview; (<NUM>) Examples, Components, and Alternatives; (<NUM>) Illustrative Combinations and Additional Examples; (<NUM>) Advantages, Features, and Benefits; and (<NUM>) Conclusion. The Examples, Components, and Alternatives section is further divided into subsections A through C, each of which is labelled accordingly.

In general, a satellite in accordance with the present teachings may include a hollow central support structure and a housing. The central support structure may be the primary structure of the satellite, supporting the housing and connecting to a launch vehicle. The central support structure may span between first and second panels of the housing, and may be connected to the housing only by the first and second panels. Payload and operational equipment of the satellite may be supported by the housing. The central support structure may be additively manufactured, and include a cylindrical wall having an array of diamond-shaped apertures. For launch, the satellite may form part of a structural satellite launch configuration.

In general, a structural satellite launch configuration in accordance with the present teachings includes two satellites, each satellite having a central support structure that is the primary structure of the satellite. The central support structures of the two satellites may be connected to form a single beam structure, which may be mounted as a cantilever beam to a payload adaptor of a launch vehicle. In other words, the central support structures of the two satellites may define a core axis and the satellites may be mounted in the launch vehicle such that the core axis is perpendicular to a launch axis of the launch vehicle.

In some examples, a structural satellite launch configuration may include a plurality a pairs of connected satellites. Each pair of connected satellites may be connected to a central ring payload adaptor of the launch vehicle, extending radially outward from the ring adaptor. In some examples, a structural satellite launch configuration may include one or more stacks of three or more satellites having connected central support structures, the one or more stacks being mounted to the launch vehicle such that a core axis defined by the central support structures is perpendicular to the launch axis of the launch vehicle.

The following sections describe selected exemplary satellites as well as related assemblies and/or methods. The examples in these sections are intended for illustration and should not be interpreted as limiting the entire scope of the present disclosure. Each section may include one or more distinct examples, and/or contextual or related information, function, and/or structure.

Examples disclosed herein may be described in the context of an illustrative satellite launch method <NUM> (see <FIG>) and an illustrative satellite <NUM> (see <FIG>). In some examples including the present example, method <NUM> includes three phases: a launch phase <NUM>, a separation or deployment phase <NUM>, and an operation phase <NUM>. Launch phase <NUM> may include transporting satellite <NUM> (alternatively, spacecraft <NUM>) from a planetary body <NUM> such as Earth to outer space <NUM>, using a launch vehicle <NUM>. In the context of Earth, outer space may comprise a region beyond the Karman line. Deployment phase <NUM> may include separating satellite <NUM> from launch vehicle <NUM> once a desired location, trajectory and/or orbit has been achieved. Operation phase <NUM> may include preparation of satellite <NUM> for operation, such as establishing communication with a controller on planetary body <NUM>, extending solar panels or instrument arms, and/or maneuvering to a desired orientation relative to the planetary body. In some examples, the method may further include design, production, and/or in-service phases.

For the purposes of this description, a system integrator may include, without limitation, any number of aerospace manufacturers and major-system subcontractors; a third party may include, without limitation, any number of vendors, subcontractors, and suppliers; and an operator may be a telecommunications company, leasing company, military entity, service organization, and so on.

As shown in <FIG>, satellite <NUM> may include a bus <NUM> with a plurality of satellite systems, a payload <NUM> and a separation system <NUM>. Examples of the plurality of systems include one or more of a primary structure <NUM>, a propulsion system <NUM>, an electrical power system <NUM>, a thermal management system <NUM>, a radiation shielding system <NUM>, and a communication system <NUM>. Each system may comprise various subsystems, such as controllers, processors, actuators, effectors, motors, generators, etc., depending on the functionality involved. Any number of other systems may be included. Although an unmanned artificial satellite example is shown, the principles disclosed herein may be applied to other aerospace vehicles and technology, such as a launch vehicle, space station, crewed spacecraft, and/or interstellar probe.

Apparatuses and methods shown or described herein may be employed during any one or more of the stages of the satellite launch method <NUM>. Two or more satellites are stacked perpendicular to a launch axis of launch vehicle <NUM> during launch phase <NUM>. Similarly, one or more examples of the apparatus or method realizations, or a combination thereof, may be utilized, for example and without limitation, while satellite <NUM> and/or launch vehicle <NUM> are in preparation prior to execution of launch method <NUM>. Also, one or more examples of the apparatuses, methods, or combinations thereof may be utilized during deployment phase <NUM> for example, by deploying a satellite radially outward from launch vehicle <NUM>, perpendicular to the launch axis of the vehicle.

As shown in <FIG>, this section describes an illustrative satellite assembly <NUM>. Satellite assembly <NUM> is an example of a structural satellite launch configuration, as described above. The assembly includes a plurality of satellite stacks <NUM>. Each stack <NUM> includes a proximal satellite <NUM> and a distal satellite <NUM>, and is connected to a payload adaptor <NUM> by a mounting plate <NUM>. In some examples, a stack may include three or more satellites.

In some examples including the example depicted in <FIG>, payload adaptor <NUM> includes a ring structure <NUM> such as the Evolved Secondary Payload Adapter (ESPA) produced by Moog, Inc. The plurality of satellite stacks <NUM> are connected to ring structure <NUM> at six mounting points <NUM>, arranged symmetrically around the ring structure. In <FIG>, one of satellite stacks <NUM> is not depicted, in order to show the corresponding mounting point <NUM>. In general, the plurality of satellite stacks may be arranged symmetrically about payload adaptor <NUM> in order to balance loads transferred to the payload adaptor.

Payload adaptor <NUM> is part of a launch vehicle, having a launch axis <NUM>. The launch axis may also be described as a longitudinal axis of the launch vehicle, as a z-axis, or as a vertical axis. Directions perpendicular to the launch axis may be described as lateral and/or horizontal.

Prior to launch, the launch axis may be aligned with a vertical direction as defined by a gravitational frame of reference. During launch, the launch axis may rotate relative to the gravitational frame of reference as the vehicle follows a non-linear launch trajectory. Therefore, for clarity in the following description, directional terms and descriptors such as "up", "down", "top", "bottom", and the like should be understood relative to the launch axis.

In some examples including the present example, ring structure <NUM> of payload adaptor <NUM> has a central axis <NUM> parallel to launch axis <NUM>. Each of satellite stacks <NUM> has a core axis <NUM>, which may also be described as a longitudinal or central axis of the stack. Core axis <NUM> of each satellite stack <NUM> extends through a center point <NUM> of ring structure <NUM>, on central axis <NUM> of the ring structure. That is, the core axes of the plurality of satellite stacks intersect at the center point of the ring structure.

The plurality of satellite stacks <NUM> may be described as horizontal stacks, branches, projection assemblies, and/or radially connected satellite groups. Each satellite stack <NUM> extends radially out from ring structure <NUM>, perpendicular to central axis <NUM> of the ring structure. That is, core axis <NUM> of each satellite stack is perpendicular to launch axis <NUM>.

Proximal satellite <NUM> of each satellite stack is releasably connected to the corresponding mounting plate <NUM> by a separation system and/or device as discussed further below. Each distal satellite <NUM> is similarly releasably connected to the corresponding proximal satellite by a separation system and/or device. Each mounting plate <NUM> is fixedly attached to one of mounting points <NUM> of ring structure <NUM>. In some examples, including the present example, the mounting plate is bolted to the ring structure. In some examples, the mounting plate may be an integral part of payload adaptor <NUM> and/or the proximal satellite may connect directly to the mounting point. In some examples, mounting plate <NUM> may support other additional payload or launch vehicle components, and/or may form part of another structure.

<FIG> depict one satellite stack <NUM>, and <FIG> depict proximal satellite <NUM> of that satellite stack. Descriptions thereof may be understood to apply equally to each of satellite stacks <NUM>, except where stated otherwise. In general, satellites of a satellite assembly as described herein may include a primary structure as described below, but may vary in payload, housing design and specifications of operational systems such as communications, shielding, and thermal regulation.

<FIG> and <FIG> are opposite isometric views of proximal satellite <NUM>. <FIG> depicts a distal side of the satellite and <FIG> a proximal side of the satellite, in the context of the satellite's orientation relative to the payload adaptor. In some examples including the present example, proximal satellite <NUM> is roughly cuboid, and includes a housing <NUM> of six planar and square or rectangular wall panels. More specifically, the satellite includes a fore panel <NUM> and an opposing aft panel <NUM>. Four equipment panels <NUM> span between the fore and aft panels.

Proximal satellite <NUM> further includes a plurality of patch antennas <NUM>, mast baffles <NUM>, and a solar array <NUM> comprised of two deployable panels. Four thrusters <NUM> are mounted in brackets <NUM> at the four corners of aft panel <NUM>. Satellite <NUM> may further include any appropriate operational or payload equipment, including but not limited to a fuel tank, star tracker, reaction wheel, heat sinks, radiator panels, and/or avionics. A majority of equipment may be mounted to interior surfaces of equipment panels <NUM>.

Also shown in <FIG> is a proximal end portion <NUM> of a cylindrical core structure <NUM> of proximal satellite <NUM>, which extends out through aft panel <NUM>. Core structure <NUM> can be seen more entirely in <FIG>, where housing <NUM> is depicted as transparent. The core structure may also be described as a support structure, hollow column, and/or central beam.

Core structure <NUM> is a hollow cylinder, spanning between aft panel <NUM> and fore panel <NUM> and enclosed in housing <NUM>. The core structure may also be described as a hollow column. The core structure defines a central axis <NUM> and is centered in proximal satellite <NUM>. Proximal end portion <NUM> is fixed to aft panel <NUM> and a distal end portion <NUM> of the core structure is fixed to fore panel <NUM>. A wall <NUM> including a plurality of apertures <NUM> extends between the two end portions. Wall <NUM> is thin relative to the diameter of the core structure, allowing the core structure to be strong and stiff but light. Apertures <NUM> may further lighten the core structure, without sacrificing desired structural properties.

Core structure <NUM> acts at the primary structure of the satellite, and is configured to structurally connect the satellite to both the launch vehicle and a distal satellite. More specifically, proximal end portion <NUM> is configured for connection to the launch vehicle payload adaptor through a mounting plate and distal end portion <NUM> is configured for connection to the core structure of another satellite. Each end portion of core structure <NUM> is configured for connection by a separation system or device. In some examples including the present example, both end portions are configured for connection by similar separation systems, as described further below. In some examples, the proximal and distal end portions may be configured for connection by different separation system or devices.

Distal end portion <NUM> includes an interface flange <NUM> which contacts an interior surface of fore panel <NUM>. In some examples including the present example, the core structure is fixed to the fore panel by a plurality of fasteners extending through apertures in interface flange <NUM> and into the fore panel. Proximal end portion <NUM> includes a plurality of interface tabs around the circumference of wall <NUM>, extending out from the wall and contacting an exterior surface of aft panel <NUM>. In some examples including the present example, the core structure is fixed to the aft panel by a plurality of fasteners extending through apertures in interface tabs <NUM> into the aft panel.

As shown in <FIG>, proximal satellite <NUM> includes a fuel tank <NUM> mounted inside core structure <NUM>. Such placement may help to maximize space efficiency in the satellite, allowing core structure <NUM> to have a large cross-section for improved strength and stiffness, without wasting interior space. Apertures <NUM> may facilitate necessary access and/or connections to the tank such as fluid connections to fill valves or electrical connections to sensors.

An outer diameter of tank <NUM> may be close to, but less than an inner diameter of wall <NUM> of core structure <NUM>. To allow the closely-fitting tank to be positioned inside core structure <NUM>, the core structure includes two parts which can be assembled around tank <NUM>. More specifically, core structure <NUM> comprises a first section <NUM> and a second section <NUM>.

First section <NUM> includes proximal portion <NUM>, second section <NUM> includes distal portion <NUM>, and the first and second sections are bolted together at an intermediate interface <NUM>. Intermediate interface <NUM> may be described as disposed part-way along the extent of wall <NUM>. Each of the first and second sections are thickened proximate intermediate interface <NUM>, to reinforce and strengthen the connection and allow the two sections to act as a single effective structural support of satellite <NUM>.

In some examples including the present example, wall <NUM> further includes apertures for a plurality of shim bolts <NUM> to center and precisely position tank <NUM> in the core structure. In general, core structure <NUM> may include any customizations or modifications appropriate to installation, support, or integration of operational equipment of proximal satellite <NUM>.

<FIG> is an isometric view of housing <NUM> of proximal satellite <NUM>, with panels <NUM>, <NUM>, <NUM> depicted as transparent. In some examples including the present example, the panels of the housing are connected by eight corner brackets <NUM> and four angle clips <NUM>. Corner brackets <NUM> include the four thruster brackets <NUM>, and are each positioned inside a corner of housing <NUM>, where a corner of either fore panel <NUM> or aft panel <NUM> meets corners of two equipment panels <NUM>. Angle clips <NUM> each extend along the inside of an edge of housing <NUM>, where edges of two equipment panels <NUM> meet.

A main body <NUM> of each corner bracket <NUM> is positioned at a corresponding corner cut-out of fore panel <NUM> or aft panel <NUM>. Sides of main portion <NUM> of the corner bracket contact interior surfaces of the three adjacent panels, and may be bonded or otherwise fixed to the panels. Corner bracket <NUM> further includes angle tabs <NUM>, configured to contact an inner surface of angle clips <NUM>. For each corner bracket <NUM>, one angle tab <NUM> may be bonded or otherwise fixed to an adjacent angle clip <NUM>. Each angle clip <NUM> may therefore be fixed to, and extend between, a first corner bracket at aft panel <NUM> and a second corner bracket at fore panel <NUM>.

Angle clips <NUM> and corner brackets <NUM> may structurally connect equipment panels <NUM> to fore panel <NUM> and aft panel <NUM>, which in turn are structurally connected to the core structure of the satellite. Equipment panels <NUM> are directly connected to the core structure. In other words, the core structure is only connected to equipment panels <NUM> through the fore and aft panels. Loads from equipment mounted to equipment panels <NUM> may be transferred through fore panel <NUM> and aft panel <NUM> to the core structure.

Housing <NUM> does not form part of the primary structure of proximal satellite <NUM>. As shown in <FIG>, core structure <NUM> of proximal satellite <NUM> connects directly to mounting plate <NUM> and the core structure of the distal satellite. Core structure <NUM> is the primary structure of proximal satellite <NUM>, and the primary load path to the launch vehicle. Housing <NUM> surrounds and encloses the core structure and satellite equipment. The housing is supported by core structure <NUM>, and not directly connected to the launch vehicle. As a result, freedom of material choice and design for housing <NUM> is significantly increased.

As shown in <FIG> and described above, in some examples including the present example housing <NUM> is roughly cuboid and composed of planar panels. In general, the housing may have any shape appropriate to house satellite systems and equipment and efficiently stow inside the launch vehicle. For example, the housing may be a polyhedron, may include curved panels, and/or may have an irregular shape. Preferably, housing <NUM> may be symmetric or approximately symmetric about central axis <NUM> and/or balanced for straightforward and tumble-free separation during deployment.

In some examples including the present example, panels <NUM>, <NUM>, <NUM> comprise a composite honeycomb sandwich material. In general, the panels may include any light-weight material or materials that are sufficiently strong to support mounted equipment. For example, the panels may be additively manufactured and/or include additively manufactured portions that may be produced by three-dimensional (3D) printing, laser sintering of a metal alloy, or other method. The panels need not be designed for the strength or stiffness required of a primary structure.

Housing <NUM> may be highly customizable to selected payload and operational satellite equipment. View ports, supports, shielding, access holes, or other modifications may be made to the housing without affecting the primary structure of the satellite. Particularly in combination with the rapid prototyping and design implementation capabilities of additively manufactured components, such freedom may significantly simplify design and reduce testing and certification times.

Of housing <NUM>, fore panel <NUM> and aft panel <NUM> may have the most design constraints. That is, the two panels or structures of housing <NUM> connected to the proximal and distal ends of the core structure may need to be configured to interface with the core structure. In some examples including the present example, the core structure interfaces with the fore and aft panels as defined by the position and direction of the satellite thrusters. In some examples, the core structure may connect to sides of the satellite such that the satellite may be described as mounted sideways to the launch vehicle, or may be mounted to any two opposing walls or wall portions of the housing.

Referring again to <FIG>, aft panel <NUM> includes a circular aperture <NUM> with six circumferential cutouts or recesses <NUM>. Circular aperture <NUM> and recesses <NUM> may allow core structure <NUM> to protrude through aft panel <NUM>, such that proximal end portion <NUM> is exterior to the aft panel as shown in <FIG>. More specifically, wall <NUM> may extend through circular aperture <NUM> and six bays <NUM> protruding out from wall <NUM> may extend through recesses <NUM>. In other words, aft panel <NUM> may include an aperture shaped to correspond to core structure <NUM>, such that proximal end portion <NUM> can extend through the aft panel. The specific shape of proximal end portion <NUM> and the aperture in aft panel <NUM> may depend on the separation system selected, as described further below.

As shown in <FIG> and <FIG>, fore panel <NUM> includes six large and three small circular apertures <NUM>, positioned to allow devices of the separation system to extend through the panel to engage distal end portion <NUM> of core structure <NUM>. Similarly to aft panel <NUM>, fore panel <NUM> may include any aperture or apertures shaped and positioned to match the selected separation system and corresponding configuration of distal end portion <NUM>.

<FIG> depicts proximal satellite <NUM> connected to distal satellite <NUM> and mounting plate <NUM>, as part of a satellite stack <NUM>. Similarly to proximal satellite <NUM>, distal satellite <NUM> includes a roughly cuboid housing <NUM> and a cylindrical core structure <NUM> with a proximal end portion <NUM> and a distal end portion <NUM>. Core structure <NUM> spans between a fore panel <NUM> and an aft panel <NUM>, with proximal end portion <NUM> extending out through the aft panel to connect to distal end portion <NUM> of proximal satellite <NUM>.

In some examples including the present example, distal satellite <NUM> is identical or substantially identical to proximal satellite <NUM> apart from the configuration of distal end portion <NUM> of core structure <NUM>. Accordingly, reference numerals for components of distal satellite <NUM> match those of corresponding components of proximal satellite <NUM>. In general, distal satellite <NUM> may include a core structure <NUM> matching or substantially matching and configured to connect to core structure <NUM> of proximal satellite <NUM>, but may otherwise differ in design from proximal satellite <NUM>. For example, payloads, operational equipment, and/or housings of the two satellites may differ.

Distal end portion <NUM> of distal satellite <NUM> is more simply configured than distal end portion <NUM> of proximal satellite <NUM>, as shown more clearly in <FIG>. Unlike distal end portion <NUM>, distal end portion <NUM> does not need to be configured for connection to another core structure, in some examples including the present example. Therefore distal end portion <NUM> includes cylindrical wall <NUM> up to a circular, annular interface flange <NUM> for connection to the fore panel of distal satellite <NUM>. This simpler shape may be desirably lighter. In some examples, distal end portion <NUM> of distal satellite <NUM> may match distal end portion <NUM> of proximal satellite <NUM> for simplicity of manufacture and/or satellite design. In some examples, satellite stack <NUM> may include three or more satellites, at least one of which may include a core configured for connection to an adjacent satellite at both proximal and distal ends.

In some examples including the present example, distal satellite <NUM> connects to proximal satellite <NUM> in the same manner as the proximal satellite connects to mounting plate <NUM>. Accordingly, proximal end portion <NUM> of core structure <NUM> of distal satellite <NUM> matches proximal end portion <NUM> of core structure <NUM> of proximal satellite <NUM>. Mounting plate <NUM> also includes a distal portion <NUM> which matches distal portion <NUM> of proximal satellite <NUM>.

Mounting plate <NUM> may act as an adaptor, facilitating structural connection between core structure <NUM> of proximal satellite <NUM> and the launch vehicle payload adaptor. The mounting plate includes a proximal portion <NUM> configured for connection to a mounting point of the payload adaptor. In some examples including the present example, proximal portion <NUM> includes a square, planar face with bolt holes at each corner. Proximal portion <NUM> and distal portion <NUM> are joined by a cylindrical center wall with supporting braces. In general, mounting plate <NUM> may have any geometry or configuration appropriate to provide a strong connection and efficient load path between the core structures of the satellites and the launch vehicle.

Together, mounting plate <NUM>, core structure <NUM>, and core structure <NUM> may act as a cantilever beam extending horizontally outward from the launch vehicle payload adaptor. The combined core structure is sufficiently stiff to support both proximal satellite <NUM> and distal satellite <NUM>, withstanding the bending moment and vibrational loading associated with launch. The joined core structures also provide a strong and simple load path to the launch vehicle. As described further below, the dimensions and design of the core structures provide the needed stiffness, with minimal weight.

Also important to the stiffness of the beam effected by the combined core structures, is the stiffness of the connections between core structure <NUM> and core structure <NUM>, and between core structure <NUM> and mounting plate <NUM>. Any effective separation system or device may be used to connect the satellites. However, a system providing direct connection between the core structures, such as is depicted in the present example, may be preferable to provide a sufficiently stiff connection.

Satellite stack <NUM> includes a proximal separation system <NUM> connecting proximal satellite <NUM> and mounting plate <NUM>, and a distal separation system <NUM> connecting distal satellite <NUM> and proximal satellite <NUM>. As shown most clearly in <FIG>, each separation system <NUM>, <NUM> includes six separable connectors <NUM> and three push-off pins <NUM>. Each separable connector <NUM> comprises a male portion <NUM> and a female portion <NUM>.

On proximal satellite <NUM>, male portions <NUM> of separable connectors <NUM> of separation system <NUM> are housed in bays <NUM> of proximal end portion <NUM> and extend out through an interface flange <NUM> of the proximal end portion to engage a corresponding female portion on the mounting plate. Female portions <NUM> of separable connectors <NUM> of separation system <NUM> are mounted in recesses in wall <NUM> of distal end portion <NUM>, on an opposite side of interface flange <NUM> from fore panel <NUM>. Corresponding apertures in interface flange <NUM> allow the respective male portions on the distal satellite to reach through the flange to female portions <NUM>.

Push-off pins <NUM> of separation system <NUM> are mounted similarly to female portions <NUM> of the separation system. That is, the push-off pins are mounted on the opposite side of interface flange <NUM> from fore panel <NUM>, and extend through corresponding apertures in the interface flange and fore panel to contact an interface flange of the proximal portion of distal satellite <NUM>. Interface flange <NUM> of proximal end portion <NUM> of proximal satellite <NUM> includes three scallops <NUM> to engage the push-off pins on mounting plate <NUM>.

Separable connectors <NUM> and push-off pins <NUM> of separation systems <NUM>, <NUM> are spaced evenly around the circumference of core structures <NUM>, <NUM>. Each device <NUM>, <NUM> of the separation systems is spring actuated for smooth and reliable separation, and connected to a control system for coordinated triggering.

As shown in <FIG>, devices <NUM>, <NUM> of separation system <NUM> are not interposed between mounting plate <NUM> and core structure <NUM>. Similarly the devices of separation system <NUM> are not interposed between core structure <NUM> and core structure <NUM>. Interface flange <NUM> of proximal end portion <NUM> of core structure <NUM> contacts mounting plate <NUM> directly. Interface flange <NUM> of distal end portion <NUM> of core structure <NUM> and interface flange <NUM> of proximal end portion <NUM> of core structure <NUM> contact the inner and exterior faces of fore panel <NUM> of proximal satellite respectively, with only the fore panel between the two flanges. The direct connection between core structures may result in the desired stiffness. Devices of the separation systems may also be individually configured to facilitate a stiff connection.

In some examples including the present example, separation system <NUM> further includes two separable housing connectors <NUM>, shown in <FIG> and <FIG>. Each housing connector <NUM> includes a first bracket <NUM> mounted to fore panel <NUM> of proximal satellite <NUM> and a second bracket mounted to aft panel <NUM> of distal satellite <NUM>. The first and second brackets <NUM>, <NUM> are connected by a spring-actuated releasable mechanism similar to separable connectors <NUM>. Together the two connected brackets <NUM>, <NUM> have an axial extent matching proximal end portion <NUM> of the distal satellite, allowing housing connectors <NUM> to bridge between the two satellites when core structure <NUM> is connected to core structure <NUM>.

Housing connectors <NUM> may be configured and/or positioned according to the geometry or other properties of housings <NUM> and <NUM>. In some examples including the present example, the two housing connectors are positioned at opposing outer lateral edges of the satellites to provide additional lateral stability to the connection between the satellites, and assist in tumble-free separation.

<FIG> is a cross-sectional view of the primary structure of satellite stack <NUM>, including mounting plate <NUM>, separation system <NUM>, core structure <NUM>, separation system <NUM>, and core structure <NUM>. Each core structure has an inner diameter <NUM> as defined by an interior surface of wall <NUM> or <NUM>. Each core structure <NUM>, <NUM> also has a length <NUM> from the interface flange at the proximal end to the interface flange at the distal end. Inner diameter <NUM> and length <NUM> are the same for core structure <NUM> and core structure <NUM>.

In some examples, the core structures have an inner diameter <NUM> of approximately <NUM> inches (approximately <NUM>) and a length of approximately <NUM> inches (approximately <NUM>). In some examples including the present example, the core structures have an inner diameter <NUM> of <NUM> inches (<NUM>) and a length of <NUM> inches (<NUM>). Walls <NUM>, <NUM> of the core structures may have a thickness of between approximately <NUM> and <NUM> thousandths of an inch (between approximately <NUM> and <NUM>). Walls <NUM>, <NUM> of the core structures may have a thickness of between <NUM> and <NUM> thousandths of an inch (between <NUM> and <NUM>). The core structures, and the cantilever beam structure formed by connecting the core structures, are strong and stiff enough to support proximal and distal satellites of <NUM> kilograms (or approximately <NUM> kilograms) each, at a vibration frequency of <NUM> hertz (or approximately <NUM> hertz). In general, core structures <NUM>, <NUM> may be any size appropriate to a satellite's size and weight. That is, the core structure design may be applicable from microsats up through full-sized satellites.

Apertures <NUM> in wall <NUM> and apertures <NUM> in wall <NUM> may help to reduce the weight of the core structures, without sacrificing structural strength. In some examples including the present example, each core structure includes two arrays of apertures <NUM>, with a first array in the first section and a second array in the second section. In some examples, the apertures may be arranged in additional arrays, and in examples such as a unitary core structure the apertures may form a single array.

Each of apertures <NUM>, <NUM> is diamond shaped. In some examples including the present example, each aperture is approximately two inches (approximately <NUM>) in length and spaced approximately one quarter inch (approximately <NUM>) from adjacent apertures, or each aperture is two inches (<NUM>) in length and spaced one quarter inch (<NUM>) from adjacent apertures. Arrays of apertures <NUM> may also be described as a mesh and/or as a diamond lattice. The diamond shape may be particularly suited to additive manufacture. Any desired aperture shape may be used, and an appropriate aperture shape may depend on a selected method of manufacture. Aperture size and spacing may be selected according to desired structural properties and/or electromagnetic properties of the core structure.

In some examples including the present example, core structures <NUM>, <NUM> are additively manufactured from metal. More specifically, the core structures may comprise laser sintered metal alloy, manufactured using direct metal laser sintering (DMLS) of an aluminum alloy. In general, the core structures may be manufactured according to any effective method and of any sufficiently strong and light material. Additive manufacture of the core structures may be particularly suited to production of the thin walls, apertures, and customized interface features.

Other components of satellites <NUM>, <NUM> may be advantageously manufactured using additive manufacturing methods such as DMLS or electron beam melting (EBM). For example, fuel tank <NUM>, angle clips <NUM>, corner brackets <NUM>, and/or panels <NUM>, <NUM>, <NUM> may be additively manufactured.

This section describes steps of an illustrative method of deploying satellites from a launch vehicle; see <FIG>. Examples of satellites, structural satellite launch configurations, and/or launch vehicle payload adaptors and mounting plates described above may be utilized in the method steps described below. Where appropriate, reference may be made to components and systems that may be used in carrying out each step. These references are for illustration, and are not intended to limit the possible ways of carrying out any particular step of the method.

<FIG> is a flowchart illustrating steps performed in an illustrative method, and may not recite the complete process or all steps of the method. Although various steps of method <NUM> are described below and depicted in <FIG>, the steps need not necessarily all be performed, and in some cases may be performed simultaneously or in a different order than the order shown.

At step <NUM>, the method includes stowing a plurality of satellites in a launch vehicle. The launch vehicle may comprise any vehicle suitable to transport a payload to space. For example, the launch vehicle may be an expendable autonomous vehicle, or may be a manned spacecraft. Step <NUM> may be performed as part of preparations for launch of the vehicle, and the plurality of satellites may be configured for connection to and launch in the vehicle. Stowing the satellites may include attaching the satellites to a payload adaptor of the launch vehicle and/or to one another using one or more separation systems and/or devices. The satellites may be stowed according to substeps <NUM>-<NUM> of step <NUM>.

Sub-step <NUM> includes stacking multiple satellites horizontally relative to a vertical axis of the launch vehicle. In other words, two or more satellites may be positioned adjacent one another along a horizontal axis. The vertical axis may correspond to an orientation of the launch vehicle during preparations for launch and/or may correspond to a launch direction or launch axis. The vertical axis may also be referred to as a primary axis of the launch vehicle. The two or more satellites may be referred to as a horizontal stack and/or lateral assembly. Only one of the satellites of the stack may be directly connected to the launch vehicle.

Sub-step <NUM> of sub-step <NUM> includes connecting cylindrical core structures of adjacent satellites of the stacked satellites. A primary structure of each satellite of the plurality of satellites may include a cylindrical core structure. Each core structure may have the same diameter, and may be configured for connection to another core structure by a separation system. Within the horizontal stack of satellites, each satellite may be connected to the adjacent satellites by the core structure. The connected core structures of the satellites of the stack may form a beam, extending horizontally out from the launch vehicle payload adaptor.

Sub-step <NUM> of step <NUM> includes assembling and attaching plural stacks around a ring structure of the launch vehicle. The ring structure may be the payload adaptor of the launch vehicle, and may include a plurality of attachment or mount points. A plurality of horizontal stacks may be assembled according to sub-step <NUM>, and one satellite of each stack may be connected to a mount point of the ring structure. The horizontal stacks of satellites and/or the horizontal axis of each stack may extend radially outward from the ring.

Step <NUM> includes carrying the plurality of satellites to space in the launch vehicle. Step <NUM> and/or method <NUM> may include finalizing launch preparations for the vehicle and/or the satellites. For example, the method may include connecting control systems and separation systems, enclosing the satellites in thermal protection, and/or stowing additional payloads. Step <NUM> may include launching the vehicle, and propelling the vehicle into space with rockets.

Space may be understood to include any region or location desirable for deployment of one or more of the carried plurality of satellites. For example, space may include, but is not limited to, a region beyond the Karman line of Earth, a region outside the atmosphere of a planetary body, or an orbit around a non-planetary body.

Step <NUM> includes separating the satellites from the launch vehicle, perpendicular to the launch vehicle axis. Separating the satellites may be performed sequentially, and may be performed by actuating in turn the separation systems that connect adjacent satellites and the separation systems that connect the stacks of satellites to the ring structure of the launch vehicle.

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

The different examples of the satellites and satellite assemblies described herein provide several advantages over known solutions for designing and mounting satellites for launch. For example, illustrative examples described herein allow a sturdy and simple satellite structural design.

Additionally, and among other benefits, illustrative examples described herein provide a stiff and lightweight primary structure.

Additionally, and among other benefits, illustrative examples described herein allow a satellite primary structure to be rapidly and inexpensively produced by additive manufacture.

Additionally, and among other benefits, illustrative examples described herein allow a strong and simple load path for stacked satellites.

Additionally, and among other benefits, illustrative examples described herein remove the primary structural function limitations from the satellite housing.

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 utilization of secondary payload space in a launch vehicle. However, not all examples described herein provide the same advantages or the same degree of advantage.

Claim 1:
A satellite assembly (<NUM>), comprising: a launch vehicle (<NUM>) having a launch axis (<NUM>); a first separation system (<NUM>; <NUM>); a first satellite apparatus (<NUM>, <NUM>); and a second satellite apparatus (<NUM>), wherein:
the first satellite apparatus (<NUM>, <NUM>), comprises: a first housing (<NUM>) including first and second opposing walls (<NUM>,<NUM>); and a first support structure (<NUM>) spanning the first and second opposing walls (<NUM>, <NUM>) and enclosed by the first housing (<NUM>);
an end portion (<NUM>) of the first support structure (<NUM>) is connected to the launch vehicle (<NUM>) by the first separation system (<NUM>, <NUM>);
the second satellite apparatus (<NUM>) comprises: a second housing (<NUM>) including first and second opposing walls (<NUM>, <NUM>), and a second support structure (<NUM>) spanning the first and second opposing walls of the second housing (<NUM>) and enclosed by the second housing (<NUM>);
a distal end portion (<NUM>) of the first support structure (<NUM>) in the first satellite apparatus (<NUM>) is connected to a proximal end portion (<NUM>) of the second support structure (<NUM>) in the second satellite apparatus (<NUM>);
wherein the first and second satellite apparatus (<NUM>, <NUM>) are stacked perpendicular to the launch axis inside the launch vehicle.