Additively manufactured satellites

Abstract

Additively manufactured satellite. A satellite is disclosed, comprising a main body and a communication device attached to the main body. The main body has an additively manufactured outer wall structure that at least partially forms an enclosed compartment, and the communication device is configured to receive and transmit data in space.

Description

Technical Field

[0001] The present invention relates to additively manufactured satellites. Background Art

[0002] For spacecraft such as satellites, space is a hostile environment. Besides the dangers of damaging radiation, orbital debris impacts, and the extreme loads of launch, thermal management is particularly challenging. Many spacecraft are constructed from multiple composite parts, such as sandwich panels, that are fastened together. This construction can be strong and lightweight. However, these composite parts have high thermal impedance and are expensive and labor-intensive to manufacture and assemble. The design and production of custom spacecraft for payload-specific features, such as equipment mounting racks, is slow and expensive.

[0003] Additive manufacturing (AM) is rapidly gaining popularity across many industries as a rapid, relatively low-cost production method. AM, sometimes referred to as three-dimensional (3D) printing, can be used to create solid objects from 3D models by incrementally building them. AM typically applies a feedstock, which is then selectively joined or melted to create the desired object. The feedstock is typically applied in layers, where the thickness of each layer can vary depending on the specific technology used.

[0004] The raw material is often in granular or powder form, applied as a layer, and then selectively melted by a heat source. In many cases, the upper surface of a bed of such material is melted, and the growing workpiece is then lowered slightly into the bed itself. A new layer of raw material is then applied to the bed, and the next layer is melted onto the previous layer. For example, the granular raw material may include a thermoplastic polymer, a metal powder, a metal alloy powder, or a ceramic powder, which may be melted using a computer-controlled heat source (e.g., a scanning laser or a scanning electron beam). Exemplary methods include selective laser melting (SLM), direct metal laser sintering (DMLS), selective laser sintering (SLS), fused deposition modeling (FDM), and electron beam melting (EBM), among others.

[0005] Traditional part designs for subtractive manufacturing or composite laminate construction may be inefficient or even impossible for AM. Depending on the process and materials used, unsupported features may collapse, fine features may not be clearly defined, and / or warping and cracking may occur. New spacecraft designs are needed to compensate for the limitations and functional differences of AM while leveraging production speed, cost reduction benefits, and design freedom. Summary of the Invention

[0006] The present disclosure provides systems, devices, and methods related to additively manufactured spacecraft, such as satellites. In some examples, the satellite may include a main body and a communication device attached to the main body. The main body may have an additively manufactured outer wall structure that at least partially forms an enclosed compartment, and the communication device may be configured to receive and transmit data in space.

[0007] In some examples, a satellite may include a housing having an additively manufactured outer wall structure, a communication device coupled to the housing, and a detachment device coupled to the wall structure. The communication device may be configured to receive and transmit data in space, and the detachment device may be configured to mount and carry the housing within a launch vehicle during a launch phase and subsequently release the housing from the launch vehicle after the launch phase.

[0008] In some examples, a method of manufacturing a satellite may include: printing a wall panel configured to form a portion of an outer wall structure of the satellite; and assembling the outer wall structure of the satellite including the wall panel. The method may also include mounting a communication device to the outer wall structure, the communication device configured to receive and transmit data in space.

[0009] The features, functions, and advantages can 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. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] Figure 1 is a schematic diagram of an exemplary satellite according to aspects of the present disclosure.

[0011] Figure 2 yes Figure 1 Block diagram of a satellite.

[0012] Figure 3 is an isometric view of an exemplary additively manufactured satellite.

[0013] Figure 4 yes Figure 3 Exploded isometric view of a satellite.

[0014] Figure 5 yes Figure 3 Isometric top view of the equipment panel of a satellite.

[0015] Figure 6 yes Figure 5 A close-up cutaway view of the device panel.

[0016] Figure 7 yes Figure 3 Isometric cutaway view of a satellite's radiator panel.

[0017] Figure 8 yes Figure 7A cross-sectional view of the radiator panel along line 7-7.

[0018] Figure 9 yes Figure 3 Isometric top view of the satellite's base panel.

[0019] Figure 10 yes Figure 9 Isometric bottom view of the base panel.

[0020] Figure 11 yes Figure 9 A cross-sectional view of the base panel taken along plane AA.

[0021] Figure 12 yes Figure 9 Another cross-sectional view of the base panel taken along plane BB.

[0022] Figure 13 is a flow chart depicting the steps of an illustrative method of additive manufacturing.

[0023] Figure 14 is a schematic diagram of an exemplary additive manufacturing apparatus.

[0024] Figure 15 is a flow chart depicting the steps of an illustrative method of additively manufacturing a spacecraft according to the present teachings. DETAILED DESCRIPTION

[0025] Various aspects and examples of additively manufactured spacecraft, such as satellites, and related apparatus and methods are described below and illustrated in the associated drawings. Unless otherwise indicated, a spacecraft and / or its various components according to the present teachings may (but are not required to) include at least one of the structures, components, functions, and / or variations described, illustrated, and / or included herein. In addition, unless specifically excluded, the process steps, structures, components, functions, and / or variations described, illustrated, and / or included herein in conjunction with the present teachings may be included in other similar apparatus and methods, including being interchangeable between the disclosed examples. The following descriptions of various examples are merely illustrative in nature and are in no way intended to limit the present disclosure, its applications, or uses. In addition, 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.

[0026] This detailed description includes the following sections: (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 subsections A through D, with each subsection labeled accordingly.

[0027] Overview

[0028] In general, a spacecraft according to the present teachings may include an additively manufactured primary structure. Primary structures may also be referred to as bodies, shells, wall structures, and / or frames. Primary structures may be described as those components of a spacecraft that are designed to transfer loads through the spacecraft to interfaces of a launch vehicle and / or deployment system and provide attachment points for payloads and associated equipment or components. Primary structures may also be described as the main load-bearing elements that provide the most direct and efficient load path from spacecraft components to launch vehicle interfaces. Examples of spacecraft may include, but are not limited to, satellites, space stations, crewed spacecraft, and / or interstellar probes.

[0029] Additive manufacturing of a spacecraft's primary structure can allow for design flexibility in the spacecraft, as well as reduce production cycle time, labor, costs, and post-production testing. In some examples, the primary structure can include multiple additively manufactured panels. The panels can conform to standard designs and / or include customized features depending on the desired functionality. For example, one or more panels can be configured for passive thermal management and heat removal. In another example, one or more panels can be configured for radiation shielding, one or more panels can be configured for structural load support, and / or one or more panels can be configured to support payload equipment.

[0030] The primary structure can be configured for additive manufacturing with limited or no sacrificial material, which can be referred to as auxiliary supports. More specifically, features of each panel can be configured for printing without auxiliary supports to minimize or eliminate wasted material and the labor of removing supports. Each panel can be additively manufactured or printed as a single, unified structure. The panels can also be described as monolithic. Structures and / or features that are separated from or added to the primary structure after manufacturing in a conventionally manufactured spacecraft can be integrated into the printed panel. For example, fastener holes, radiation point shielding, localized reinforcements or stiffeners, access points and / or equipment mounting brackets can be printed as part of the monolithic panel structure.

[0031] Panels for the primary structure can be designed to be printed simultaneously in a selected additive manufacturing machine or printer. In some examples, auxiliary structures and / or equipment for the satellite can also be additively manufactured. For example, propellant tanks, thruster brackets, and / or communication antennas can be additively manufactured.

[0032] Examples, Parts, and Alternatives

[0033] The following sections describe selected aspects of exemplary additively manufactured spacecraft and related systems 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 may include one or more different examples and / or contextual or related information, functionality, and / or structure.

[0034] A. Exemplary Satellites and Associated Methods

[0035] The examples disclosed herein may be used in an exemplary satellite launch method (see Figure 1 ) and exemplary satellite 100 (see Figure 1 and Figure 2 ) is described in the context of FIG. As described above, satellite 100 is an example of a spacecraft. In this example, the method includes three phases: a launch phase 20, a separation phase 30, and a deployment phase 40. Launch phase 20 may include transporting satellite 100 from a planetary body 120, such as Earth, to outer space 122 (also referred to as space) using a launch vehicle 124. In some examples, launch vehicle 124 may be a rocket-powered vehicle. In the context of Earth, outer space may be a region beyond the Kármán line. Separation phase 30 may include separating satellite 100 from launch vehicle 124 once a desired position, trajectory, and / or orbit is achieved. Deployment phase 40 may include preparing satellite 100 for operation, such as establishing communication with a controller on planetary body 120, extending solar panels or instrument arms, and / or maneuvering into a desired orientation relative to the planetary body. In some examples, the method may also include design, production, and / or commissioning phases.

[0036] The various processes of the launch method can be performed or implemented by a system integrator, a third party, and / or an operator (e.g., a customer). For the purposes of this description, a system integrator may include (but is not limited to) any number of aviation manufacturers and master system distributors; a third party may include (but is not limited to) any number of vendors, subcontractors, and suppliers; and an operator may be a telecommunications company, a leasing company, a military entity, a service organization, etc. The devices and methods shown or described herein may be employed during any one or more stages of the satellite launch method.

[0037] like Figure 2 As shown, satellite 100 may include a bus 102 for connecting to multiple satellite systems, a payload 104, and a separation system 106 or separation device. Examples of multiple systems include one or more of a main structure 108 (e.g., a main body), a propulsion system 110, a power system 112, a thermal management system 114, a radiation shielding system 116, and a communication system 118 or communication device. Depending on the functions involved, each system may include various subsystems such as controllers, processors, actuators, actuators, motors, generators, etc. 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 technologies, such as launch vehicles, space stations, manned spacecraft, and / or interstellar probes.

[0038] B. Example additively manufactured satellite

[0039] like Figures 3 to 12As shown, this section describes an exemplary additively manufactured satellite 200. As described above, satellite 200 is an example of an additively manufactured spacecraft. Figure 3 As best shown in Example A, satellite 200 is generally cube-shaped and may be described as a CubeSat and / or a microsatellite. The satellite includes a body 210 that forms the primary structure of the satellite as described above and is an example of primary structure 108 as described in Example A. Body 210 may also be described as a shell.

[0040] The body 210 is composed of a plurality of additively manufactured wall panels 212 fastened together. The wall panels 212 can be described as the outer wall structure of the satellite. In the depicted example, the width of the body 210 is between approximately 10 and 30 inches (250 and 800 mm). In some examples, the width of the body can be between approximately 16 and 20 inches (400 and 500 mm).

[0041] In this example, satellite 200 is designed to closely replicate the specifications of existing satellite designs. This replication can allow existing methods of launch and flight control, payload configuration, etc. to be used with satellite 200, thereby facilitating the rapid implementation and / or replacement of traditionally manufactured satellites. In some examples, satellite 200 can be designed independently of traditionally manufactured satellite geometries or specifications to facilitate use with updated and / or new methods, payloads, etc.

[0042] like Figure 3 As shown, separation system 214, solar panels 216 and panel antenna 218 are mounted to body 210. Separation system 214 is an example of separation system 106 as described in Example A, and can also be described as a launch vehicle interface ring. The separation system is configured to serve as the only connection between satellite 200 and the rocket-powered vehicle during launch, and then facilitate disconnection of the satellite from the vehicle. In the depicted example, separation system 214 is a belt-type separation system such as a clamping ring or a maneuvering light belt (MLB). In some examples, the satellite may employ other separation systems such as a distributor-type Quadpack or a Canisterized Satellite Distributor (CSD).

[0043] As described in Example A, solar panels 216 are configured to supply power to satellite 200 and are integrated with the satellite's electrical system (e.g., power system 112). Depending on the desired functionality, the electrical system may also include one or more batteries, controllers, transformers, switches, printed circuit boards, wiring, etc. The electrical system may provide power to one or more satellite systems and / or payload equipment.

[0044] Panel antenna 218 is configured to cooperate with the satellite's communication system to transmit and receive data in space. The communication system described in Example A is an example of communication system 118. Depending on the desired functionality, the communication system may include one or more processors, encoders, modulators, transmitters, receivers, data storage devices, additional antennas, etc. In some examples, in addition to or in place of panel antenna 218, the communication system of satellite 200 may include an additively manufactured antenna. For example, the communication system may include an antenna as disclosed in U.S. Patent Publication No. 2019 / 0291186A1, which is incorporated herein by reference in its entirety.

[0045] Figure 4 is an exploded view of satellite 200. As shown, the wall panels 212 of body 210 include an equipment panel 300, four radiator panels 400, and a base panel 500. The equipment panel can also be described as a top panel or front panel, and the base panel can also be described as a bottom panel or rear panel. In this example, wall panels 212 are shaped for a cuboidal satellite. In other words, each panel is approximately square in shape. In some examples, the panels can have other shapes suitable for other satellite designs. For example, the panels can have triangular, pentagonal, or irregular shapes.

[0046] Wall panels 212 are assembled into the main body 210, forming an enclosed interior compartment 220. Satellite and payload equipment can be housed within the enclosed compartment, mounted to and protected by the wall panels 212. Each wall panel is configured to serve as a structural component of the main body, and the wall panels are assembled so that loads on the main body 210 are efficiently transferred between the panels. In this example, the wall panels 212 are directly fastened together by four corner posts 222.

[0047] The radiator panels 400 are indirectly fastened together by the corner posts 222. The radiator panels are directly fastened to each of the equipment panel 300 and the base panel 500. As discussed further below, with reference to Figure 5 and Figure 9 The device panel and the base panel each include a sidewall configured to overlap the radiator panel. Fasteners extending through aligned holes in the sidewalls and the radiator panel secure the panels together.

[0048] Corner posts 222 are elongated members that each extend from a corner of base panel 500 to a corresponding corner of equipment panel 300. These posts can also be described as brackets or clips and have a right-angled or L-shaped cross-section. They are similar in shape to angle irons and each has two legs. In the depicted example, the corner posts are additively manufactured.

[0049] Each leg includes a plurality of holes arranged in a line and configured to align with a plurality of corresponding holes in one of the radiator panels 400, such that fasteners extending through the aligned holes secure the panel to the corner post. Each corner post 222 overlaps a pair of radiator panels 400 that meet at a corresponding corner. Each corner post is disposed on the interior of the radiator panel and on the exterior of the base panel 500 and the device panel 300.

[0050] To reduce unnecessary weight and material, most of the legs of corner posts 222 are notched or recessed depending on the location of the holes. Where additional structural strength or load tolerance is required, the corner posts may have solid and / or reinforced legs. In this example, solar panel 216 is mounted at the corner of body 210 where two radiator panels 400 meet, near one of corner posts 222. The solar panel can be fastened to the corner post and radiator panel. For improved support, the corner posts include solid, non-recessed legs.

[0051] In some examples, all of the wall panels 212 can be directly fastened together. However, the use of one or more structural components, such as corner posts 222, can facilitate the removal of one or more of the wall panels 212 during assembly of the satellite 200 (e.g., to allow payload equipment to be installed in the enclosed interior compartment 220). These structural components can be additively manufactured so that all of the body 210 can be quickly and flexibly designed and produced.

[0052] like Figure 4 As shown, satellite 200 also includes a propulsion system 224. In the depicted example, the propulsion system includes a propellant tank 226 and four thrusters mounted in a bracket 228. The propulsion system may also include regulators, valves, a feed system, etc., depending on the desired functionality. Propellant tank 226 is configured to store chemical propellants (e.g., hydrazine). Both the propellant tank and bracket 228 are additively manufactured and composed of laser-sintered titanium alloy. In some examples, satellite 200 can be configured for alternative propulsion technologies such as electrospray, Hall effect thrusters, or ion engines.

[0053] At launch, the propellant tank 226 and the stored propellant may represent a substantial portion of the weight of the satellite 200. Consequently, the propellant tank may generate significant vertical loads during launch. Accordingly, the propellant tank 226 is mounted to the base panel 500 opposite and in axial alignment with the separation system 214. The base panel is configured to efficiently transfer loads from the propellant tank to the separation system and launch vehicle, as described below with reference to FIG. Figure 11 Further discussion.

[0054] In the depicted example, the wall panels 212 each comprise a laser-sintered aluminum alloy and are printed using direct metal laser sintering (DMLS). Such metal alloys can provide a favorable strength-to-weight ratio. Generally, the panels can comprise any material having properties suitable for a satellite (e.g., strength, stiffness, and weight) and can be manufactured using any effective additive manufacturing method. For example, the wall panels can be produced from a polymer using fused deposition modeling (FDM) or from a titanium alloy using electron beam melting (EBM).

[0055] Each wall panel 212 is monolithic. That is, each panel is printed as a single unit. Each panel can be printed in a series of layers perpendicular to the build axis. The build axis can be defined by the orientation of the panel relative to the printer or other additive manufacturing equipment during printing.

[0056] Some or all of the wall panels 212 can be configured for simultaneous printing. More specifically, a selected subset of panels can be designed to be printed with a build axis and / or footprint that allows the panels of that subset to be printed in a single deposition. In this example, all four radiator panels 400 can be printed in a single deposition or printing cycle. In some examples, all components of the body 210 can be designed to be printed in a selected number of sequential depositions on a single printer. For example, on the same printer, the radiator panel 400 can be printed in a first deposition, the device panel 300 and the corner posts 222 can be printed in a second deposition, and the base panel 500 can be printed in a third deposition.

[0057] Each wall panel 212 may include one or more customized structural features 242. The customized features may vary depending on the mounting, connection, shielding, and / or view factor requirements of the satellite's payload and / or operational equipment, as well as launch methods, weight constraints, or any other relevant considerations. Additive manufacturing of the wall panels 212 may facilitate these customized features, thereby allowing a standardized design to be quickly and inexpensively modified for each satellite produced.

[0058] The equipment panel 300 is configured to support and shield equipment (not depicted) during launch and operation. The equipment may include payload equipment such as sensors, optics, and data processors, and / or may include satellite equipment such as electrical and communication systems. The equipment panel may serve as a structural support for the equipment to which the equipment is fastened or otherwise mounted. The equipment panel may also shield the satellite equipment from space radiation such as solar wind and cosmic rays. The panel may be configured to block radiation based on the satellite's intended orbit and / or the radiation tolerance of the shielded equipment. That is, the panel may block sufficient radiation to reduce the radiation level inside the satellite to a safe level for the installed equipment. For example, in low Earth orbit for non-radiation hardened electronic equipment, the panel may block at least 0.5 to 1 megarad of radiation per year. As another example, in geosynchronous orbit for radiation hardened electronic equipment, the panel may block at least approximately 500 megarads per year.

[0059] Radiator panel 400 is configured to serve as passive thermal management for satellite 200. The panel's integral metal sandwich structure allows heat to be conducted from the equipment installed in interior compartment 220 to the panel's outer surface and radiated into space. In some examples, the radiator panel can be configured to radiate heat at a rate of at least 5 watts per square foot, 20 to 40 watts per square foot, and / or at least 50 watts per square foot. The rate of thermal radiation or heat dissipation can be adjusted and / or tailored based on the heat generation properties of the equipment installed in the satellite. The radiator panel can also be configured to act as a radiation shield and / or a micrometeoroid and orbital debris (MMOD) shield. In some examples, the radiator panel can be integrated into the satellite's thermal management system, including passive and / or active thermal management such as heat pipes, cold plates, and / or thermoelectric coolers.

[0060] The base panel 500 is configured to support vertical loads from the equipment panel 300 and the radiator panel 400, as well as the equipment mounted on the panels. The base panel is configured to connect to the separation system 214 and transfer the loads from the body 210 through the separation system to the launch vehicle. The base panel 500 is also configured to support the propulsion system 224 (including the propellant tank 226 and the thruster bracket 228).

[0061] Each wall panel 212 is configured and positioned according to the equipment and systems of satellite 200. That is, the design and location of each panel can be selected based on the nature or requirements of the specific equipment mounted to and / or adjacent to the panel, the structural or orientation requirements of satellite systems such as propulsion and communications, and / or the expected conditions during launch and operation. For example, in the depicted example, equipment panel 300 is located on the top side of satellite 200, opposite separation system 214, so that the panel is aligned with the launch axis. For another example, equipment panel 300 can be located on the side of satellite 200 expected to face the sun in order to shield radiation-sensitive equipment from the solar wind.

[0062] Figure 5 FIG3 is an isometric view of the outside of an equipment panel 300. The panel includes a cover panel 330, which is generally square and flat and can also be described as a plate. The cover panel 330 has a first side 333, a second side 335, and four linear outer edges. A stiffening structure 336 extends from the cover panel's first side 333 on the panel's outside 332.

[0063] In this example, most of the stiffening structures 336 form a grid of equilateral triangles 340. The stiffening structures may also be described as forming an array of equilateral triangles. The covering panel 330 and the stiffening structures 336 together may be described as an iso-grid or an iso-grid panel. In general, the stiffening structures 336 may be formed in any pattern suitable for the desired stiffness of the equipment panel 300. For example, the stiffening structures may form a rectangular grid and / or may form a grid of varying density to allow for positionally varying stiffness. Some patterns of stiffening structures 336 may be preferred because they are more efficient for additive manufacturing, as described below with reference to Figure 6 Further described.

[0064] The undepicted inner side of the cover panel 330 may be flat, except for the custom structural features 242. Positioning the stiffening structure on the outer side of the cover panel may facilitate mounting equipment on the inner side and increase the usable space inside the satellite.

[0065] In the depicted example, the customized structural features 242 of the equipment panel 300 include a plurality of threaded fastener holes 344, a fluid exchange connection 346, an external mounting recess 348, and a star tracker viewing window 350. Other possible customized structural features not included in this example include electrical connections, solar panel and / or antenna mounting brackets, and hole patterns that match the communication array. The customized structural features 242 can be built into the cover panel 330 and / or the stiffening structure 336. For example, the fluid exchange connection 346 includes a hole through the cover panel 330 and a modification to the stiffening member 336 to define a recess that receives a seal.

[0066] Four side walls 338 extend from the outer edges of the cover panel 330 and include a plurality of fastener holes 352. In this example, when the equipment panel 300 is assembled into the body of the satellite, the side walls extend toward the interior of the satellite. The side walls 338 are configured to facilitate connection of the equipment panel 300 to other panels of the satellite body. Figure 4 Each sidewall overlaps one of the radiator panels 400, contacting the inside of the panel so that the fastener holes 352 align with corresponding holes printed in the radiator panel. Fastener assemblies extending through the aligned holes secure the two panels together.

[0067] At the corner where the two side walls 338 of the device panel 300 meet, each of the two side walls deviates slightly from being flat to form a corner recess 374, which can be Figure 5 Each corner recess 374 is configured to receive the end of one of the corner posts 222. Figure 4 As discussed, the corner posts are received between the side walls 338 and the heat sink panel 400. The corner recesses 374 allow the side walls 338 to also directly contact the heat sink panel without gaps.

[0068] Similar to corner posts 222, side walls 338 have notched edges to allow for desired locations of fastener holes 352 relative to cover panels 330 while minimizing the material and weight of equipment panel 300. In some examples, equipment panel 300 may include other connection features in addition to or in place of side walls 338. For example, instead of side walls 338, a satellite may include four horizontally oriented corner posts.

[0069] Figure 6 FIG3 is a partial cross-sectional view of equipment panel 300, showing stiffening structure 336 in greater detail. Stiffening structure 336 can be described as a uniform grid flange stiffener and / or a T-beam. In combination with decking panel 330, the stiffening structure can be described as forming an I-beam and / or as a uniform grid panel with flanges. The stiffening structure includes a web portion or web 360 and a flange portion or cap 362.

[0070] Webs 360 extend perpendicularly to the cladding panels 330 from the outer sides 332 thereof. Caps 362 are centered on webs 360, spaced apart from the cladding panels 330 by the webs, and extend parallel to the cladding panels. The caps can also be described as extending in two opposing directions from the distal ends of webs 360. Webs 360 and caps 362 are each generally planar and / or plate-like and extend linearly to form a grid of triangles 340. At the intersections of the grid, the webs and caps can be described as forming nodes having circular and / or hexagonal shapes.

[0071] In some examples, the stiffening structure 336 may have alternative geometries. For example, the stiffening structure may include only the web 360, or the cap 362 may extend from the distal end of the web in only one direction. The depicted I-beam geometry of the stiffening structure may be preferred because it provides stiffness with minimal added weight.

[0072] The stiffening structure 336 can be configured for printing without the use of auxiliary supports. Both the linear portions and the nodes of the web and cover can be configured for printing without the use of sacrificial or auxiliary supports. More specifically, all surfaces of the stiffening structure 336 can be oriented relative to the build axis 364 of the device panel 300 so that the stiffening structure can be printed without auxiliary supports. For example, the web 360, cover 362 and overall pattern of the stiffening structure can form an angle of no more than 45 degrees relative to the build axis 364. A grid of equilateral triangles 340 can facilitate this orientation, where other patterns, such as squares, may not have an orientation that meets this angular requirement.

[0073] The dimensions of the stiffening structure 336 can be adjusted to achieve the desired stiffness of the equipment panel 300. The web 360 has a depth 366 and the cover 362 has a width 368. In this example, the web depth 366 is between approximately 0.1 and 0.3 inches, and the cover width 368 is between approximately 0.25 and 0.75 inches. Each triangle in the grid of equilateral triangles 340 has a side length 370. In this example, the side length 370 is between approximately 1 and 5 inches.

[0074] Varying the cap width 368 can change the moment of inertia of the stiffening structure, thereby changing the stiffness of the equipment panel.Varying the triangle side length 370 can change the density of the stiffening structure 336 on the cover panel 330, thereby changing the stiffness of the equipment panel.

[0075] The thickness of the equipment panel 300 can also be adjusted to achieve desired panel properties. The cover panel 330 has a thickness 372. In this example, the web 360 and the cover 362 each have the same thickness as the cover panel 330 (between approximately 0.05 and 0.125 inches). This range can provide adequate radiation shielding for typical satellite equipment without adding unnecessary weight. The appropriate thickness can depend at least in part on the material of the cover panel 330. In some examples, the web and cover can have different thicknesses.

[0076] In this example, each of the dimensions depth 366, width 368, length 370, and thickness 372 is constant across the device panel 300. In some examples, one or more of these dimensions can vary across the panel to achieve localized or targeted changes in the properties of the panel. For example, a region of the panel can have a reduced length 370 and increased width 368 to provide the additional rigidity needed to support particularly bulky equipment mounted near that region. For another example, a region of the panel can have an increased thickness 372 to provide additional shielding for particularly radiation-sensitive equipment.

[0077] Figure 7 and Figure 8is a cross-sectional view of radiator panel 400. The following description can be understood to apply to any of the four radiator panels. Radiator panel 400 is an example of an additively manufactured truss panel and includes a first skin 410 and a second skin 412 connected by a truss structure 414. The radiator panel can be described as a unitary sandwich structure. As further described below, the individual skins and truss structures can be shaped, characterized, and / or configured according to the desired properties of radiator panel 400.

[0078] Fabricating skins 410, 412 and trusses 414 as a unified metal structure can improve the thermal performance and heat removal of the panel by eliminating the impedance caused by the adhesive materials used in traditionally manufactured sandwich panels. In particular, aluminum alloys can provide efficient heat radiation and / or heat removal. These thermal properties can be particularly beneficial for thermal management in a space environment.

[0079] In this example, the radiator panel 400 is approximately half an inch thick. Each of the skins 410, 412 is planar and has an outer surface 418 and an inner surface 420. The first skin 410 is parallel to the second skin 412. In this example, each skin has a uniform thickness throughout and has the same thickness of approximately 0.02 inches. In some examples, the thickness of one or both skins may vary, and / or one skin may have a greater thickness than the other.

[0080] The outer surface 418 of each of the skins 410, 412 includes a stiffening texture 419. The texture is a regular, repeating pattern on the outer surface except where interrupted by custom features of the radiator panel. For example, custom features may include recesses for solar panel mounting brackets, fastener holes corresponding to corner posts or other panels, windows for panel antennas, etc. In this example, the stiffening texture 419 is a square grid of linear sections that includes circular sections at each intersection of the linear sections. The texture is raised from the outer surface 418 but does not overhang the surface. Thus, the texture can improve the stiffness of the skins 410, 412 without reducing radiation efficiency or creating light traps on the radiator panel 400.

[0081] The truss structure 414 includes a plurality of truss members 424 extending between nodes 426. Figure 8 As shown, each truss member extends between a node on the inner surface 420 of the first skin 410 and a node on the inner surface of the second skin 412. Each truss member 424 is approximately cylindrical in shape.

[0082] Truss structure 414 comprises an array of core structures. Each core structure may comprise a framework of truss members 424 arranged according to any suitable geometry. The geometry of the core structure may be uniform across radiator panel 400 or may vary depending on the position within the panel. For example, the core structure may vary depending on the expected non-uniform loading of radiator panel 400. The core structures may be arranged in a grid, a repeating pattern, and / or in an efficient manner.

[0083] In some examples, the core structure can be a box frame. Such a frame can include four pairs of crisscrossed cylindrical truss members, each truss member extending diagonally between skins 410, 412. In the depicted example, the core structure is a pyramid frame 434 arranged in a grid. Each core pyramid 434 includes four truss members 424.

[0084] like Figure 8 As shown, each truss member 424 extends from a separate base node 438 to a single common upper node 440. The base node 438 is located at the corner of a square, and each core pyramid 434 can be described as having a base at the first skin and an apex at the second skin. Alternatively, the truss members 424 of the truss structure 414 can be considered to belong to an array of core pyramids, each having a base at the second skin and an apex at the first skin. The radiator panel 400 can accordingly have a symmetrical response to the load.

[0085] In this example, the truss structure 414 also includes a plurality of triangular stiffeners 428 on the inner surface 420 of the first skin 410 and the second skin 412. In this example, each stiffener has a maximum thickness of approximately 0.04 inches, or twice the thickness of the skins 410, 412. The maximum thickness of the stiffeners 428 is also approximately equal to the diameter of the truss member 424.

[0086] Each node 426 is located at a stiffener, and the radiator panel 400 includes a checkerboard pattern of core pyramids 434 and an offset grid pattern of stiffeners 428. The first skin 410 is covered by the array of core pyramids without gaps or overlaps, with the base of each core pyramid meeting the base of an adjacent core pyramid along the stiffeners 428 of the first skin 410.

[0087] The stiffeners 428 may improve the structural properties (eg, stiffness and buckling stability) of the skins 410, 412 and may strengthen the connection between the truss members 424 and the skins. Including stiffeners may allow the thickness of the skins to be reduced without sacrificing desired structural properties.

[0088] Radiator panel 400 has a major axis 422 extending parallel to skins 410, 412, and a minor axis 430 perpendicular to the major axis and the skins. In this example, major axis 422 also serves as the build axis for radiator panel 400. In other words, the panel is printed such that major axis 422 is aligned with the build direction of the additive manufacturing apparatus. Because major axis 422 is parallel to skins 410, 412, the skins can be described as being printed vertically. Each of the outer surface 418 and inner surface 420 of first skin 410 and second skin 412 can extend in the build direction throughout the print.

[0089] This orientation can allow the skin to be printed with a good surface finish and limited warping without auxiliary supports.This orientation can also reduce the footprint size of the radiator panel 400 and allow multiple panels to be printed simultaneously in a single deposition run.

[0090] The stiffeners 428 and core pyramids 434 are configured to be self-supporting during printing of the heat sink panel 400. More specifically, the triangular cross-sectional shape of the stiffeners is shaped and oriented to allow for self-support, and the truss members 424 of the core pyramids each extend at an angle selected to allow for self-support. Figure 8 As shown, each side of the triangular cross-sectional shape of the stiffener 428 extends at a stiffener angle 432 relative to the secondary axis 430. The angle 432 may be at least 35 degrees. In the depicted example, the angle 432 is approximately 40 degrees.

[0091] Figure 8 Also shown is a first truss angle 442 between truss member 424 and minor axis 430. First truss angle 442 lies in a plane defined by the major and minor axes of radiator panel 400. A second truss angle (not shown) is also defined between truss member 424 and minor axis 430 in a plane perpendicular to the major axis but including the minor axis. Each truss angle can be at least 35 degrees and no more than 50 degrees. An angle of 35 degrees or greater allows the truss members to be printed without auxiliary supports. On the other hand, an angle of no more than 50 degrees ensures that the truss formed by the truss members is sufficiently strong. In this example, core pyramid 434 has a square base, resulting in the first and second truss angles being equal. In this example, the truss angles are each 40 degrees.

[0092] Each core pyramid 434 can also be configured according to the desired structural properties of the radiator panel 400. For example, the first and / or second truss angles can be increased or decreased to adjust the panel to a desired stiffness. For another example, the diameter of each truss member 424 can be increased to improve panel strength or decreased to reduce the overall panel weight.

[0093] Figure 9 and Figure 10The inner side 510 and the outer side 512 of the base panel 500 are shown separately. When the base panel is assembled into a satellite, the inner side 510 faces the closed interior compartment of the satellite and the outer side 512 faces the external environment.

[0094] The base panel 500 includes an inner skin 514, an outer skin 516, and four side walls 518. The inner skin and side walls define the square shape of the base panel, while the outer skin has a generally undulating cross or x-shape. The inner skin 514 has a similar cross-shaped main portion 520 and four raised corner portions 522. The inner skin 514 and the outer skin 516 meet and join at the perimeter of the outer skin and inner skin main portion 520. The main portion 520 of the inner skin 514 is flat, and the outer skin 516 is convex relative to the inner skin.

[0095] The base panel 500 also includes a central interface ring 524 and a plurality of thruster mount interface rings 526. In this example, the base panel includes four thruster mount interface rings corresponding to the satellite's four thrusters. Each thruster mount interface ring 526 is disposed in one of the raised corner portions 522. The location of the thruster mount interface rings and / or the shape and profile of the inner and outer skins 514, 516 may be determined by the number and type of thrusters and / or thruster mounts used in the satellite. Preferably, the rings may be arranged symmetrically within the base panel 500.

[0096] Each thruster bracket interface ring 526 includes a circular aperture 528 and a surrounding annular mounting surface 530. The aperture 528 may also be described as an opening or hole through the base panel 500. The aperture 528 extends through the raised corner portion 522 of the inner skin 514. The mounting surface 530 is shaped and reinforced to conform to and support the outer edge of the thruster bracket when mounted to the base panel 500 over the aperture 528. A plurality of fastener holes also extend through the raised corner portion 522 of the inner skin 514 adjacent the mounting surface 530 to facilitate installation of the thruster bracket.

[0097] The central interface ring 524 includes a larger circular hole 532, an inner annular surface 534, and a stepped outer annular surface 536. The inner annular surface 534 includes the inner side of the annular portion 535 of the inner skin 514, and the outer annular surface 536 includes the outer side of the annular portion 537 of the outer skin 516. The central interface ring is configured for connection between the propellant tank 226 and the separation system 214 (see Figure 4 ). The propellant tank is mounted to the annular inner surface 534, and the separation system is mounted to the stepped annular outer surface 536. The size and / or shape of each annular surface can be determined according to the specifications of the corresponding mounted structure. For example, before printing the base panel 500, the width of the inner annular surface 534 in the panel design can be modified to match the connector flange of the new propellant tank.

[0098] As shown in the cross section of the base panel 500 along the plane AA Figure 11 As shown, base panel 500 is hollow. The main portion 520 of inner skin 514 is spaced apart from outer skin 516, except where the two skins are joined at the periphery. At a central interface ring 524, the inner and outer skins are joined by a vertical annular wall 540. The annular portion 535 of inner skin 514, the annular portion 537 of outer skin 516, and the annular wall 540 together form an I-beam structure. This structure strengthens base panel 500 and facilitates payload transfer between the propellant tank mounted to the inner annular surface 534 and the separation system mounted to the outer annular surface 536.

[0099] Annular wall 540 also includes a plurality of fastener posts 542 circumferentially spaced evenly around central interface ring 524. Each post is centrally located on annular wall 540, extends from inner skin 514 to outer skin 516, and includes a fastener hole. Each hole extends from inner annular surface 534 through a corresponding fastener post to outer annular surface 536. The fastener posts facilitate secure mounting of the propellant tank and separation system, as well as strengthening annular wall 540 and improving load transfer.

[0100] like Figure 12 As can be seen in FIG. 5A (another cross-section of the base panel along plane BB), the hollow interior of the base panel 500 is divided into separate interior cavities 544 by a plurality of stiffening ribs 546. In this example, the base panel 500 includes four interior cavities 544 and four stiffening ribs 546. Each stiffening rib is flat, extending vertically between the outer skin 516 and the inner skin 514 and horizontally between the annular wall 540 and the outer periphery of the outer skin. The stiffening ribs 546 can be described as being each centered relative to an arm of the cruciform outer skin and extending parallel to the respective arm.

[0101] The stiffening ribs 546 can carry shear loads passing through the base panel 500 and increase panel stiffness. The depth of the internal cavity 544 and the base panel 500 can improve the base panel's resistance to bending loads. The hollow design of the base panel 500 can save weight, material costs, and printing time while maintaining the desired panel stiffness and strength. The shape and location of the stiffening ribs 546 and / or internal cavity 544 can be determined based on the shape of the outer skin 516, the number and type of thrusters mounted to the base panel, the mass of the propellant tanks and stored propellant, expected bending loads, and / or any relevant structural considerations. Any number and / or pattern of ribs can be used.

[0102] In this example, the inner skin 514, outer skin 516, sidewalls 518, and stiffeners 546 have approximately the same thickness. Each structure is approximately 0.08 inches thick. In some examples, the thickness of these structures may be different and / or the thickness of any one structure may vary. Preferably, the inner skin 514 and outer skin 516 may be at least approximately 0.05 inches thick and no more than approximately 0.125 inches thick. The annular portions 535, 537, and annular wall 540 are reinforced relative to the inner and outer skins and may be between approximately 1.5 and 2 times the thickness of the inner and / or outer skins. The depth of the internal cavity 544 tapers to a maximum depth of between approximately 0.5 and 2 inches.

[0103] Refer again Figure 9 and Figure 10 , sidewall 518 includes a plurality of fastener holes 548. In this example, when base panel 500 is assembled into the body of the satellite, the sidewalls extend toward the interior of the satellite. Sidewall 518 is configured to facilitate connection of base panel 500 to other panels of the satellite body. Return to Brief Reference Figure 4 , each sidewall overlaps one of the radiator panels 400, contacting the inside of the panel so that the fastener holes 548 align with the corresponding holes printed in the radiator panel. Fastener assemblies extending through the aligned holes secure the two panels together.

[0104] At the corner where the two side walls 518 of the base panel 500 meet, each of the two side walls deviates slightly from being flat to form a corner recess 550. Each corner recess 550 is configured to receive the end of one of the corner posts 222. Figure 4 As discussed, and similar to the device panel 300, the corner posts are received between the side walls 518 and the heat sink panel 400. The corner recesses 550 allow the side walls 518 to also directly contact the heat sink panel without gaps.

[0105] Similar to corner posts 222, sidewalls 518 have notched edges to allow for desired locations of fastener holes 548 relative to inner skin 514 while minimizing the material and weight of base panel 500. In some examples, base panel 500 may include other connection features in addition to or in lieu of sidewalls 518. For example, instead of sidewalls 518, a satellite may include four horizontally oriented corner posts.

[0106] As described above, the base panel 500 may also include one or more customized structural features 242. In the depicted example, the customized structural features 242 of the base panel 500 include Figure 9 The electrical connector 552 shown, Figures 9 to 12 The access port 554 is shown extending through the panel and Figure 10 Mounting point 556 is shown for devices to connect via access ports.

[0107] C. Exemplary Methods of Additive Manufacturing

[0108] See also Figure 13 , this section describes the steps of an illustrative method 600 for additive manufacturing of a workpiece. Figure 14 Aspects of the exemplary additive manufacturing apparatus depicted in the accompanying drawings can be used in the method steps described below. Where appropriate, references are made to components and systems that can be used to perform various steps. These references are for illustrative purposes only and are not intended to limit the possible manner in which any particular step of the method may be performed.

[0109] In some examples, method 600 may also be referred to as a 3D printing method. The terms additive manufacturing and 3D printing may both be understood to include processes that create objects by adding material in successive layers. Additive manufacturing may be understood as a broader term that encompasses 3D printing. In some examples, the terms additive manufacturing and 3D printing may be used interchangeably. In the present disclosure, printing and / or printing steps may be understood to include creating by any additive manufacturing method. Examples of additive manufacturing processes include (but are not limited to) material extrusion, powder bed fusion, material jetting, binder jetting, directed energy deposition, photopolymerization, and sheet lamination.

[0110] Figure 13 is a flow chart showing steps performed in an exemplary method and may not list the complete process or all steps of the method. Figure 13 Various steps of method 600 are depicted in FIG. 6 , but these steps need not necessarily be performed in their entirety and in some cases may be performed simultaneously or in an order different from that shown.

[0111] At step 610, digital information describing a plurality of ordered layers is received. The digital information may be provided by Figure 14 The computer controller 712 of the additive manufacturing device 710 depicted in FIG. 7A receives the digital design information. The additive manufacturing device may also be referred to as a printer or a fabricator. The computer controller 712 may include any data processing system configured to receive digital design information and control the functions of the printer 710. Figure 14 The exemplary computer controller shown includes a processor 714 for controlling printer functions and a memory 716 for storing received data.

[0112] The received information may include geometric data and / or design details of a plurality of two-dimensional patterns of layers that make up a three-dimensional object, wherein the three-dimensional object is a workpiece to be manufactured 728. The layers may also be described as cross-sections or slices. The plurality of layers may be ordered such that the layers may be numbered or organized from the first layer to the last layer.

[0113] Step 612 of method 600 includes depositing feedstock on a build platform 718 located in a build environment 720 of printer 710. The build platform can include a support movable along a build axis 722 by computer controller 712. The build platform can have a planar surface perpendicular to the build axis 722.

[0114] The feedstock can be any material suitable for additive manufacturing, typically in the form of a fluid or powder, and includes, but is not limited to, photopolymer resins, thermoplastics, plaster, ceramics, and metals. The material can be dispensed from a feedstock source 724, such as a hopper, tank, or powder bed. For example, aluminum powder can be brushed from a powder bed onto a build platform 718 by a brush arm actuated by computer controller 712.

[0115] The raw material can be evenly distributed on build platform 718 or can be deposited in a selected pattern. Deposition can be performed under the control of computer controller 712. In some examples, build platform 718 can be immersed in the raw material, and deposition can be achieved by gravity or fluid pressure. In some examples, print head 726 connected to raw material source 724 can deposit the raw material in a pattern corresponding to the first layer of the ordered plurality of layers.

[0116] At step 614, the feedstock is altered to produce a first layer. In other words, according to the design information describing the first layer of the ordered plurality of layers and directed by computer controller 712, the deposited material is physically altered to realize the first layer as a physical object on the build platform.

[0117] The material may be acted upon by a print head 726 of the printer 710, which is controlled by a computer controller 712. For example, the print head may include a laser that cures a photopolymer by exposure to light or sinters a metal powder by exposure to heat. The print head may be directed by the computer controller 712 to follow a path defined for the first layer in the received digital information and / or a path calculated by the processor 714 based on the received digital information.

[0118] Step 616 includes repositioning the build platform. In some examples, build platform 718 may begin at a selected distance from print head 726. The selected distance may be determined by the processes performed by the print head. After generating a layer, the build platform may be repositioned by computer controller 712 along build axis 722 by the thickness of the layer away from print head 726. In other words, the build platform may be moved so that the top surface of the generated layer is the selected distance from print head 726.

[0119] In some examples, build platform 718 can begin aligned with another element of printer 710 (e.g., a material dispensing component). After generating a layer, build platform can be repositioned by computer controller 712 along build axis 722 so that the top surface of the generated layer is aligned with another element of printer 710. In some examples, at step 616, print head 726 can be repositioned instead of or in addition to build platform 718. In some examples, step 616 can be skipped.

[0120] At step 618, a feedstock is deposited over the layer generated in the previous step of method 600. As described with respect to step 612, the feedstock can be any suitable material and can be deposited in any suitable manner. At step 620, the feedstock is changed to generate the next layer, as previously described with respect to step 614.

[0121] Steps 616 to 620 may be repeated to generate each of the multiple layers of the received digital information until a final layer is generated. The generated first through last layers may then constitute a workpiece 728 as described in the received digital information. The workpiece may be removed from the printer and post-processed as desired. For example, the workpiece may be machined from a build plate of a build platform, and further details or a smooth surface may be refined by machining or other methods.

[0122] Workpiece 728 manufactured according to method 600 may possess different structural properties compared to workpieces manufactured according to conventional manufacturing methods (e.g., machining, molding, and / or assembly). For example, all parts and / or features of workpiece 728 may be integral and / or monolithic. In another example, workpiece 728 may include multiple layers of molten material, wherein each layer is perpendicular to the build axis of the workpiece. In another example, workpiece 728 may include microstructural anisotropy caused by the directionality of the manufacturing process.

[0123] D. Exemplary Method of Manufacturing a Satellite

[0124] See also Figure 15 This section describes the steps of an exemplary method 800 for manufacturing a satellite. Aspects of the satellite, panel, and / or additive manufacturing methods and apparatus described above can be used in the following method steps. Where appropriate, references are made to components and systems that can be used to perform various steps. These references are for illustrative purposes only and are not intended to limit the possible ways of performing any particular step of the method.

[0125] Figure 15 is a flow chart showing steps performed in an exemplary method and may not list the complete process or all steps of the method. Figure 15 Various steps of method 800 are depicted in FIG. 8 , but these steps need not necessarily be performed in their entirety and, in some cases, may be performed simultaneously or in an order different from that shown.

[0126] At step 810, the method includes printing a wall panel. In some examples, the wall panel may be square or rectangular and generally flat. The wall panel may be configured as a primary load-bearing element or primary structure. Printing the wall panel may include printing standard structural features based on a standard panel design, such as those described below in optional sub-steps 812-820. Printing the wall panel may also include printing custom features incorporated into the panel design to accommodate specific structural or functional requirements of the panel. For example, step 810 may include printing a star tracker observation window, antenna mounting brackets, O-ring recesses, MMOD point shielding, and / or electrical connector interfaces.

[0127] Optional sub-steps 812-816 may be performed to print an integral sandwich panel. Such a panel may be particularly suitable for portions of a satellite that benefit from passive heat removal and low weight but have minimal radiation shielding. Sub-step 812 includes printing a first cover panel, and sub-step 814 includes printing a second cover panel. The cover panels may also be described as skins and / or walls. The first cover panel may be spaced apart from and parallel to the second cover panel. Each cover panel may have a generally flat surface and a limited thickness. The thickness of the cover panels may be equal, may vary, and / or may vary across the cover panel.

[0128] Each cover panel may have an inner side and an outer side, wherein each of the inner side and the outer side extends parallel to the build direction during printing. Substeps 812, 814 may also be described as printing the cover panels perpendicularly, and / or the cover panels may be described as being parallel to the build direction during printing.

[0129] Sub-step 816 includes printing a truss structure. The truss can be connected to the first cladding panel and can be connected to the second cladding panel, thereby connecting the first cladding panel to the second cladding panel. The truss can be described as being sandwiched between the first and second cladding panels and / or forming the core of the wall panel.

[0130] The truss may include a plurality of elongated and / or cross-shaped members arranged in an array of core structures. Each core structure may include a geometric frame such as a pyramidal frame or a cubic frame. The plurality of elongated members may be arranged so that the members serve as structural trusses and transfer loads between the first and second cover sheets. The individual members may be configured and / or oriented so that the truss can be printed without auxiliary supports. That is, the truss can be printed without subsequently removing sacrificial support structures.

[0131] The trusses can be printed simultaneously with the first and second cover panels. That is, substeps 812-816 can be performed simultaneously. The individual material layers deposited during printing can include a portion of the first cover panel, a portion of the trusses, and a portion of the second cover panel. The first and second cover panels and the trusses can be printed as a single, integral structure. That is, the cover panels and trusses can be printed together to form a printed material panel without joints or seams.

[0132] Optional sub-step 818 may be performed to print isogrid panels. Such panels may be particularly suitable for portions of a satellite that would benefit from greater radiation shielding. Sub-step 818 includes printing the isogrid including the cover panels and stiffening structure.

[0133] The cover panel may include a first side and a second side. In some examples, the first side may include an outer side or exterior side, and the second side may include an inner side or interior side. The cover panel may be printed at an angle between approximately 30 and 45 degrees to the build direction. The thickness of the cover panel may be selected based on the desired radiation shielding or blocking properties of the wall panel.

[0134] The stiffening structure may be printed on the first side of the cover sheet. The stiffening structure may comprise a regular, repeating grid or lattice extending from the cover sheet and / or may comprise local shapes and features. At least a portion of the stiffening structure may form a grid of equilateral triangles.

[0135] In some examples, printing the stiffening structure may include printing a web portion and a flange portion. The web portion may extend from the sheathing panel, and the flange portion may extend from a distal end of the web portion. The stiffening structure may be described as a T-beam or, with the sheathing panel, form an I-beam. The size of the stiffening structure (e.g., the width of the flange portion) may be selected based on the desired stiffness of the wall panel.

[0136] The stiffening structure can be printed simultaneously with the cover sheet. Each layer of material deposited during printing can include a portion of the cover sheet and a portion of the stiffening structure. The cover sheet and stiffening structure can be printed as a single, integral structure. In other words, the cover sheet and stiffening structure can be printed together to form a printed material panel without joints or seams.

[0137] Optional sub-step 820 may be performed to print a panel including a connection structure for mounting the satellite to an external structure (e.g., a launch vehicle). Sub-step 820 includes printing a connection structure configured to connect the panel to a separation system. The connection structure may be configured for the selected separation system. For example, when a belt separation system is selected, the connection structure may include an interface ring having an annular mounting surface. The connection structure may also include features designed to efficiently transfer loads between the satellite and the separation system (e.g., an I-beam structure). Printing such a panel may also include printing a deep cross-section, preferably with a hollow interior. This configuration may improve resistance to bending loads while minimizing the increase in panel weight.

[0138] Optional step 822 of method 800 includes printing multiple wall panels. This step may include sequentially repeating step 810 and / or printing multiple panels simultaneously. For example, six panels may be printed in a single printer during three depositions. A first panel may be printed according to step 810 and optional sub-step 818, and then a second panel may be printed according to step 810 and optional sub-step 820. Four additional panels may be printed in a single deposition, each printed according to step 810 and optional sub-steps 812-816.

[0139] Step 824 of the method comprises assembling the satellite's exterior wall structure, which includes the wall panels printed in step 810 and any wall panels printed in optional step 822. The wall panels may be assembled in any effective manner to form the satellite's primary structure. For example, the wall panels may be fastened together directly or indirectly. For another example, the wall panels may be bonded, welded, or otherwise secured together. The assembled wall structure may include only the wall panels, may include the wall panels and additively manufactured components, and / or may include the wall panels and one or more conventionally manufactured components.

[0140] At step 826, the method includes mounting a communication device to the exterior wall structure. The communication device may include an antenna (e.g., a cup dipole antenna, a command horn antenna, or an array antenna) and / or other elements of a communication system (e.g., a processor, an encoder, a modulator, a transmitter, or a receiver). The communication device may be configured to facilitate communication between a satellite and a planetary body and / or another spacecraft in space. In some examples, the communication device may use radio waves, microwaves, and / or any other electromagnetic radiation to transmit and / or receive data.

[0141] The communication device can be mounted to the exterior wall structure in any effective manner. For example, the antenna can be directly fastened to the wall panel, can be printed as an integral structure of the wall panel, and / or can be affixed to a satellite deployable mechanism. The communication device can be mounted to the exterior of the exterior wall structure or can be mounted in an interior compartment enclosed by the exterior wall structure.

[0142] Step 828 includes installing the exowall structure in a launch vehicle. A launch vehicle may include any vehicle or system configured to transport a satellite from a planetary body into space. For example, a launch vehicle may include a rocket configured to propel a satellite from the Earth's surface through the Kármán line into space. The satellite may be described as the payload of the launch vehicle.

[0143] The exostructure can be mounted to a launch vehicle with a payload adapter and / or separation system. For example, a belt-type separation system such as a clamp, motorized light belt (MLB), KSRC, or mechanical locking system (MLS), a dispenser-type separation system such as a Quadpack or Canisterized Satellite Dispenser (CSD), and / or a pyrotechnic separation system can be used.

[0144] Illustrative Combinations and Additional Examples

[0145] This section describes additional aspects and features of additively manufactured satellites, which are presented as a series of paragraphs (without limitation), some or all of which may be designated alphanumerically for clarity and efficiency. Each of these paragraphs may be combined in any suitable manner with one or more other paragraphs and / or with the disclosure elsewhere in this application. Some of the following paragraphs explicitly reference and further qualify other paragraphs, thereby providing (without limitation) examples of some suitable combinations.

[0146] A0. A satellite comprising:

[0147] a body having an additively manufactured outer wall structure at least partially forming an enclosed compartment, and

[0148] A communication device, attached to the body, is configured to receive and transmit data in space.

[0149] A1. A satellite according to A0, wherein the outer wall structure includes a plurality of side panels, each of which is additively manufactured by laser sintering aluminum alloy.

[0150] A2. A satellite according to A1, wherein each side panel radiates heat from the interior of the compartment at a rate of at least 5 watts per square foot.

[0151] A3. A satellite according to A1, wherein each side panel radiates heat from the interior of the compartment at a rate between approximately 20 and 40 watts per square foot.

[0152] A4. A satellite according to A1, wherein each side panel radiates heat from the interior of the compartment at a rate of at least 45 watts per square foot.

[0153] A5. A satellite according to any one of A0 to A4, wherein the outer wall structure includes a plurality of side panels, each side panel being integral.

[0154] A6. A satellite according to A5, wherein each side panel includes first and second cover panels connected by an internal truss structure.

[0155] A7. A satellite according to any one of A0 to A6, further comprising a propulsion system configured to propel the satellite in space.

[0156] A8. A satellite according to A7, wherein the propulsion system includes an additively manufactured propellant tank.

[0157] A9. A satellite according to any one of A0 to A8, wherein the wall structure comprises additively manufactured isogrid panels.

[0158] A10. A satellite according to A9, wherein the isogrid panels include a cover panel composed of a laser sintered aluminum alloy and having a thickness sufficient to block at least 0.5 Mrad / year of radiation originating from outside the compartment.

[0159] A11. A satellite according to A10, wherein the thickness is sufficient to block at least 1 Mrad / year of radiation.

[0160] A12. A satellite according to A10 or A11, wherein the thickness is sufficient to block at least 100 Mrad / year of radiation.

[0161] A13. A satellite according to any one of A10 to A12, wherein the thickness is sufficient to block at least 500 Mrad / year of radiation.

[0162] A14. A satellite according to any one of A9 to A13, wherein the isogrid panel includes a cover panel and an isogrid stiffening structure on the outside of the satellite body.

[0163] A15. A satellite according to A14, wherein the isogrid stiffening structure comprises an I-beam structure forming an array of equilateral triangles.

[0164] A16. A satellite according to any one of A0 to A15, wherein the communication device includes an additively manufactured antenna mounted inside the compartment.

[0165] A17. A satellite according to any one of A0 to A16, wherein the communication device includes an additively manufactured antenna mounted inside the compartment.

[0166] A18. A satellite according to any one of A0 to A17, wherein the entire outer wall structure is additively manufactured by laser sintering aluminum alloy.

[0167] A19. A satellite according to A18, wherein the outer wall structure is cubical and includes four side panels, a front panel and a rear panel.

[0168] A20. A satellite according to any one of A0 to A19, wherein the outer wall structure includes an additively manufactured rear panel connected to a separation system, which is configured to install and carry the body inside the launch vehicle during the launch phase and release the body from the launch vehicle after the launch phase.

[0169] A21. A satellite according to A20, wherein the additively manufactured rear panel is hollow.

[0170] A22. A satellite according to A20 or A21, wherein the rear panel has one or more openings aligned with the thruster brackets of the satellite.

[0171] A23. A satellite according to A22, wherein one or more thruster brackets are composed of laser sintered titanium.

[0172] B0. A satellite comprising:

[0173] a housing having an additively manufactured outer wall structure,

[0174] a communication device connected to the housing, configured to receive and transmit data in space, and

[0175] A separation device, connected to the wall structure, is configured to mount and carry the shell inside the launch vehicle during a launch phase and then release the shell from the launch vehicle after the launch phase.

[0176] B1. A satellite according to B0, wherein the outer wall structure includes additively manufactured side panels configured to radiate heat from the interior of the shell.

[0177] B2. The satellite of B1, wherein each side panel comprises an integral sandwich structure.

[0178] B3. A satellite according to any one of B0 to B2, wherein the outer wall structure includes panels configured to block space radiation from entering the shell.

[0179] B4. A satellite according to any one of B0 to B3, wherein the outer wall structure includes panels connected to the separation device.

[0180] C0. A spacecraft launch assembly, the spacecraft launch assembly comprising:

[0181] rocket-powered vehicles,

[0182] a payload comprising an additively manufactured wall structure, and

[0183] A separation device connects the wall structure to the rocket-powered vehicle and is configured to install and carry the payload inside the rocket-powered vehicle during a launch phase and then release the payload from the rocket-powered vehicle after the launch phase.

[0184] C1. The spacecraft launch assembly of C0, wherein the payload comprises a satellite.

[0185] C2. A spacecraft launch assembly according to C0 or C1, wherein the satellite includes an additively manufactured antenna.

[0186] D0. A method for manufacturing a satellite, the method comprising:

[0187] printing wall panels configured to form part of an outer wall structure of the satellite,

[0188] Assembling the satellite's outer wall structure including the wall panels,

[0189] A communication device is mounted to the exterior wall structure, the communication device being configured to receive and transmit data in space.

[0190] D1. The method according to D0, wherein the printing step comprises:

[0191] A plurality of wall panels are printed that are configured to form an entire exterior wall structure of the satellite.

[0192] D2. The method of D1, wherein the plurality of wall panels are printed simultaneously.

[0193] D3. The method of D2, wherein the plurality of wall panels are printed on a single printer.

[0194] D4. The method according to any one of D0 to D3, wherein the panel is composed of a laser sintered aluminum alloy.

[0195] D5. The method according to any one of D0 to D4, wherein the wall panel is a unitary sandwich structure.

[0196] D6. The method according to any one of D0 to D5, wherein the printing step comprises:

[0197] Print the first cover sheet,

[0198] Print the second cover sheet, and

[0199] Print the truss structure connecting the first and second cladding panels.

[0200] D7. The method according to any one of D0 to D6, wherein the printing step comprises:

[0201] Prints a uniform grid including stiffening structures on the cladding panels.

[0202] D8. The method according to any one of D0 to D7, wherein the printing step comprises:

[0203] A structure is printed on the wall panel, the structure being configured for connection to the separation device.

[0204] D9. The method according to D, further comprising:

[0205] Install the outer wall structure inside the launch vehicle.

[0206] E0. A satellite comprising:

[0207] an outer skin structure comprising wall panels comprising a laser sintered metal alloy, and

[0208] A communication device inside the skin structure is configured to send and receive data in space.

[0209] E1. A satellite according to E0, wherein the wall panels are composed of laser sintered aluminum alloy.

[0210] E2. A satellite according to E0 or E1, wherein the wall panels are composed of laser sintered titanium alloy.

[0211] E3. A satellite according to any one of E0 to E2, wherein the communication device includes an additively manufactured antenna.

[0212] E4. A satellite according to any one of E0 to E3, wherein the outer skin structure is entirely composed of a laser sintered metal alloy.

[0213] E5. A satellite according to any one of E0 to E4, wherein the wall panel comprises an integral sandwich structure.

[0214] Advantages, Features, and Benefits

[0215] The various examples of additively manufactured satellites described herein provide a number of advantages over known solutions for manufacturing satellites. For example, the illustrative examples described herein allow for rapid, on-demand production of satellites.

[0216] Additionally, the illustrative examples described herein reduce production cycle time, including reducing the number of manufacturing steps, contact labor time, and post-production testing, among other benefits.

[0217] Additionally, the illustrative examples described herein allow for design agility and rapid customization of standard designs to specific payloads, among other benefits.

[0218] Additionally, among other benefits, the illustrative examples described herein reduce cost and part count by integrating previously separately manufactured components and post-processed features into the additive manufacturing of a satellite's primary structure.

[0219] Additionally, the illustrative examples described herein allow for a high level of customization and local design, among other benefits.

[0220] Additionally, among other benefits, the illustrative examples described herein allow some or all major structural components to be printed simultaneously in a single deposition.

[0221] Additionally, the illustrative examples described herein allow for printing without sacrificing auxiliary supports, among other benefits.

[0222] There is no known system or device that can perform these functions, especially for a high precision, robust and lightweight satellite primary structure.However, not all examples described herein offer the same advantages or the same degree of advantages.

[0223] in conclusion

[0224] The disclosure set forth above may encompass a plurality of different examples with independent practicality. Although each of these has been disclosed in its preferred form, the specific examples disclosed and illustrated herein should not be considered in a limiting sense, because numerous changes may be made. With respect to the use of section headings within this disclosure, these headings are only used for organizational purposes. The subject matter of the present disclosure includes all novel and non-obvious combinations and sub-combinations of the various elements, features, functions and / or properties disclosed herein. The accompanying claims specifically point out certain combinations and sub-combinations that are considered novel and non-obvious. Other combinations and sub-combinations of features, functions, elements and / or properties may be claimed in applications claiming priority to this application or a related application. Regardless of whether these claims are wider, narrower, equal or different in scope than the original claims, these claims are also considered to be included in the subject matter of the present disclosure.

Claims

1. A satellite, comprising: a body having an additively manufactured outer wall structure at least partially forming an enclosed compartment; propellant tanks; as well as a communication device attached to the subject, the communication device being configured to receive and transmit data in space, wherein: The outer wall structure includes an additively manufactured rear panel connected to a separation system configured to mount and carry the body inside a launch vehicle during a launch phase and to release the body from the launch vehicle after the launch phase, The additively manufactured rear panel includes an inner skin, an outer skin and side walls, The additively manufactured aft panel includes a central interface ring and a plurality of thruster support interface rings, the central interface ring including an inner annular surface of the inner skin and an outer annular surface of the outer skin, and The central interface ring is connected to the propellant tank and the separation system, wherein the propellant tank is mounted to the inner annular surface and the separation system is mounted to the outer annular surface.

2. The satellite according to claim 1, wherein: The outer wall structure includes a plurality of side panels, and each side panel is integral.

3. The satellite according to claim 1, wherein: The outer wall structure includes a plurality of side panels, each of which is manufactured by laser sintering aluminum alloy.

4. The satellite according to claim 3, wherein: Each side panel radiates heat from the interior of the compartment at a rate of at least 5 watts per square foot.

5. The satellite according to claim 1, wherein The outer wall structure includes additively manufactured isogrid panels.

6. The satellite according to claim 5, wherein: The isogrid panel includes a cover sheet constructed of a laser sintered aluminum alloy and having a thickness sufficient to block at least 0.5 Mrad / year of radiation originating from outside the compartment.

7. The satellite according to claim 1, wherein: The entire outer wall structure is manufactured by laser sintering aluminum alloy.

8. The satellite according to claim 7, wherein: The outer wall structure is in a cubic shape and includes four side panels, a front panel and a rear panel.

9. The satellite according to claim 1, wherein: The rear panel has one or more holes aligned with the satellite's thruster mounts.

10. A method for manufacturing a satellite, the method comprising the following steps: a printing step of printing a wall panel configured to form a portion of an outer wall structure of the satellite, an assembling step of assembling the outer wall structure of the satellite including the wall panels, a mounting step of mounting a communication device to the outer wall structure, wherein the communication device is configured to receive and transmit data in space; The outer wall structure includes an additively manufactured rear panel connected to a separation system configured to mount and carry the body inside a launch vehicle during a launch phase and to release the body from the launch vehicle after the launch phase, The additively manufactured rear panel includes an inner skin, an outer skin and side walls, The additively manufactured aft panel includes a central interface ring and a plurality of thruster support interface rings, the central interface ring including an inner annular surface of the inner skin and an outer annular surface of the outer skin, and The central interface ring is connected to a propellant tank of the satellite and to the separation system, wherein the propellant tank is mounted to the inner annular surface and the separation system is mounted to the outer annular surface.

11. The method according to claim 10, wherein: The printing step comprises: A plurality of wall panels configured to form an entire exterior wall structure of the satellite are printed.

12. The method according to claim 11, wherein The plurality of wall panels are printed simultaneously.

13. The method according to claim 12, wherein: The plurality of wall panels are printed on a single printer.

14. The method according to claim 10, wherein: The wall panels are composed of laser sintered aluminum alloy.

Images