Systems for reducing reflected sunlight from space vehicles

Abstract

A space vehicle includes a space vehicle body and a radiator, the space vehicle body including an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, wherein at least a portion of the space vehicle body includes a panel including a sheet metal; the radiator being positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations configured such that, when the space vehicle is evaluated under certain conditions, the corrugations reflect sufficient sunlight away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0.1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle lacking the plurality of corrugations and otherwise equivalent to the space vehicle, when evaluated under the same conditions

Description

SYSTEMS FOR REDUCING REFLECTED SUNLIGHT FROM SPACE VEHICLESCross-Reference

[0001] This application is an international (PCT) patent application relating to and claiming the benefit of commonly-owned, co-pending France Patent Application No. 2309115, filed on August 30, 2023, and entitled “SYSTEMS FOR REDUCING REFLECTED SUNLIGHT FROM SPACE VEHICLES,” the contents of which are incorporated herein by reference in their entirety.Field of the Invention

[0002] The field of the invention relates to space vehicle thermal management systems and methods for manufacturing such vehicles and systems. More particularly, the field invention relates to space vehicles, such as satellites, having radiators that reduce reflection of sunlight toward Earth.Background of the Invention

[0003] Conventional radiators used in space vehicles, such as satellites, have flat surfaces. Such radiators are often covered by coating materials having desired thermo-optical properties.Brief Description of the Figures

[0004] Some embodiments of the invention are herein described, by way of example only, with reference to the accompanying drawings. With specific reference now to the drawings in detail, it is stressed that the particulars shown are by way of example and for purposes of illustrative discussion of embodiments of the invention. In this regard, the description taken with the drawings makes apparent to those skilled in the art how embodiments of the invention may be practiced.

[0005] Figure 1 shows a perspective view of an exemplary space vehicle.

[0006] Figure 2 shows a perspective view of an exemplary space vehicle.

[0007] Figure 3 shows a perspective view of a panel of an exemplary space vehicle.

[0008] Figure 4 shows an exploded perspective view of panels of an exemplary space vehicle.

[0009] Figure 5 shows an exploded perspective view of panels of an exemplary space vehicle.

[0010] Figure 6 shows elements of an exemplary space vehicle.

[0011] Figure 7 shows elements of an exemplary space vehicle including stamped grooves and heat pipes.

[0012] Figure 8A shows a perspective view of elements of an exemplary space vehicle including stamped grooves that are shaped to receive heat pipes.

[0013] Figure 8B shows a top view of the elements of the exemplary space vehicle shown in Figure 8A.

[0014] Figure 9 shows a cross-section of an exemplary radiator corrugation.

[0015] Figure 10 shows a cross-section of an exemplary radiator corrugation.

[0016] Figure 11 shows a cross-section of an exemplary panel corrugation.

[0017] Figure 12 shows a perspective view of an exemplary corrugated radiator

[0018] Figure 13A shows a perspective view of an exemplary space vehicle including corrugated radiators.

[0019] Figure 13B shows a detailed perspective view of a portion of the space vehicle of Figure 13A.

[0020] Figure 14A shows an exemplary corrugation.

[0021] Figure 14B shows an exemplary corrugation.

[0022] Figure 15A shows sunlight incident on a flat reflective surface.

[0023] Figure 15B shows sunlight incident on an exemplary corrugated radiator.

[0024] Figure 15C shows an alternate view of sunlight incident on the exemplary corrugated radiator of Figure 15B.

[0025] Figure 16A shows an exemplary patern of an exemplary corrugated radiator.

[0026] Figure 16B shows an exemplary patern of an exemplary corrugated radiator.

[0027] Figure 17A shows a diagram showing parameters relevant to evaluation of satellite magnitude.

[0028] Figure 17B shows a detailed view of a portion of the diagram of Figure 17A.

[0029] Figure 17C shows a detailed view of a portion of the diagram of Figure 17A.

[0030] Figure 18A shows reflection of sunlight from an exemplary satellite at a first location.

[0031] Figure 18B shows reflection of sunlight from an exemplary satellite at a second location.

[0032] Figure 18C shows reflection of sunlight from an exemplary satellite at a third location.

[0033] Figure 19A shows a detailed view of a portion of the diagram of Figure 17A.

[0034] Figure 19B shows a graph characterizing reduction of reflected sun flux provided by a satellite including a corrugated radiator.Summary of the Invention

[0035] In some embodiments, a space vehicle includes an upper panel; a lower panel fixed to the upper panel around at least a portion of respective perimeters thereof to thereby form an enclosed structure having an interior volume; two stacking pillars at opposite sides of the enclosed structure; a stiffener extending across the interior volume, wherein the stiffener is fixed to the upper panel and the lower panel, and wherein the stiffener extends from a first end proximate a first one of the two stacking rings to a second end proximate a second one of the two stacking rings; and a plurality of payload elements, wherein each of the plurality of payload elements is fixed to at least one of the upper panel, the lower panel, or the stiffener, and wherein at least some of the plurality of payload elements lack a casing.

[0036] In some embodiments, the upper panel and the lower panel are fabricated by pressforming.

[0037] In some embodiments, a space vehicle includes a space vehicle body, wherein the space vehicle body includes an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein at least a portion of the space vehicle body includes a panel including a sheet metal; an onboard equipment component positioned within the interior of the space vehicle body; a radiator positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations, and wherein the plurality of corrugations are angled so as to reflect at least a portion of incident sunlight away from Earth; and a heat transport element positioned along the interior surface of the space vehicle body so as to convey heat away from the onboard equipment component and toward the radiator.

[0038] In some embodiments, the heat transport element is oriented transversely to the plurality of corrugations of the radiator.

[0039] In some embodiments, the heat transport element is oriented along the plurality of corrugations of the radiator.

[0040] In some embodiments, the heat transport element is positioned within one of the plurality of corrugations of the radiator.

[0041] In some embodiments, the heat transport element is defined by a material of the radiator and a material of the panel underlying the material of the radiator. In some embodiments, the heat transport element also includes a heat-conductive liquid positioned within a space defined between the material of the radiator and the material of the panel. In some embodiments, the heat-conductive liquid has a thermal conductivity greater than 0.2 W / m K at standard temperature and pressure. In some embodiments, the heat-conductive liquid includes ammonia.

[0042] In some embodiments, the heat transport element includes a heat pipe. In some embodiments, the heat pipe includes at least one of a copper-water heat pipe or an aluminumammonia heat pipe.

[0043] In some embodiments, the heat transport element includes a carbon nanotube-based material.

[0044] In some embodiments, the heat transport element has a heat flux density that is in a range of from 2 W / m2to 20 W / m2.

[0045] In some embodiments, the heat transport element has a thermal conductivity that is at least 1,000 W / m / K.

[0046] In some embodiments, the heat transport element has a heat transport capacity that is a range of from 10 W to 1,000 W.

[0047] In some embodiments, the space vehicle also includes a conductive filler positioned between the heat transport element and the interior surface of the space vehicle body so as to adhere the heat transport element to the interior surface of the space vehicle body. In some embodiments, the conductive filler has a thermal conductivity that is in a range of from 100 W / m / K to 1,000 W / m / K in plane.

[0048] In some embodiments, the plurality of corrugations are further angled so as to reduce a proportion of incident sunlight that shines on the radiator.

[0049] In some embodiments, the onboard equipment component is one of a platform equipment component or a payload component.

[0050] In some embodiments, a space vehicle includes a space vehicle body and a radiator, wherein the space vehicle body includes an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein at least a portion of the space vehicle body includes a panel including a sheet metal; and wherein the radiator is positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations, wherein the plurality of corrugations are configured such that, when the space vehicle is evaluated under evaluation conditions of: (1) sunlight illuminating Earth in a manner of an equinox date, (2) the space vehicle in an equatorial plane, and (3) the space vehicle positioned at a boundary of a dark region and a sunlight region, the plurality of corrugations reflect a sufficient portion of the sunlight incident on the space vehicle away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0.1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle, wherein the comparison space vehicle lacks the plurality of corrugations and is otherwise equivalent to the space vehicle, and wherein the reflected sunlight flux of the comparison spacevehicle is determined by evaluation of the comparison space vehicle under the evaluation conditions.

[0051] In some embodiments, the reduction factor is in a range of from 0.14 to 0.18.

[0052] In some embodiments, the plurality of corrugations includes a right triangular corrugation. In some embodiments, the right triangular corrugation includes a right triangle having a short side that is positioned along the exterior surface of the space vehicle body such that a long side of the right triangle and a hypotenuse of the right triangle are exposed to space surrounding the space vehicle. In some embodiments, an angle between the short side and the hypotenuse is in a range of from 45 degrees to 70 degrees. In some embodiments, the angle is in a range of from 55 degrees to 65 degrees.

[0053] In some embodiments, the space vehicle also includes a heat transport element positioned along the interior surface of the space vehicle body so as to convey heat away from an onboard equipment component and toward the radiator. In some embodiments, the heat transport element is positioned within one of the plurality of corrugations of the radiator.

[0054] In some embodiments, a space vehicle includes a space vehicle body and a radiator, wherein the space vehicle body includes an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein at least a portion of the space vehicle body includes a panel including a sheet metal; and wherein the radiator is positioned on the exterior surface of the space vehicle body, wherein the radiator includes a plurality of corrugations, wherein the plurality of corrugations are configured such that, when the space vehicle is evaluated under evaluation conditions of: (1) sunlight illuminating Earth in a manner of an equinox date, (2) the space vehicle in an equatorial plane, and (3) the space vehicle positioned at a boundary of a dark region and a sunlight region, the plurality of corrugations reflect a sufficient portion of the sunlight incident on the space vehicle away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0.1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle, wherein the comparison space vehicle lacks the plurality of corrugations and is otherwise equivalent to the space vehicle, and wherein the reflected sunlight flux of the comparison space vehicle is determined by evaluation of the comparison space vehicle under the evaluation conditions, and wherein the plurality of corrugations provide an increase in effective surfacearea by a factor that is in a range of from 2 to 4 as compared to a surface area of the comparison space vehicle.

[0055] In some embodiments, the reduction factor is in a range of from 0.14 to 0.18.

[0056] In some embodiments, the plurality of corrugations includes a right triangular corrugation. In some embodiments, the right triangular corrugation includes a right triangle having a short side that is positioned along the exterior surface of the space vehicle body such that a long side of the right triangle and a hypotenuse of the right triangle are exposed to space surrounding the space vehicle. In some embodiments, an angle between the short side and the hypotenuse is in a range of from 45 degrees to 70 degrees. In some embodiments, the angle is in a range of from 55 degrees to 65 degrees.

[0057] In some embodiments, the space vehicle also includes a heat transport element positioned along the interior surface of the space vehicle body so as to convey heat away from an onboard equipment component and toward the radiator. In some embodiments, the heat transport element is positioned within one of the plurality of corrugations of the radiator.

[0058] In some embodiments, the factor is in a range of from 3 to 4.Detailed Description of the Invention

[0059] The following description provides exemplary embodiments only, and is not intended to limit the scope, applicability, or configuration of the disclosure. Rather, the following description of the exemplary embodiments will provide those skilled in the art with an enabling description for implementing one or more exemplary embodiments. It will be understood that various changes may be made in the function and arrangement of elements without departing from the spirit and scope of the presently disclosed embodiments. Embodiment examples are described as follows with reference to the figures. Identical, similar, or identically acting elements in the various figures are identified with identical reference numbers and a repeated description of these elements is omitted in part to avoid redundancies.

[0060] Throughout the specification, the following terms take the meanings explicitly associated herein, unless the context clearly dictates otherwise. The phrases “in one embodiment” and “in some embodiments” as used herein do not necessarily refer to the sameembodiment(s), though they may. Furthermore, the phrases “in another embodiment” and “in some other embodiments” as used herein do not necessarily refer to a different embodiment, although they may. Thus, as described below, various embodiments may be readily combined, without departing from the scope or spirit of the present disclosure.

[0061] In addition, the term “based on” is not exclusive and allows for being based on additional factors not described, unless the context clearly dictates otherwise. In addition, throughout the specification, the meaning of “a,” “an,” and “the” include plural references. The meaning of “in” includes “in” and “on.”

[0062] As used herein, the terms “and” and “or” may be used interchangeably to refer to a set of items in both the conjunctive and disjunctive in order to encompass the full description of combinations and alternatives of the items. By way of example, a set of items may be listed with the disjunctive “or,” or with the conjunction “and.” In either case, the set is to be interpreted as meaning each of the items singularly as alternatives, as well as any combination of the listed items.

[0063] Disjunctive language such as the phrase “at least one of X, Y, or Z,” unless specifically stated otherwise, is otherwise understood with the context as used in general to present that an item, term, etc., may be either X, Y, or Z, or any combination thereof (e.g., X, Y, and / or Z). Thus, such disjunctive language is not generally intended to, and should not, imply that certain embodiments require at least one of X, at least one of Y, or at least one of Z to each be present.

[0064] The exemplary embodiments described herein relate to radiators for space vehicles, such as satellites. More particularly, the exemplary embodiments described herein relate to corrugated radiators for space vehicles. The exemplary embodiments also relate to space vehicles having corrugated radiators directly overlaying heat transport elements, such as heat pipes.

[0065] Figures 1 and 2 show perspective views of opposite sides of an exemplary space vehicle 100. In some embodiments, an exemplary space vehicle 100 includes two panels 110, 120 forming an outer shell of the space vehicle 100. In some embodiments, the first and second panels 110, 120 are not honeycomb panels. In some embodiments, the first and second panels 110, 120 includes a metal. In some embodiments, the panels 110, 120 include a metal alloy.In some embodiments, the metal alloy includes at least one of an aluminum alloy, a steel alloy, and / or combinations thereof. In some embodiments, the panels 110, 120 include an aluminum alloy. In some embodiments, the aluminum alloy includes aluminum and zinc. In some embodiments, the aluminum alloy is a 7000-series aluminum alloy. In some embodiments, the aluminum alloy is 7075 aluminum. In some embodiments, the panels 110, 120 are press- formed. In some embodiments, the panels 110, 120 are press-formed from sheet aluminum. In some embodiments, the metal alloy includes an alloy other than steel and / or aluminum. In some embodiments, the panels 110, 120 include a composite. In some embodiments, the panels 110, 120 include a fiber composite. In some embodiments, the fiber composite includes at least one of a carbon fiber composite and / or a glass fiber composite.

[0066] In some embodiments, the sheet aluminum has a thickness that is from 0.1 millimeter to 10 millimeters. In some embodiments, the sheet aluminum has a thickness that is from 0.1 millimeter to 1 millimeter. In some embodiments, the sheet aluminum has a thickness that is from 1 millimeter to 3 millimeters. In some embodiments, the sheet aluminum has a thickness that is from 3 millimeters to 5 millimeters. In some embodiments, the sheet aluminum has a thickness that is from 5 millimeters to 10 millimeters. In some embodiments, the sheet aluminum is 1 millimeter in thickness. In some embodiments, the sheet aluminum is 2 millimeters in thickness. In some embodiments, the sheet aluminum has a first thickness (e.g., 0.1 millimeter) in some regions and has one or more regions having a greater second thickness (e.g., 2 millimeters) provided by the addition of a corresponding thickening panel or panels.

[0067] In some embodiments, the panels 110, 120 are joined to one another around the respective perimeters thereof to form an enclosed structure. In some embodiments, the panels 110, 120 are connected by a mechanical joining technique such as welding, joining with fasteners (e.g., bolts, rivets, etc.), or adhering (e.g., with a glue or other adhesive). Figure 3 shows an exploded perspective view of the exemplary panel 110. Figure 4 shows an exploded perspective view of the exemplary panels 110 and 120, together with an earth deck panel 130 that is configured to be positioned at a location of the space vehicle 100 that faces toward Earth when the space vehicle 100 is in orbit. Figure 5 shows an additional perspective view of the panels 110 and 120 and the earth deck panel 130.

[0068] Referring back to Figures 1 and 2, in some embodiments, the exterior of the space vehicle 100 includes one or more stacking pillars 140, a solar array mounting interface 150, and one or more antenna modules 160 (e.g., transmitting antenna modules and receiving antenna modules). In some embodiments, the one or more stacking pillars 140 includes two stacking pillars 140. In some embodiments, respective pass-throughs are formed in the panels 110, 120 to allow external elements of the space vehicle 100 to be coupled (e.g., communicatively) to internal elements of the space vehicle 100.

[0069] In some embodiments, an exemplary space vehicle 100 includes onboard equipment, such as platform equipment and payloads. Figure 6 shows an exemplary space vehicle 100 including various onboard equipment 600, such as platform equipment and payloads, positioned at various locations therein. In some embodiments, exemplary onboard equipment includes an on-board computer, a power conditioning and distribution unit, an on-board processor, an optical terminal (such as an optical head), one or more reaction wheels, a power processing unit, one or more gateway interface units, and / or a battery. The specific equipment listed above is only exemplary, and the equipment that will be included in the exemplary space vehicle 100 will vary among different implementations. Additionally, the specific locations of the onboard equipment 600 shown in Figure 6 are only exemplary, and the specific locations of onboard equipment 600 will vary among different implementations. In some embodiments, at least some items of the onboard equipment 600 have their respective casings removed and are mounted within the exemplary space vehicle by mounting directly to one of the panels.

[0070] In some embodiments, an exemplary space vehicle includes an upper panel and a lower panel that are attached to one another to form an enclosed body within which platform hardware and payloads are accommodated. In some embodiments, the interior of an exemplary space vehicle (e.g., an interior space as defined between two joined panels) is pressurized. In some embodiments, the exterior of the exemplary space vehicle (e.g., the perimeter of the panels 110, 120 described above) is hermetically sealed in order to allow the interior to be pressurized. In some embodiments the interior is pressurized to a pressure that is in a range of from 1 bar to 2 bar. In some embodiments, the interior is pressurized with a gas. In some embodiments, the gas is a convection gas. In some embodiments, the convection gas facilitates convection of heat within the space vehicle.

[0071] In some embodiments, the panels 110, 120 include grooves formed in surfaces that are configured to face the interior of the space vehicle 100. In some embodiments, the grooves 700 are configured (e.g., sized and shaped) to receive heat transport elements, such as heat pipes, therein. In some embodiments, the grooves 700 are sized and shaped to receive heat pipes that have a l l mm to 15 mm square cross-section. In some embodiments, grooves 700 are also formed in the earth deck panel. Figure 7 shows an exemplary panel 710 and earth deck panel 720 with grooves 700 formed therein. In some embodiments, such as shown in Figure 7, the grooves 700 are oriented linearly, such as in parallel lines.

[0072] In some embodiments, heat transport elements, such as heat pipes, are embedded in at least some of the grooves 700. In the embodiment shown in Figure 7, heat transport elements 740 are embedded in some of the grooves 700. In some embodiments, the heat transport elements 740 include copper- water heat pipes. In some embodiments, the heat transport elements include a carbon nanotube-based material. In some embodiments, the heat transport elements 740 include aluminum-ammonia heat pipes such as the heat pipes commercialized by Advanced Cooling Technologies of Lancaster, Pennsylvania, or by Euro Heat Pipes of Nivelles, Belgium. For clarity, Figure 7 includes a reference numeral identifying a single one of the heat transport elements 740, but this term and the accompanying description refers to other exemplary heat pipes as well. In some embodiments, a heat spreader, such as a heat spreader made from carbon nanotubes, is used in place of the heat transport elements 740.

[0073] In some embodiments, the heat transport elements 740 are secured in the grooves 700 using a conductive fdler. In some embodiments, a suitable conductive fdler provides sufficient thermal conductivity to convey heat from a payload component located adjacent to one of the grooves 700 to one of the heat transport elements 740, and also provides sufficient adhesion to provide sufficient thermal conductivity while compensating for some degree of surface roughness either within the grooves 700 and / or on the surface of the heat transport elements 740. In some embodiments, the conductive filler has a thermal conductivity that is in a range of from 1 W / m / K to 30 W / m / K cross-plane (e.g., in a direction across the conductive filler between one of the heat transport elements 740 and the surface of a surrounding one of the grooves 700). In some embodiments, the conductive filler has a thermal conductivity that is in a range of from 100 W / m / K to 1,000 W / m / K in plane (e.g., in a direction along the conductive filler perpendicular to the cross-plane direction). In some embodiments, the conductive fillerhas sufficient adhesion to compensate for surface roughness that is in a range of from 0 microns (e.g., a perfectly smooth surface with no pores, gaps, etc.) to 500 microns. In some embodiments, suitable conductive fillers include the filler commercialized under the trade name MAPSIL by Map Space Coatings of Mazeres, France, the filler commercialized under the trade name SIGRAFLEX by SGL Carbon of Charlotte, North Carolina, the filler commercialized under the trade name T-PLI by Laird Technologies of Chesterfield, Missouri, and the filler commercialized under the trade name THERM-A-GAP by Parker Chomerics of Woburn, Massachusetts.

[0074] In some embodiments, the panels of an exemplary space vehicle include equipment fixation locations, which are locations to which equipment (e.g., platform elements and payloads) may be secured. Figures 8A and 8B shows a perspective view and a top view, respectively of an embodiment of an exemplary panel 800 having instrument fixation locations 810 and grooves 700 (only one of each of which is specifically identified in Figures 8A and 8B for clarity). In some embodiments, the grooves 700 are positioned proximate to (e.g., extending away from) the payload fixation locations 810 such that heat transport elements received therein will be thermally coupled to payload elements secured at the payload fixation locations 810, thereby to conduct heat away from such payload elements. In some embodiments, the space vehicle includes one or more radiators configured to radiate heat from the space vehicle into space. In some embodiments, the grooves 700 extend from the payload fixation locations 810 and toward a location from which heat is to be radiated, such as a radiator.

[0075] In some embodiments, an exemplary thermal control system includes corrugated surfaces. In some embodiments, an exemplary thermal control system includes corrugated radiators. In some embodiments, an exemplary thermal control system includes corrugated radiators positioned directly adjacent to heat transport elements to thereby allow heat within the heat transport elements to be directly conducted to the corrugated radiators. In some embodiments, an exemplary thermal control system includes corrugated radiators that are integrally formed with heat transport elements (e.g., heat pipes are positioned within corrugations of a corrugated radiator and / or heat transport elements are at least partially formed from the material of a corrugated radiator).

[0076] Figures 9-11 show exemplary embodiments of thermal control arrangements. Figures 9-11 include only a small portion of an actual implementation of a thermal control system, in that each figure shows a cross-sectional view of only one corrugation of a system that will include multiple corrugations. In some embodiments, an actual implementation including the thermal control arrangements shown in Figures 9-11 will include several of the corrugations adjacent to one another. Additionally, while Figures 9-11 show cross-sectional views, exemplary corrugations implemented into an exemplary space vehicle will have some depth in the view direction of the cross-sectional views shown in Figures 9-11. Figure 12 shows a perspective view of an exemplary corrugated arrangement 1200 including five corrugations, shown in a manner so as to represent the depth of the corrugated arrangement 1200.

[0077] Figure 9 shows an exemplary embodiment of a thermal control arrangement 900. The thermal control arrangement 900 includes a generally linear heat transport element 910. In some embodiments, the heat transport element 910 is a copper-water heat pipe. In some embodiments, the heat transport element 910 includes a carbon nanotube-based material. In some embodiments, the heat transport element 910 is an ammonium-aluminum heat pipe. In some embodiments, the heat transport element 910 is an oscillating heat pipe (“OHP”). The thermal control arrangement 900 also includes a corrugated radiator 920. In some embodiments, the corrugated radiator 920 is formed by press-forming as described above. In some embodiments, the corrugated radiator 920 is formed from a metal. In some embodiments, the corrugated radiator 920 is formed from an aluminum alloy. In some embodiments, the aluminum alloy is a 2000 series aluminum alloy, 6000 series aluminum alloy or a 7000 series aluminum alloy. In the embodiment shown in Figure 9, the heat transport element 910 is oriented transversely to the corrugations of the corrugated radiator 920.

[0078] Figure 10 shows an exemplary embodiment of a thermal control arrangement 1000. The thermal control arrangement 1000 includes a corrugated heat transport element 1010. In some embodiments, the heat transport element 1010 is a copper-water heat pipe. In some embodiments, the heat transport element 1010 includes a carbon nanotube -based material. In some embodiments, the heat transport element 1010 is an ammonium-aluminum heat pipe. In some embodiments, the heat transport element 1010 is an oscillating heat pipe (“OHP”). The thermal control arrangement 1000 also includes a corrugated radiator 1020. In some embodiments, the corrugated radiator 1020 is formed by press-forming as described above. Insome embodiments, the corrugated radiator 1020 is formed from a metal. In some embodiments, the corrugated radiator 1020 is formed from an aluminum alloy. In some embodiments, the aluminum alloy is a 2000 series aluminum alloy, 6000 series aluminum alloy or a 7000 series aluminum alloy.

[0079] Figure 11 shows an exemplary embodiment of a thermal control arrangement 1100. The thermal control arrangement 1100 includes a corrugated panel portion 1110 and an flat panel portion 1120 that is positioned adjacent to and underlying the corrugated panel portion 1110. In some embodiments, the corrugated panel portion 1110 and the flat panel portion 1120 are formed by press-forming as described above. In some embodiments, a gap 1130 is formed between the corrugated panel portion 1110 and the flat panel portion 1120. In some embodiments, the gap 1130 is fdled with a heat-conductive liquid that is arranged so as to act as an OHP, thereby forming a heat transport element from the corrugated panel portion 1110, the flat panel portion 1120, and the heat-conductive liquid. In some embodiments, the heat- conductive liquid is a liquid having athermal conductivity greater than 0.2 W / m K at standard temperature and pressure. In some embodiments, the heat-conductive liquid is ammonia. In some embodiments, the corrugated panel portion 1110 is formed from an aluminum alloy. In some embodiments, the corrugated panel portion 1110 is formed from a metal. In some embodiments, the corrugated panel portion 1110 is formed from an aluminum alloy. In some embodiments, the aluminum alloy is a 2000 series aluminum alloy, 6000 series aluminum alloy or a 7000 series aluminum alloy.

[0080] Figure 13 A shows a perspective view of an exemplary space vehicle 1300 having corrugated radiator regions 1310 and 1320. Figure 13B shows an alternate perspective view of the corrugated radiator region 1310. As described above with reference to Figures 9-11, in some embodiments, the corrugated radiator regions 1310 and 1320 overlay heat transport elements so as to draw heat from such heat transport elements and radiate such heat into space.

[0081] In some embodiments, the corrugations of a corrugated radiator as described herein are optimized to increase the heat dissipation capability of a radiator. In some embodiments, the use of a corrugated surfaces increases the equivalent heat dissipation area of a radiator. Figure 14A shows a cross-sectional view of an exemplary corrugation pattern 1400 illustrating the increase in surface area of a corrugated radiator to thereby increase the heat distribution. Thecorrugation pattern 1400 shown in Figure 14A is defined by a footprint length 1410, a height 1420, and an angle 1430. The footprint length 1410, height 1420, and angle 1430 combine to produce corrugations having a first side length 1440 and a second side length 1450, the sum of which is the effective length of the corrugation pattern 1400. In some embodiments, for a given footprint length 1410, the effective length can be tuned by selecting the height 1420 and the angle 1430. In some embodiments, for given values of the footprint length 1410 and the height 1420, the angle 1430 can be selected to increase the effective surface area (e.g., the sum of the first side length 1440 and the second side length 1450) by a ratio of 1 + ^2 (i.e., about 2.41) as compared to the footprint length 1410. In some embodiments, such an increase in effective surface area thereby increases the heat dissipation capability of such a radiator by the same ratio as compared to a planar radiator of the same size.

[0082] Figure 14B shows a cross-sectional view of an exemplary corrugation pattern 1460. The corrugation pattern 14B includes a series of right triangular corrugations defined by a footprint length 1470 (e.g., a short side of the triangle positioned along the exterior surface of the space vehicle), a height 1480, and an angle 1490. In some embodiments, the angle 1490 is between 45 degrees and 70 degrees. In some embodiments, the angle is between 55 degrees and 65 degrees. In some embodiments, the angle 1490 is 60 degrees. In embodiments in which the angle 1490 is 60 degrees, the corrugation pattern has a surface length of 2 + A / 3 (e.g., about 3.73, which is the length of the long side of the triangle plus the length of the hypotenuse of the triangle, both of which are exposed to space), thereby increasing the effective surface area by a ratio of 2 + A / 3 (e.g., about 3.73) as compared to the footprint length 1470. In some embodiments, such an increase in effective surface area thereby increases the heat dissipation capability of such a radiator by the same ratio as compared to a planar radiator of the same size. In some embodiments, a corrugated radiator provides an increase in effective surface area by a factor that is in a range of from 2 to 4, or that is in a range of from 2 to 3.5, or that is in a range of from 2 to 3, or that is in a range of from 2 to 2.5, or that is a in a range of from 2.5 to 4, or that is a range of from 2.5 to 3.5, or that is in a range of from 2.5 to 3, or that is in a range of from 3 to 4, or that is in a range of from 3 to 3.5, or that is in a range of from 3.5 to 4.

[0083] In some embodiments, a corrugated radiator is optimized to increase heat dissipation capability as described above with reference to Figures 14A and 14B, and also to limit the brightness of a satellite as viewed from an observer on the ground. Figure 15A shows reflectionof sunlight off a flat surface. As shown in Figure 15 A, incident sunlight 1500 reflects off a flat surface 1510 to produce reflected sunlight 1520 that projects toward an observer on Earth. Figure 15B shows reflection of the same incident sunlight 1500 off a corrugated surface (e.g., a surface having corrugations similar to the corrugation pattern 1400 described above with reference to Figure 14A. As shown in Figure 15B, a surface includes corrugations 1530 similar to the corrugation pattern 1400 described above with reference to Figure 14A. As a result of the presence of the corrugations 1530, the incident sunlight 1500 reflects off the corrugations 1530 to produce reflected sunlight 1550 that projects away from Earth; no reflected sunlight projects in a direction 1555 toward Earth that is the same direction as the direction of the reflected sunlight 1520 shown in Figure 15 A.

[0084] In addition to reflecting sunlight away from an observer on Earth, in some embodiments, a corrugated radiator reduces the proportion of a radiator surface on which the sun reflects. Figure 15C shows additional detail of the manner in which sunlight is incident upon a surface including the corrugations 1530 shown in Figure 15B. Figure 15C shows the incident sunlight 1500 projecting in the same direction as shown in Figures 15A and 15B, and also shows additional parallel rays 1502, 1504, 1506. As shown in Figure 15C, the rays 1502, 1504, and 1506 are defined by respective peaks 1532, 1534, 1536 of the corrugations 1530. Each of the peaks 1532, 1534, 1536 obstructs a portion of its own corrugation as well as a portion of a subsequent corrugation such that the sunlight only shines on a portion of each corrugation. For example, referring to a corrugation 1560 that includes the peak 1534, the peak 1534 obstructs the sunlight from shining on a shaded portion 1562 of the corrugation 1560 as well as on a shaded portion 1564 of an adjacent corrugation, leaving only an exposed portion 1566 of the adjacent corrugation exposed to the sunlight. As such, the portion of the corrugations 1530 reflecting the sun is reduced as compared to, for example, a flat surface 1510 as described above with reference to Figure 15A. As a result, in some embodiments, the corrugations 1530 reduce solar flux received by a radiator including the corrugations 1530.

[0085] In some embodiments, a corrugated radiator includes corrugations have a shielding angle that is selected to allow a white painted Earth deck with low reflection to the ground or a black coated Earth deck with low incidence of solar flux for a worst hot case beta angle. In some embodiments, the shielding angle is in a range of from 20 degrees to 40 degrees, or is in a range of from 25 degrees to 35 degrees, or is about 30 degrees, or is 30 degrees.

[0086] Typically, observed brightness of an object in space, such as a satellite, is expressed quantitatively in terms of apparent magnitude. Magnitude of an object can be measured by the measurement technique of photometry, which measures light in the ultraviolet, visible, and / or infrared spectrums. Magnitude is a logarithmic scale using the brightness of the star Vega as a reference value of zero, under which brighter objects have lower values. According to the magnitude scale, the sun as seen from Earth has a magnitude of about -27, the full moon as seen from Earth at perihelion has a magnitude of -12.9, the new moon as seen from Earth has a magnitude of -2.5, the faintest stars visible in an urban neighborhood with the naked eye have magnitude of +3 to +4, and the faintest starts observable by a typical naked eye under very good conditions have a magnitude of about +6.5. Accordingly, the contribution of any given object, such as a satellite, to light pollution can be evaluated based on its apparent magnitude, such as based on its apparent magnitude under typical operating conditions (e.g., object position and orientation; angle of incident sunlight; location of observer). Similarly, the relative brightness of two objects (and, thus, the relative contribution to such objects to light pollution) can be evaluated by comparing the apparent magnitude of the two objects under the same conditions (e.g., at the same position and orientation, as exposed to the same incident sunlight, and as observed from the same location).

[0087] In some embodiments, a space vehicle (e.g., the space vehicle 1300) including a radiator having corrugations (e.g., the corrugations in the corrugated radiator regions 1310 and 1320) appears less bright to an observer than a comparable space vehicle that lacks corrugations but is otherwise equivalent. In some embodiments, the effect of corrugations (e.g., the corrugations in the corrugated radiator regions 1310 and 1320) on brightness can be expressed in terms of the relationship between the apparent magnitude of a space vehicle including the corrugations (e.g., the space vehicle 1300) and the apparent magnitude of a comparable space vehicle that lacks corrugations but is otherwise equivalent. In some embodiments, the effect of corrugations (e.g., the corrugations in the corrugated radiator regions 1310 and 1320) on brightness can be expressed solely in terms of the apparent magnitude of a space vehicle including the corrugations (e.g., the space vehicle 1300) without reference to a comparable space vehicle that lacks corrugations.

[0088] In some embodiments, magnitude of a space vehicle can be determined (e.g., computationally or by observation) at a single set of conditions, which may be, for example,an arbitrary set of conditions, a “worst-case” set of conditions, or a set of conditions selected as representative of a range of conditions. In some embodiments, magnitude of a space vehicle can be determined by determining (e.g., computationally or by observation) magnitude at various conditions across a range of conditions (e.g., for a variety of satellite positions and for a variety of observer locations) and selecting the worst-case magnitude (e.g., the lowest magnitude value, which is indicative of the highest brightness) as the magnitude of the space vehicle.

[0089] In some embodiments, a satellite is possibly visible by an observer on the Earth when the following conditions are met: (1) the sky is dark at the observer’s location (e.g., it is night where the observer is located), (2) the satellite is positioned at a location where it is visible by the observer (e.g., the satellite is not positioned such that the Earth obstructs line of sight between the satellite and the observer), (3) the satellite is illuminated by the sun (e.g., the satellite is not positioned such that the Earth obstructs sunlight from illuminating the satellite), and (4) the satellite is oriented such that at least one surface of the satellite reflects incident sunlight toward the Earth.

[0090] Figure 17A shows a diagram 1700 illustrating these conditions. It should be noted that Figure 17A is a two-dimensional representation of three-dimensional space, but the general concepts described with reference to Figure 17A remain applicable to three-dimensional analysis. In the diagram 1700, the Earth 1705 is viewed from above the North Pole, and incident sunlight 1710 illuminates the Earth as shown, producing a sunlight region 1715 and a dark region 1720. An arc 1725 shows the region on the surface of the Earth where the sky is dark, and wherein the above condition (1) is accordingly met for any location within the arc 1725 as a result. An example observer location 1730 within the arc 1725 is indicated, and a half circle 1735 indicates the region of the sky that is visible to an observer in the observer location 1730. In other words, the half circle 1735 indicates the possible locations where the above condition (2) is met for an observer in the observer location 1730. A circle 1740 indicates the set of locations at which a satellite crosses the equatorial plane. The circle 1740 indicates a satellite orbiting at an altitude of about 1,200 kilometers; a similar circle will exist for satellites orbiting at different altitudes. Within the circle 1740, an arc 1745 indicates the region of the orbit in which a satellite is both (a) within the sunlight region 1715 (e.g., the abovecondition (3) is met) and (b) visible to an observer in the observer location 1730 (e.g., the above condition (2) is met).

[0091] Continuing to refer to Figure 17A, while a satellite is positioned within the arc 1745, it may be positioned and oriented such that at least one surface of the satellite reflects incident sunlight toward the Earth (e.g., the above condition (4) is met), or may be positioned and oriented such that no surface of the satellite reflects incident sunlight toward the earth (e.g., the above condition (4) is not met). Figure 17A shows a first position 1750 within the arc 1745 and a second position 1755 within the arc 1745. Figures 17B and 17C show detailed views of a satellite 1760, with Figure 17B showing the satellite 1760 oriented as positioned in the first position 1750 and Figure 17C showing the satellite 1760 oriented as positioned in the second position 1755. As shown in Figures 17B and 17C, the satellite 1760 includes a nadir / earth deck surface 1765, a zenith surface 1770, and a side surface 1775.

[0092] Referring to Figure 17B, when the satellite 1760 is positioned at the first position 1750, the incident sunlight 1710 is incident on the side surface 1775 and the zenith surface 1770. As shown in Figure 17B, sunlight incident on the zenith surface 1770 is reflected in a direction 1780a away from the Earth, while sunlight incident on the side surface 1775 is reflected in a direction 1785 toward the Earth, the direction 1785 is towards a region of the Earth that is within the sunlight region 1715. Thus, when the satellite 1760 is positioned in the first position 1750, the above condition (4) is not met. Referring now to Figure 17C, when the satellite 1760 is positioned at the second position 1755, the incident sunlight 1710 is incident on the side surface 1775 and the nadir surface 1765. As shown in Figure 17C, sunlight incident on the side surface 1775 is reflected in a direction 1780b away from the Earth, while sunlight incident on the nadir surface 1765 is reflected toward the Earth, and is reflected in a direction 1790 towards a region of the Earth that is within the dark region 1720. Thus, when the satellite 1760 is positioned in the second position 1750, the above condition (4) is met. Referring back to Figure 17A, an arc 1795 shows the portion of the arc 1745 within which the above condition (4) is met. Summarizing the above, condition (1) is met for observer locations within the arc 1725; condition (3) is met for satellite locations within the arc 1745, condition (4) is met for satellite locations within the arc 1795, and condition (2) is or is not met depending on the specific locations of the observer and the satellite.

[0093] In some embodiments, the apparent brightness of a space vehicle, such as a satellite, is characterized in terms of reflected sun flux, e.g., the quantity or proportion of sunlight incident on the space vehicle that is reflected toward the Earth. In some embodiments, the reflected sun flux will depend on the shape, position, and orientation of the space vehicle. In some embodiments, to compare the apparent brightness of one space vehicle to that of another, the reflected sun flux of the two space vehicles are compared under identical conditions (e.g., comparing the two space vehicles when in the same position and being exposed to the same incident sunlight). In some embodiments, to compare the apparent brightness of one space vehicle to that of another, the reflected sun flux of the two space vehicles are compared under a “worst case” set of conditions, i.e., conditions that would be expected to result in maximum reflection of incident sunlight toward the Earth.

[0094] In some embodiments, a space vehicle including corrugations such as the corrugation pattern 1460 described above with reference to Figure 14B will reflect sunlight toward the Earth in a manner that it would contribute to light pollution while it is positioned within the arc 1795 discussed above with reference to Figure 17A. This is the case because, if the satellite were to be positioned within the portion of the arc 1745 that extends past the arc 1795, any reflected sunlight would reflect toward the portion of the Earth that is illuminated by sunlight (e.g., condition (1) described above would not be met).

[0095] Figures 18A, 18B, and 18C show the path of incident and reflected sunlight for a satellite including a corrugated radiator having the corrugation pattern 1460 at respective positions 1800, 1810, and 1820 within the arc 1795 (see Figure 17A). Referring first to Figure 18A, which shows the satellite 1802 when the satellite is positioned at position 1800, the only reflection is the side surface 1804. Accordingly, as the side surface 1804 faces in a direction parallel to the orbital motion of the satellite 1802 (e.g., perpendicular to the direction from the satellite 1802 toward the Earth), no light reflects toward Earth.

[0096] Referring now to Figure 18B, which shows the satellite 1812 when the satellite is positioned in the position 1810, sunlight is incident on a first portion 1814 of each of the corrugations (the first portion 1814 only illustrated on one corrugation for clarity). As may be seen in Figure 18B, because of the orientation of the first portion 1814, reflected sunlight 1816 that reflects from the first portion 1814 does not reflect towards the Earth, but, rather, travelstoward a second portion 1818 of each of the corrugations. At least a portion of the reflected sunlight 1816 reflects off the second portion 1818 and becomes re-reflected sunlight 1819, some of which may be directed toward the Earth. However, the amount of the re-reflected sunlight 1819 (e.g., a reflected sun flux) is significantly less than would be re-reflected towards the Earth in the case of a planar radiator.

[0097] Referring now to Figure 18C, which shows the satellite 1822 when the satellite is positioned in the position 1820, as it the case in Figure 18B, sunlight is incident on a first portion 1824 of each of the corrugations (the first portion 1824 only illustrated on one corrugation for clarity). As may be seen in Figure 18C, because of the orientation of the first portion 1824, reflected sunlight 1826 that reflects from the first portion 1824 does not reflect towards the Earth, but, rather, travels toward a second portion 1828 of each of the corrugations. At least a portion of the reflected sunlight 1826 reflects off the second portion 1828 and becomes re-reflected sunlight 1829, some of which may be directed toward the Earth. As is the case for Figure 18C, the amount of the re-reflected sunlight 1829 (e.g., a reflected sun flux) is significantly less than would be re-reflected towards the Earth in the case of a planar radiator.

[0098] Because the size of the first portion 1824 is larger than the size of the first portion 1814, the amount of the re-reflected sunlight 1829 that may travel towards the Earth when the satellite is in the position 1820 is larger than the amount of the re-reflected sunlight 1819 that may travel towards the Earth when the satellite is in the position 1810. Thus, by comparison of Figures 18A, 18B, and 18C, it may be concluded that the position 1820 is a suitable “worst case” position for comparing different space vehicles to one another to evaluate relative brightness. To summarize, the position 1820 can be characterized as a position (1) in an equatorial plane, (2) at a boundary between a dark region and a sunlight region, and (3) the Earth oriented as on an equinox date.

[0099] For a given satellite configuration, orientation, and position, the amount of incident sunlight that is eventually reflected toward the Earth (e.g., the reflected sun flux) can be determined on the basis of geometry, the law of reflection, and Lambert’s Law. Figure 19A shows a detailed view of a portion of Figure 17A including the arc 1795. As shown in Figure 19A, the position of a satellite within the arc 1795 can be expressed in terms of an angle 1900 (with the position line shown for the angle 1900 in Figure 19A indicating a satellite at theposition 1810). Figure 19A also shows the location 1910 of an observer on the Earth that is directly below the position 1820, which presents a likely “worst case” observation location.

[0100] Figure 19B shows a graph 1950 characterizing the reduction in reflected sun flux provided by a satellite having corrugated radiators as described above with reference to Figure 14B, including having an angle 1490 of 60 degrees. The horizontal axis 1960 of the graph 1950 corresponds to the position of the satellite along the arc 1795 as characterized by the value of the angle 1900 shown in Figure 19A. The graph 1950 includes a data plot 1970 against a vertical axis 1980. The data plot 1970 is indicative of the ratio of the reflected sun flux provided by the satellite having corrugated radiators as described above to the reflected sun flux provided by a satellite that includes non-corrugated (e.g., planar) radiators and is otherwise equivalent to the satellite described above (e.g., has the same geometry, is formed from the same materials, includes the same coatings, etc.). For example, at an angle of 20 degrees as measured along the horizontal axis 1960, the data plot 1970 indicates a value of 0.09 along the vertical axis. This indicates that, when a satellite having corrugated radiators as described above with reference to Figure 14B and having a corrugation angle of 60 degrees is positioned at a position along the arc 1795 having an angle 1900 of 20 degrees, as described above with reference to Figure 19A, the reflected sun flux towards the Earth will be 0.09 times the reflected sun flux towards the Earth from a satellite that lacks corrugated radiators but is otherwise equivalent.

[0101] As discussed above, the position 1820 shown in Figures 17A and 19A provides a likely “worst case” location for reflected sun flux toward the Earth, irrespective of satellite design. In some embodiments, a space vehicle (e.g., the space vehicle 1300) having a corrugated radiator (e.g., including the corrugations in the corrugated radiator regions 1310 and 1320, such as the corrugations shown in Figures 14A and 14B), when evaluated as described above, exhibits reflected sun flux that is less than the reflected sun flux exhibited by a space vehicle that lacks a corrugated radiator but is otherwise comparable (e.g., includes substantially the same geometry, substantially the same materials, substantially the same coatings, etc.) by a factor that is in a range of from 0. 1 to 0.2, or is in a range of from 0. 12 to 0.2, or is in a range of from 0.14 to 0.2, or is in a range of from 0.16 to 0.2, or is in a range of from 0.18 to 0.2, or is in a range of from 0.1 to 0.18, or is in a range of from 0.12 to 0.18, or is in a range of from 0.14 to 0.18, or is in a range of from 0.16 to 0.18, or is in a range of from 0.1 to 0.16, or is in arange of from 0.12 to 0.16, or is in a range of from 0.14 to 0.16, or is in a range of from 0. 1 to 0.14, or is in a range of from 0.12 to 0. 14, or is in a range of from 0. 1 to 0. 12.

[0102] In some embodiments, a space vehicle (e.g., the space vehicle 1300) having a corrugated radiator (e.g., including the corrugations in the corrugated radiator regions 1310 and 1320, such as the corrugations shown in Figures 14A and 14B), when evaluated as described above, exhibits reflected sun flux that is less than the reflected sun flux exhibited by a space vehicle that lacks a corrugated radiator but is otherwise comparable (e.g., includes substantially the same geometry, substantially the same materials, substantially the same coatings, etc.), and also provides increased effective surface area for heat dissipation, wherein the reduction in reflected sun flux is by a factor that is in a range of from 0.1 to 0.2, or is in a range of from 0.12 to 0.2, or is in a range of from 0.14 to 0.2, or is in a range of from 0.16 to 0.2, or is in a range of from 0.18 to 0.2, or is in a range of from 0.1 to 0.18, or is in a range of from 0.12 to 0.18, or is in a range of from 0.14 to 0.18, or is in a range of from 0.16 to 0.18, or is in a range of from 0.1 to 0.16, or is in a range of from 0.12 to 0.16, or is in a range of from 0.14 to 0.16, or is in a range of from 0.1 to 0. 14, or is in a range of from 0.12 to 0. 14, or is in a range of from 0.1 to 0.12, and wherein the increase in effective surface area is by a factor that is in a range of from 2 to 4, or that is in a range of from 2 to 3.5, or that is in a range of from 2 to 3, or that is in a range of from 2 to 2.5, or that is a in a range of from 2.5 to 4, or that is a range of from 2.5 to 3.5, or that is in a range of from 2.5 to 3, or that is in a range of from 3 to 4, or that is in a range of from 3 to 3.5, or that is in a range of from 3.5 to 4.

[0103] In some embodiments, such as described above with reference to Figures 12, 13A, and 13B, a corrugated radiator includes a linear corrugation pattern (e.g., a pattern including a series of parallel corrugations). In other embodiments, a corrugated radiator includes a two- dimensional corrugation pattern. Figure 16A shows a top view of an embodiment of a corrugated radiator 1600 having a square pattern defined by alternating rising portions 1610, 1612 (e.g., analogous to the portions identified with reference numeral 1440 in Figure 14A) and descending portions 1620, 1622 (e.g., analogous to the portions identified with reference numeral 1450 in Figure 14A). Figure 16B shows a top view of a corrugated radiator 1650 having a circular pattern defined by alternating rising portions 1660, 1662 (e.g., analogous to the portions identified with reference numeral 1440 in Figure 14A) and descending portions 1670, 1672 (e.g., analogous to the portions identified with reference numeral 1450 in Figure14A). The specific paterns of the corrugated radiators 1600 and 1650 shown in Figures 16A and 16B, respectively, are only exemplary, and other embodiments of corrugated radiators may have differing two-dimensional paterns without departing from the broader principles described herein.

[0104] In some embodiments, any of the heat transport elements described herein (e.g. the heat transport elements 740, 910, and 1010) has a heat flux density that is in the range of from 2 W / m2to 20 W / m2, or is in the range of from 2 W / m2to 10 W / m2, or is in the range of from 5 W / m2to 20 W / m2, or is in the range of from 5 W / m2to 10 W / m2. In some embodiments, any of the heat transport elements described herein (e.g. the heat transport elements 740, 910, and 1010) has a thermal conductivity that is greater than 1,000 W / m / K, or is greater than 2,000 W / m / K, or is greater than 3,000 W / m / K, or is between 1,000 W / m / K and 100,000 W / m / K, or is between 2,000 W / m / K and 100,000 W / m / K, or is between 3,000 W / m / K and 100,000 W / m / K. In some embodiments, any of the heat transport elements described herein (e.g. the heat transport elements 740, 910, and 1010) has a heat transport capacity that is in the range of from 10 W to 1,000 W, or is in the range of from 50 W to 1,000 W, or is in the range of from 10 W to 500 W, or is in the range of from 50 W to 500 W, or is in the range of from 10 W to 300 W, or is in the range of from 50 W to 300 W. In some embodiments, any of the heat transport elements described herein (e.g., the heat transport elements 740, 910, and 1010) has a cross-sectional size (e.g., an exterior side length of a square heat pipe or an exterior diameter of a circular heat pipe) that is in the range of from 5 mm to 50 mm, and can have a cross- sectional size selected based on the amount of heat required to be conveyed thereby. In some embodiments, any of the heat transport elements described herein (e.g., the heat transport elements 740, 910, and 1010) is linear, L-shaped, U-shaped, or has a more complex shape, in order to allow for such heat transport elements to be positioned and to convey heat from and to appropriate locations within a satellite structure (e.g., from a payload component to a radiator).

[0105] In some embodiments, an exemplary space vehicle is made by a process including providing a quantity of sheet aluminum; forming at least a first panel and a second panel from the sheet aluminum by press-forming, wherein the first panel includes a groove formed therein; positioning a heat transport element in the groove; forming a radiator overlaying the heat transport element; providing at least one onboard equipment component, wherein the at least one onboard equipment component has no outer casing; fixing the at least one onboardequipment component to the first panel so as to overlay the heat transport element; and sealing the first panel to the second panel to form the space vehicle.

[0106] By incorporating grooves for receiving heat transport elements directly into the material of the panels (e.g., by forming such grooves by stamping the material of the panels), heat transport elements may be incorporated where appropriate and without inclusion of additional elements to accommodate the heat transport elements. Additionally, by positioning corrugated radiators directly over heat transport elements, radiation of heat generated by onboard equipment into space is improved.

[0107] While a number of embodiments of the present invention have been described, it is understood that these embodiments are illustrative only, and not restrictive, and that many modifications may become apparent to those of ordinary skill in the art. For example, all dimensions discussed herein are provided as examples only, and are intended to be illustrative and not restrictive.

Claims

ClaimsWhat is claimed is:

1. A space vehicle, comprising: a space vehicle body, wherein the space vehicle body comprises an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein at least a portion of the space vehicle body comprises a panel comprising a sheet metal; and a radiator positioned on the exterior surface of the space vehicle body, wherein the radiator comprises a plurality of corrugations, wherein the plurality of corrugations are configured such that, when the space vehicle is evaluated under evaluation conditions of:(1) sunlight illuminating Earth in a manner of an equinox date,(2) the space vehicle in an equatorial plane, and(3) the space vehicle positioned at a boundary of a dark region and a sunlight region, the plurality of corrugations reflect a sufficient portion of the sunlight incident on the space vehicle away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0. 1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle, wherein the comparison space vehicle lacks the plurality of corrugations and is otherwise equivalent to the space vehicle, and wherein the reflected sunlight flux of the comparison space vehicle is determined by evaluation of the comparison space vehicle under the evaluation conditions.

2. The space vehicle of claim 1, wherein the reduction factor is in a range of from 0. 14 to 0.18.

3. The space vehicle of claim 1, wherein the plurality of corrugations comprises a right triangular corrugation.

4. The space vehicle of claim 3, wherein the right triangular corrugation comprises a right triangle having a short side that is positioned along the exterior surface of the space vehicle body such that a long side of the right triangle and a hypotenuse of the right triangle are exposed to space surrounding the space vehicle.

5. The space vehicle of claim 4, wherein an angle between the short side and the hypotenuse is in a range of from 45 degrees to 70 degrees.

6. The space vehicle of claim 5, wherein the angle is in a range of from 55 degrees to 65 degrees.

7. The space vehicle of claim 1, further comprising a heat transport element positioned along the interior surface of the space vehicle body so as to convey heat away from an onboard equipment component and toward the radiator.

8. The space vehicle of claim 7, wherein the heat transport element is positioned within one of the plurality of corrugations of the radiator.

9. A space vehicle, comprising: a space vehicle body, wherein the space vehicle body comprises an interior surface defining an interior of the space vehicle body and an exterior surface opposite the interior surface, and wherein at least a portion of the space vehicle body comprises a panel comprising a sheet metal; and a radiator positioned on the exterior surface of the space vehicle body, wherein the radiator comprises a plurality of corrugations,wherein the plurality of corrugations are configured such that, when the space vehicle is evaluated under evaluation conditions of:(1) sunlight illuminating Earth in a manner of an equinox date,(2) the space vehicle in an equatorial plane, and(3) the space vehicle positioned at a boundary of a dark region and a sunlight region, the plurality of corrugations reflect a sufficient portion of the sunlight incident on the space vehicle away from the Earth such that a reflected sunlight flux is reduced by a reduction factor that is in a range of from 0. 1 to 0.2 as compared to a reflected sunlight flux of a comparison space vehicle, wherein the comparison space vehicle lacks the plurality of corrugations and is otherwise equivalent to the space vehicle, and wherein the reflected sunlight flux of the comparison space vehicle is determined by evaluation of the comparison space vehicle under the evaluation conditions, and wherein the plurality of corrugations provide an increase in effective surface area by a factor that is in a range of from 2 to 4 as compared to a surface area of the comparison space vehicle.

10. The space vehicle of claim 9, wherein the reduction factor is in a range of from 0.14 to 0.18.

11. The space vehicle of claim 9, wherein the plurality of corrugations comprises a right triangular corrugation.

12. The space vehicle of claim 11, wherein the right triangular corrugation comprises a right triangle having a short side that is positioned along the exterior surface of the space vehicle body such that a long side of the right triangle and a hypotenuse of the right triangle are exposed to space surrounding the space vehicle.

13. The space vehicle of claim 12, wherein an angle between the short side and the hypotenuse is in a range of from 45 degrees to 70 degrees.

14. The space vehicle of claim 13, wherein the angle is in a range of from 55 degrees to 65 degrees.

15. The space vehicle of claim 9, further comprising a heat transport element positioned along the interior surface of the space vehicle body so as to convey heat away from an onboard equipment component and toward the radiator.

16. The space vehicle of claim 15, wherein the heat transport element is positioned within one of the plurality of corrugations of the radiator.

17. The space vehicle of claim 9, wherein the factor is in a range of from 3 to 4.

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