Patent Publication Number: US-2015069187-A1

Title: Hosted instrument radiator system

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims the priority benefit of U.S. Patent Application No. 61/875,578, filed on Sep. 9, 2013, the entirety of which is incorporated herein by reference. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     Not applicable. 
     FIELD 
     The present disclosure generally relates to an instrument mounting system for a satellite, and in particular to, for example, without limitation, a hosted instrument radiator system. 
     BACKGROUND 
     Accommodation of precision instruments as hosted payloads on commercial or communications satellites promises to dramatically reduce the costs associated with deploying such instrument based systems. Precision instruments may require different accommodation resources than those generally available on commercial satellites. Commercial satellite customers are reluctant to offer their satellites as instrument hosts if doing so would add significant cost, risk, or schedule impact. 
     The description provided in the background section, including without limitation, any problems, features, solutions or information, should not be assumed to be prior art merely because it is mentioned in or associated with the background section. The background section may include information that describes one or more aspects of the subject technology. 
     SUMMARY 
     The description in this summary section may provide some illustrative examples of the disclosure. This section is not intended to be a broad overview or to identify essential elements of the disclosure. 
     In accordance with an aspect of some embodiments of the inventions disclosed herein is the realization that the rejection of waste thermal energy from a hosted instrument is one of the most challenging problems associated with hosting a precision instrument. Waste heat must be removed from the instrument using a thermal sink operating at a temperature well below those typically available on commercial satellites, often below 20° C., and a separate thermal energy removal system for the hosted instrument must often be added to the host satellite. These auxiliary radiator systems are typically expensive and consume large amounts of valuable physical space on the satellite. 
     Additionally, an aspect of some embodiments is the realization that some hosted instruments are very sensitive to torques generated by satellite thermal gradients and material coefficient of thermal expansion (CTE) effects. It is sometimes necessary to minimize the mechanical coupling between spacecraft-provided waste heat radiators and the hosted payload. 
     Therefore, according to some embodiments, an instrument mounting system is provided that can comprise a thermal radiator component to provide independent heat rejection capability for an instrument that is hosted on a geosynchronous satellite. 
     The system can comprise an instrument mounting surface. In some embodiments, the instrument mounting surface can be coupled to the radiator component. The instrument mounting surface can be thermally coupled to the radiator component, for example, via heat pipes that may include one or both of rigid or flexible elements. 
     In some embodiments, the system can be mounted on the hosting satellite so as to provide the hosted instrument with a clear field of view of the Earth while rejecting heat in one or more directions, such as one or both of the north and south directions. 
     The system may also provide structural support for elements or components of one or both of the hosted instrument or the hosting satellite. 
     The system may be integrated with the instrument prior to installation of the instrument on the hosting satellite, thereby allowing the system to be tested and integration issues resolved without impact on the timelines of the hosting satellite or the launch vehicle. 
     The system may provide higher performance thermal control than that available on the hosting satellite while avoiding the need to modify the host satellite to accommodate the instrument. This may provide a greater number of possible hosting satellites. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The accompanying drawings, which are included to provide further understanding and are incorporated in and constitute a part of this specification, illustrate disclosed embodiments and together with the description serve to explain the principles of the disclosed embodiments. In the drawings: 
         FIGS. 1-2  depict a commercially hosted infrared payload mounted on a commercial GEO satellite. 
         FIGS. 3-4  depict typical Earth-deck mounting constraints imposed on a hosted instrument. 
         FIG. 5  depicts a satellite with an instrument mounting assembly in a deployed configuration, according to some embodiments. 
         FIGS. 6A-6D  depict top, front, side, and isometric views of a typical instrument to be hosted on a satellite. 
         FIG. 7A  shows an enlarged view of the mounting assembly of  FIG. 5  in the stowed configuration, according to some embodiments. 
         FIG. 7B  depicts the mounting assembly of  FIG. 7A  in a deployed configuration, according to some embodiments. 
         FIG. 8A  depicts another exemplary radiator assembly, according to some embodiments. 
         FIG. 8B  depicts an instrument mounted on the instrument mounting panel of the mounting assembly of  FIG. 8A , according to some embodiments. 
         FIG. 9A  is an enlarged side view of the mounting assembly of  FIG. 8A , according to some embodiments. 
         FIG. 9B  is a further enlargement of the heat pipe at the corner of the mounting assembly of  FIG. 9A , according to some embodiments. 
         FIGS. 10A-10C  depict an example instrument attached to an exemplary mounting assembly and mounted on a host satellite within a typical launch vehicle fairing, according to some embodiments. 
     
    
    
     DETAILED DESCRIPTION 
     The present disclosure generally relates to a radiator system provided for a secondary electronic package, such as a scientific instrument or payload, hosted on a multi-function satellite, such as a communication or commercial satellite. 
     The detailed description set forth below is intended as a description of various configurations of the subject technology and is not intended to represent the only configurations in which the subject technology may be practiced. The appended drawings are incorporated herein and constitute a part of the detailed description. The detailed description includes specific details for the purpose of providing a thorough understanding of the subject technology. However, it will be apparent to those skilled in the art that the subject technology may be practiced without these specific details. In some instances, well-known structures and components are shown in block diagram form in order to avoid obscuring the concepts of the subject technology. Like components are labeled with identical element numbers for ease of understanding. 
     In accordance with some embodiments disclosed herein is the realization that host commercial satellites present an opportunity for hosting one or more secondary instruments, sensors, or payloads. The ability to host an additional payload on a satellite can generate additional revenue for the host satellite and provide a budget-efficient opportunity for another party to have a payload carried by the host satellite without incurring the full cost the host satellite. Accordingly, in some embodiments, payloads can be placed on existing satellites and thereby enable a party to enjoy substantial cost savings. These opportunities can stimulate exciting new ventures and give companies the ability to pursue previously unreachable goals. 
     However, an aspect of at least some embodiments disclosed herein is the realization that the ability to place an additional payload on a host satellite creates substantial risk for the host satellite. In order to take advantage of such opportunities, the payload and the system used to incorporate the payload must meet strict requirements associated with the host satellite. In particular, some embodiments disclosed herein recognize the requirement that a host satellite be configured in an efficient, space-saving configuration that ensures that all components of the satellite and accompanying payload can function properly and reliably. Such satellites are built to strict specifications that are demanding and sensitive. 
     Therefore, an aspect of at least some embodiments disclosed herein is the realization that in order to enable a satellite to host an additional payload, the payload must provide mere noise or insignificant effects to the satellite and instrumentation thereof. The noise or effects shall be considered insignificant if the satellite and its instrumentation are able to perform their intended function. 
     In making this possible, some embodiments minimize the thermal distortion or vibration of the system such that the additional payload can be accommodated on the host satellite with minimal adverse effects. The additional payload may be sensitive to thermal-distortion-induced torques originating from the host satellite or the hosted payload thermal radiator system (provided by the host satellite). Thus, some embodiments advantageously avoid the need for thermal distortion reducing modifications to the host satellite and therefore reduce the impact of any necessary accommodations, thereby increasing the likelihood that a given host satellite qualifies or could be used with some embodiment disclosed herein. Indeed, as noted above, a candidate hosted instrument should be accommodated in the most simplistic and non-impactful way possible in order to maximize its chances of finding a host. 
     Some embodiments of the disclosed system provide a thermal control capability that is otherwise not available in a hosting satellite. Some embodiments of the disclosed system may also provide structural support for instrument elements, for example a deployable antenna or non-deployable sunshade. Some embodiments allow the system to integrate the instrument with a thermal control system and/or a structural support in advance of mounting the system on the hosting satellite. The instrument and system may be integrated and tested prior to installation on the hosting satellite, possibly reducing one or more of the time, schedule, risk, or cost elements of the integrated satellite program. Such advantages can decouple the mounting and assembly activities from the timeline of the hosting satellite and launch vehicle preparation, thereby taking the instrument off of a critical path schedule. This may also advantageously reduce the risk and cost of integration and timeline disruption, as problems with the instrument can be worked out without delay of the entire program. 
     Further, some embodiments of the disclosed radiator system can provide a self-contained thermal control system for an instrument to be hosted on a GEO satellite. The system can comprise a radiator component that enables the system to have one or more fields of view (“FOVs”) directed toward the Earth and/or a select lateral direction, for example north, south, east, or west, and reject heat, for example, to one or both of the northern space or southern space. 
     In current satellite configurations, satellites are most efficient if the instrumentation and existing equipment occupies the entirety of the available space; the satellite should be made as compact as possible while being as large as necessary. Accordingly, the prior art has not attempted to create an available space on a host satellite for a secondary instrument or payload, especially for a large-scale instrument or payload or in situations in which the host satellite requires large components. 
     However, in accordance with some embodiments disclosed herein, an instrument mounting system can enable large-scale payloads to be hosted on host satellites. For example, the size of the payload or instrument that can be hosted on the satellite can be of an instrument class that is large-scale, or much larger compared to the instrumentation of the satellite itself 
     Referring to  FIGS. 1-2 , a host satellite  100  can be configured to support or carry a small-scale hosted payload  110 . The satellite  100  and hosted payload  110  illustrated in  FIG. 1  demonstrates that there have been no known instances of large-scale hosted precision instruments or payloads. As used herein, “large-scale” payloads can refer to payloads that are at least about, 0.7 m 3 , 0.8 m 3 , 0.9 m 3 , 1 m 3 , 1.1 m 3 , 1.2 m 3 , or larger in size. No such large-scale hosted payload has been used on prior satellites or other implementations thereof, whether combined with large antenna or large thermal radiator systems. 
     Further, none of the prior systems, such as that illustrated in  FIG. 2 , included a large-scale payload in combination with any substantial or large antennae or radiator systems. Thus, as illustrated in  FIGS. 1-2 , prior systems have not been capable of supporting large-scale payloads, especially when the host satellite  100  required large-scale antennae, control modules, radiators, or other electronics. 
     In accordance with some embodiments disclosed herein, a system is provided that can enable a host satellite to support a larger payload or instrument simultaneously with large host satellite payload antennas or equipment. Further, some embodiments of the system can be capable of isolating radiator panel thermal distortion torques or translations from the payload or instrument. For example, some embodiments can use, a deflectable heating system, such as a flexible heat pipe system. Additionally, some embodiments can provide a system that is uniquely configured to accommodate a large-scale payload or instrument by utilizing unique radiator panel shapes that are configured to maximize capacity within launch vehicle fairing constraints. Furthermore, some embodiments of the system can enable the integration of a high-value payload or instrument with significantly higher size, mass, power, or thermal resource requirements, onto a typical commercial geosynchronous communications satellite. 
     Referring now to  FIGS. 3-4 , a satellite  120  is shown that comprises a deck area  122  that is substantially limited in size and typically used for host mission antennas and spacecraft sensors. As such, the satellite  120  has little to no space for accommodating any additional payload or instrument thereon.  FIGS. 4  is a top view of the satellite  120 . 
     As shown in  FIGS. 3-4 , the satellite  120  can comprise spacecraft primary thermal radiators  130  that have north and south faces, primary mission communication antenna feeds  132 , one or more primary mission communications antenna reflectors  134  that need a clear view to earth, satellite earth sensors  136  that need a clear view to earth, primary mission communication antennas  138  that also need a clear view to earth, and other related equipment. While excess spacecraft thermal energy is radiated to space in the north and south directions  140 ,  142 , other equipment of the satellite  120  requires a clear view to the earth. Traditionally, such constraints made it impossible to host any large-scale payloads or instruments. 
     However, in accordance with some embodiments disclosed herein, a system is provided by which an innovative solution to this conflict between the host satellite and the hosted payload is resolved. In particular, some embodiments allow the field of view, area, and thermal energy rejection requirements to the resolved such that a host satellite can carry a large-scale hosted payloads while ensuring that all of the host equipment and the payload itself are capable of operating within the field of view, area, and thermal energy rejection requirements. 
       FIG. 5  illustrates an embodiment of a system  200  that can be used with a satellite  202  in order to accommodate a large-scale payload  300  along with various host satellite components  210 . 
     As illustrated in  FIGS. 6A-6D , the payload  300  can comprise a precision instrument that has significant mounting, thermal, or field-of-view constraints that must be accommodated by the host satellite  202 . The hosted payload  300  can comprise two separate thermal dissipation rejection systems. A first thermal dissipation rejection system of the payload  300  can comprise a north or south facing self-contained instrument thermal radiator  302  to reject heat directly to space, as shown in  FIGS. 6A-6B  and  6 D. The thermal radiator  302 , when the payload  300  is positioned on the host satellite  202 , would generally require that the thermal radiator  302  have a mostly clear field of view to space. Such a clear field of view would therefore allow excess thermal energy to be radiated directly to space from the radiator  302 . The second thermal dissipation rejection system can eliminate heat through a surface  304  and a mounting panel or conductive interface  306 , as shown in FIGS.  6 A and  6 C- 6 D. The payload can also comprise an anti-Earth or Zenith direction surface  304  through which a portion of the payload&#39;s excess thermal energy can be dissipated, such as by conduction, to a host spacecraft. In some embodiments, the host spacecraft can receive conducted heat if the temperature at the interface between the host spacecraft or satellite  202  and the payload  300  is less than about 20° C. The conductive interface  306 , due to instrument electronics sensitivities, may preferably run at temperatures (at or below 20° C.), which are well below those used on the host satellite (up to 71° C. or even higher). 
     In addition, the conductive interface  306  can provide a mounting structure that enables the interface  306  to be coupled to the Zenith direction surface  304  of the payload  300 . The mounting structure of the conductive interface  306  can be configured to enable the payload  300  to be mechanically mounted or anchored onto the host satellite  202 . 
     Furthermore, the payload  300  can also comprise an instrument aperture  310 . The instrument aperture can allow the payload to perform a desired function and will require that the instrument aperture  310  have a clear view to the earth. 
     Referring to FIGS.  5  and  7 A- 7 B, in some embodiments, the system  200  can comprise a thermal radiator  400  that can be configured to carry or support components of the host satellite  202 , such as antenna, electronics, optical, or other components of the host satellite  202 . In some embodiments, the radiator  400  can enable one or more of the components  210  of the host satellite  202  to be structurally mounted onto the satellite  202  and allow these components to have satisfied the required field of view, area, and thermal heat rejection requirements. 
     An embodiment of a radiator component of the system is shown in FIGS.  5  and  7 A- 9 B. As illustrated therein, the radiator  400  can comprise a first or host instrument side  410  along which the host components  210  can be mounted or supported and a second or hosted instrument side  412 . As demonstrated, the field of view, area, and thermal heat rejection requirements can be satisfied by positioning the host equipment or components  210  along the first side  410  of the radiator  400 . Such requirements can be met and satisfied by positioning these components  210  along the first side  410  of the radiator  400  while positioning the host instrument or payload  300  along the second side  412  of the radiator  400 . 
     In some embodiments, the hosting satellite  202  may utilize the structure of the radiator  400  to mount elements, for example antenna feed horns or secondary antennae. In some embodiments, the structure of the radiator  400  may provide separate mounting surfaces, e.g, an Earth deck, into a first portion utilized by the hosting satellite and a second portion utilized by the hosted instrument. 
     As noted above, the host components  210  can comprise host satellite secondary antennas host satellite secondary antennas can comprise, for example, host satellite secondary antennas  220 , host satellite antenna feed horns  222 , and a host satellite antenna reflector  224 . In addition, the system  200  can also be configured such that the radiator  400  comprises a reflector deployment mechanism  226  that allows the reflector  224  to rotate from a stowed position (illustrated in  FIG. 7A ) to a deployed position (illustrated in  FIG. 7B ). In such an embodiment, the system  200  can comprise one or more antenna reflector launch configuration tie-down points  230  that allow the reflector  224  to be maintained in the stowed or launch configuration until released and deployed. The embodiment illustrated in  FIGS. 7A-7B  illustrates an example of various host equipment configurations and needs that can be satisfied in accordance with some embodiments of the system  200 . 
     As also illustrated in  FIGS. 7A-7B , the system can comprise a base structure  500  that can be coupled to the radiator  400  such that the radiator  400  extends from the base structure at an angle generally transverse relative to the base structure  500 . For example, in some embodiments, the radiator  400  can be oriented substantially perpendicular relative to the base structure  500 . As shown in  FIG. 5 , the base structure  500  can be coupled to a base panel or nadir deck  240  of the host satellite  202 . The base structure  500  can allow the system  200  to be coupled to the host satellite  202 . The base structure  500  can comprise a face or panel, as well as internal structural members. The base structure  500  can be oriented parallel relative to the nadir deck  240  when coupled to the host satellite. Further, some embodiments may be implemented in which the system does not comprise a base structure  500 . Instead, the system  200 , as shown in  FIGS. 8A-8B  with the hosted payload attached, would attach to the spacecraft&#39;s nadir deck  240  directly without the use of a base structure  500 . 
     The components  210  can generally be very sensitive to anything in their field of view. In particular, the components  210  generally require a clear field of view to the earth. Referring briefly to  FIG. 9B , which illustrates a side view of a portion of the system  200 , the hosted payloads  300  can have a required field of view  340  that extends away from the mounting structure  306  of the system  200 . The field of view  340  can be accommodated by the profile of the radiator  400 . For example, a left edge  342  of the field of view can extend transversely relative to the radiator  400  without intersecting the radiator  400 , as shown in  FIG. 9B . Further, a right edge  344  of the field of view  340  can extend generally parallel relative to the radiator  400  in a path that does not intersect with the radiator  400  or a plane of the radiator  400 . Accordingly, the radiator  400  can be configured such that its profile permits a free field of view of the instrument aperture  310  of the payload  300 . Additionally, the host components  210  can also benefit from the profile of the radiator  400 . The profile of the radiator  400  can be configured such that it does not interfere with the field of view of the host components  210 . For example, the field of view of the reflector  224 , similar to the field of view  340  of the payloads  300  can extend above the radiator  400  without interference from the radiator  400 . The other components, such as the horns and antennas of the equipment  210  can similarly be free of any interference with the profile of the radiator  400 . 
     Referring still to  FIGS. 7A-7B , the radiator  400  can be thermally independent of the components  210  and/or the satellite  202  and run at lower temperatures to suit the needs of the hosted payload  300 . The radiator  400  can be mechanically independent from the payload  300 . The radiator  400  can be thermally coupled to the base of the payload  300  (through the conductive interface  306 ). Therefore, the conductive interface  306  can run at nominally the same temperature as the radiator  400  (e.g., heat pipes can nominally isothermalize along their length). The temperature that the conductive interface  306  and the radiator  400  run at can be cooler (e.g., sometimes about 20° C.) than the satellite itself (e.g., sometimes about 71° C.), but warmer than the thermal radiator  302 . Therefore, in some embodiments, the radiator  400  can be thermally independent from the satellite  202 , but not the conductive interface  304 . Further, in some embodiments, while the host payload components  210  may not produce significant thermal dissipations, the radiator  400  could and may need to radiate any thermal dissipation produced by components  210 . In accordance with an aspect of at least some embodiments disclosed herein, the radiator  400  can therefore serve a dual use by radiating excess heat while also providing a structure onto which the components  210  can be mechanically mounted. The components  210  can be mounted to the radiator  400  and therefore be mechanically and thermally coupled to the radiator  400 . However, some of the components  210  can be dissipative and experience temperatures well above those seen on the radiator  400  (e.g, transmit feed horn(s) can be coupled to the radiator  400  in locations via a path with relatively high thermal impedance). 
     As discussed above with respect to  FIGS. 7A-7B , the mounting structure or interface  306 , due to instrument electronics sensitivities, may preferably run at temperatures (at or below 20° C.), which may be well below those used on the host satellite (up to 71° C. or even higher). The payload  300  may run at a lower temperature than an operating temperature of the equipment  210  of the host satellite  202 . Thus, the radiator  400  can provide the system  200  with a space efficiency advantage, allowing the host equipment or components  210  to be mounted thereon instead of an additional structure independent of the hosted payload thermal radiator accommodation system. 
       FIGS. 5 and 7B  depict an exemplary instrument mounting assembly in a deployed configuration, according to some embodiments. This example configuration provides an integrated thermal radiator and antenna mounting structure. The radiator  400  can provide structural mounting of antenna components, which conserves valuable satellite deck mounting space, and is compatible with both the host and instrument FOV constraints. As discussed further herein, heat removal may be accomplished using simple constant conductance heat pipes, if mechanical decoupling is not required. If minimization of mechanical coupling is desired, for example, to reduce to torque on instrument due to radiator and spacecraft thermal distortion, flexible heat pipes may be employed. 
       FIGS. 10A-10C  depict an example instrument attached to an exemplary mounting assembly and mounted on a host satellite within a typical launch vehicle fairing, according to some embodiments. 
     In accordance with some embodiments, the radiator  400  can have a design that is specifically configured such that the radiator  400  maximizes available space within the launch envelope. For example,  FIGS. 10A-10C  illustrate examples of a typical commercial satellite with a hosted instrument that is stowed for launch.  FIG. 10A  illustrates a launch vehicle  600  having a fairing  602  that defines a large envelope  604  within the fairing  602 . In these figures, the system  200  is illustrated mounted onto a host satellite  202 . The system  200  comprises the radiator  400 , which can comprise a perimeter profile  430  that can closely approximate an internal contour or shape of the launch envelope  604 . The launch envelope  604  defines a volumetric constraint into which the system  200 , including host components  210  and the payload  300 , must be fitted. 
     An aspect of at least some embodiments disclosed herein is the realization that although prior satellites must be fitted into the launch envelopes into which they are placed, some embodiments disclosed herein are configured such that the system  200  closely approximates the launch envelope in order to maximize radiator size while accommodating host equipment and a secondary instrument or payload. For example, some embodiments comprise a radiator  400  that is uniquely positioned within an upper portion  602  of the fairing. Further, in some embodiments, as discussed herein, the host components  210 , the radiator  400 , and the hosted payloads  300  can approximate the internal shape of the upper portion of the launch vehicle fairing  600 . Prior systems provides no related designs or solutions, which is consistent with Applicant&#39;s understanding that the problem being solved in some embodiments is a new problem and the solutions presented herein have been unnecessary to this point and are only now being solved using embodiments of the inventions disclosed herein. 
     In accordance with some embodiments disclosed herein, the radiator  400  can comprise a shape that is specifically contoured to typical launch vehicle fairing constraints and designed to maximize the dedicated instrument thermal radiator size and capacity of the radiator  400 . The radiator  400  can comprise one of a variety of unique profiles based on the required launch vehicle fairing compatibility. 
     For example, some embodiments provide a system  200  that includes a thermal radiator  400  that has a tapered or tapering perimeter  430  that can fit within a launch envelope  604  such that the perimeter  430  closely approximates an inner contour or surface of the launch envelope  604 . 
     Referring now to  FIGS. 8A-9B , the system  200  can be configured to provide a conductive heat removal capability, according to some embodiments. 
       FIG. 8A  depicts an exemplary radiator assembly, according to some embodiments. A planar instrument mounting surface is provided that, in some embodiments, may also provide attachment features to secure the mounting assembly to the host satellite. In some embodiments, the instrument mounting surface may be perpendicular to an Earth-pointing axis. A planar thermal radiator is provided at a right angle to the instrument mounting surface. In some embodiments, the plane of the radiator may be perpendicular to a north-south axis such that heat may be radiated from the radiator in one or both of a north and a south direction. 
     The instrument mounting surface and the radiator can be thermally connected by one or more heat pipes. A portion of the one or more heat pipes may be embedded in the radiator or, in some embodiments, thermally coupled to the surface of the radiator. In some embodiments, a second portion of the one or more heat pipes may be flanged and attached to instrument mounting plate using mechanical fasteners and thermal adhesive or filler material. In some embodiments, the instrument end of the heat pipes may directly embedded in an instrument mounting panel. In some embodiments the right angles in the heat pipes are typical rigid heat pipe if thermal distortion isolation is not a concern. In some embodiments the right angles in the heat pipes may be a flexible heat pipe. In some embodiments, a “loop” heat pipe may be used for a portion of the heat pipes. 
       FIG. 8B  depicts an instrument mounted on the instrument mounting panel of the mounting assembly of  FIG. 8A , according to some embodiments. In some embodiments, the radiator system provides a clear view of Earth, i.e. along the nadir or Earth pointing axis, while rejecting heat in one or both of north and south directions. 
     As shown in  FIGS. 8A-9B , the system  200  can provide heat removal using thermal radiator components. For example, the heat removal can be accomplished using conductive heat removal, in some embodiments. 
     Referring to  FIGS. 8A-8B , the system  200  comprises the mounting structure  306  that can be configured to remove or facilitate removal of heat from the payload  300 . For example, the waste heat coming out of the payload  300  can move through the system in one or more pathways. 
     For example, as noted above with regard to  FIGS. 6A-6D , a first pathway for heat removal can be through the thermal radiator  302 . The thermal radiator  302  can comprise a very low temperature radiator that can provide removal of waste heat from the most sensitive elements or components of the payload  300 . These elements of the payload  300  can be running at a temperature of about 77 Kelvin, which therefore requires very low temperature independent heat removal. 
     A second pathway can be provided for heat removal from components of the payload  300  out of baseplate of instrument and at a cooler temp than satellite. Such components can include electronics and computer support equipment, which can be mechanically isolated from the radiator  400  and require or use a radiator that operates at a lower temperature than the host spacecraft  202 . The removal of the heat through the second pathway can be accomplished using some embodiments disclosed herein. 
     Further, as illustrated in  FIGS. 8A-8B , the second pathway for heat removal can include the mounting structure  306 . The mounting structure  306  can be coupled to the zenith direction face of the payload  300 . In accordance with some embodiments, the mounting structure  306  can include one or more heat pipes. For example, the mounting structure  306  can comprise a honeycomb structure with heat pipes. In some embodiments, the heat pipes can be embedded into one or more structures of the system  200 . In another embodiment, the mounting structure  306  may be a solid, or other non-honeycomb plate, and the heat pipes would be attached directly to the bottom of the plate. The heat pipes may have flanges  324  (visible in  FIG. 9B  in the end view). The flanges  324  can be attached to the mounting structure  306  with mechanical fasteners and/or thermally conductive filler/adhesive. 
     Referring to  FIG. 9B , some embodiments can comprise the mounting structure  306  and a heat pipe assembly  310  that can be coupled to or embedded in the mounting structure  306 . Further, as illustrated in  FIG. 9B , the heat pipe assembly  310  can be coupled to or embedded in the mounting structure  306  and/or in the radiator  400 . In some embodiments, the heat pipe assembly  310  can be attached to an instrument mounting structure  306  interposed between the mounting structure  306  and the nadir deck  240  or  500  of the satellite  202 . 
     The heat pipe assembly  310  can comprise heat pipes that can be positioned about at least one surface of the payload  300  to provide conductive heat removal from the at least one surface of the payload  300 . Further, the heat pipe assembly  310  can comprise one or more constant conductance heat pipes. 
     In some embodiments, the heat pipe assembly  310  (which extends along the mounting structure  306 ) can be secured to the payload  300  using one or more mechanical fasteners. Further, the interface between the heat pipe assembly  310 , the instrument mounting structure  306 , and the payload  300  can also be filled with thermal adhesive or other filler material. 
     The heat pipe assembly  310  can be coupled to the mounting structure  306  and/or the radiator  400 , which can extend transversely relative to each other. An individual heat pipe  320  is shown in the side view of  FIG. 9B . The heat pipe assembly  310  can comprise between four to  20  individual heat pipes, and in some embodiments, ten heat pipes. Each pipe can, as seen in  FIG. 8A , have a L-shaped body that runs along a bottom of mounting structure  306  (see e.g.,  FIG. 9B ) and make a 90 degree bend (see e.g.,  FIG. 9B ), and run in the nadir direction internal to radiator panel  400 . Further, the heat pipe assembly  310  can comprise a standard rigid constant conductance heat pipe, a flexible pipe, or a loop heat pipe system. Thus, as an alternative to a flanged heat pipe design that is mechanically fastened to the radiator panel, the heat pipe assembly  310  can be configured such that one or more pipes of the assembly  310  bend at a right angle using a typical rigid heat pipe (if thermal distortion isolation is not a concern), a flexible heat pipe (if thermal distortion isolation is a concern, and to facilitate assembly and testing of the system  200 ) or a loop heat pipe system. 
     Accordingly, some embodiments can provide heat removal using a two-sided instrument thermal radiator  400  and an embedded heat pipe system  310 . 
     Additionally, the mounting structure  306  can facilitate heat removal and provide a means for mechanically interconnecting the payload  300  with the nadir deck  240  of the satellite. 
     In addition, some embodiments can be configured such that the heat pipe assembly  310  is flexible. A flexible heat pipe assembly can provide a compliant interface between the heat removal system and components supported on the radiator  400 , and/or the mounting structure  306 . In such embodiments, distortions induced by differential heating of the components of the system  200  can be avoided. For example, expanding and contracting movements of the heat pipe assembly  310  can cause torques or forces to which the instruments of the payload  300  may be sensitive. Thus, in some embodiments, a flexible heat pipe assembly can be used to alleviate torsion, tension, or other forces that could otherwise affect elements of the payload  300 . 
     Referring again to  FIGS. 8A and 9B , the system can comprise one or more mounts  322  coupled to the mounting structure  306  to secure the payload  300  relative to the nadir deck  240  of the satellite  202 . The mounts  322  can comprise one or more kinematic mounts that effectively isolate the system from spacecraft intra-body torques. Thus, loads or forces associated with launching the satellite  202 , which may be exerted on the system  200  and the payload  300 , can be passed through the mounts  322 . In some embodiments, the mounts  322  can comprise kinematic mounts. Further, the system  200 , although illustrated as comprising only mounts  322 , can, in some embodiments, comprise more or fewer mounts  322 , such as one, two, three, four, five, six, seven, eight, nine, ten or more. 
     This application includes description that is provided to enable a person of ordinary skill in the art to practice the various aspects described herein. While the foregoing has described what are considered to be the best mode and/or other examples, it is understood that various modifications to these aspects will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other aspects. It is understood that the specific order or hierarchy of steps or blocks in the processes disclosed is an illustration of exemplary approaches. Based upon design preferences, it is understood that the specific order or hierarchy of steps or blocks in the processes may be rearranged. The accompanying method claims present elements of the various steps in a sample order, and are not meant to be limited to the specific order or hierarchy presented. Thus, the claims are not intended to be limited to the aspects shown herein, but are to be accorded the full scope consistent with the language of the claims. 
     Headings and subheadings, if any, are used for convenience only and do not limit the invention. 
     Reference to an element in the singular is not intended to mean “one and only one” unless specifically so stated, but rather “one or more.” Use of the articles “a” and “an” is to be interpreted as equivalent to the phrase “at least one.” Unless specifically stated otherwise, the terms “a set” and “some” refer to one or more. 
     Terms such as “top,” “bottom,” “upper,” “lower,” “left,” “right,” “front,” “rear” and the like as used in this disclosure should be understood as referring to an arbitrary frame of reference, rather than to the ordinary gravitational frame of reference. Thus, a top surface, a bottom surface, a front surface, and a rear surface may extend upwardly, downwardly, diagonally, or horizontally in a gravitational frame of reference. 
     Although the relationships among various components are described herein and/or are illustrated as being orthogonal or perpendicular, those components can be arranged in other configurations in some embodiments. For example, the angles formed between the referenced components can be greater or less than 90 degrees in some embodiments. 
     Although various components are illustrated as being flat and/or straight, those components can have other configurations, such as curved or tapered for example, in some embodiments. 
     Pronouns in the masculine (e.g., his) include the feminine and neuter gender (e.g., her and its) and vice versa. All structural and functional equivalents to the elements of the various aspects described throughout this disclosure that are known or later come to be known to those of ordinary skill in the art are expressly incorporated herein by reference and are intended to be encompassed by the claims. Moreover, nothing disclosed herein is intended to be dedicated to the public regardless of whether such disclosure is explicitly recited in the claims. No claim element is to be construed under the provisions of 35 U.S.C. §112, sixth paragraph, unless the element is expressly recited using the phrase “means for” or, in the case of a method claim, the element is recited using the phrase “operation for.” 
     Phrases such as an aspect, the aspect, another aspect, some aspects, one or more aspects, an implementation, the implementation, another implementation, some implementations, one or more implementations, an embodiment, the embodiment, another embodiment, some embodiments, one or more embodiments, a configuration, the configuration, another configuration, some configurations, one or more configurations, the subject technology, the disclosure, the present disclosure, other variations thereof and alike are for convenience and do not imply that a disclosure relating to such phrase(s) is essential to the subject technology or that such disclosure applies to all configurations of the subject technology. A disclosure relating to such phrase(s) may apply to all configurations, or one or more configurations. A disclosure relating to such phrase(s) may provide one or more examples. A phrase such as an aspect or some aspects may refer to one or more aspects and vice versa, and this applies similarly to other foregoing phrases. 
     The word “exemplary” is used herein to mean “serving as an example or illustration.” Any aspect or design described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other aspects or designs. 
     All structural and functional equivalents to the elements of the various aspects described throughout this disclosure that are known or later come to be known to those of ordinary skill in the art are expressly incorporated herein by reference and are intended to be encompassed by the claims. Moreover, nothing disclosed herein is intended to be dedicated to the public regardless of whether such disclosure is explicitly recited in the claims. No claim element is to be construed under the provisions of 35 U.S.C. §112, sixth paragraph, unless the element is expressly recited using the phrase “means for” or, in the case of a method claim, the element is recited using the phrase “step for.” Furthermore, to the extent that the term “include,” “have,” or the like is used in the description or the claims, such term is intended to be inclusive in a manner similar to the term “comprise” as “comprise” is interpreted when employed as a transitional word in a claim. 
     Although embodiments of the present disclosure have been described and illustrated in detail, it is to be clearly understood that the same is by way of illustration and example only and is not to be taken by way of limitation, the scope of the present invention being limited only by the terms of the appended claims.