Patent Publication Number: US-2015068703-A1

Title: Thermal management system and method of assembling the same

Description:
BACKGROUND 
     The subject matter described herein relates generally to thermal management systems and, more particularly, to a thermal management system for cooling a heat source onboard an aircraft. 
     Many aircraft utilize electronic systems that generate significant amounts of heat, and the heat should be dissipated from the electronic systems in order to keep the systems functioning properly. At least some known cooling systems for dissipating heat from the electronics on an aircraft include forced-air cooling systems (e.g., cooling systems that blow air over the electronics using ductwork) or forced-liquid cooling systems (e.g., cooling systems that channel liquid coolant through a cooling circuit using a pump). However, in some circumstances, these known cooling systems are ineffective at cooling the electronics and make inefficient use of the heat generated by the electronics. 
     BRIEF DESCRIPTION 
     In one aspect, a thermal management system for cooling a heat source onboard an aircraft is provided. The aircraft has a frame and a skin coupled to the frame such that the skin includes a first segment and a second segment. The thermal management system includes a first network of heat pipes coupled in conductive heat transfer with the heat source and the first segment of the skin. The first network of heat pipes is configured to heat the first segment of the skin using heat from the heat source. The thermal management system further includes a second network of heat pipes coupled in conductive heat transfer with the heat source and the second segment of the skin. The second network of heat pipes is configured to heat the second segment of the skin using heat from the heat source. The thermal management system is configured to selectively deactivate the first network of heat pipes and the second network of heat pipes. 
     In another aspect, a method of assembling a thermal management system for cooling a heat source onboard an aircraft is provided. The aircraft includes a frame and a skin coupled to the frame such that the skin has a first segment and a second segment. The method includes coupling a first network of heat pipes in conductive heat transfer with the heat source and the first segment of the skin. The first network of heat pipes is configured to heat the first segment of the skin using heat from the heat source. The method further includes coupling a second network of heat pipes in conductive heat transfer with the heat source and the second segment of the skin. The second network of heat pipes is configured to heat the second segment of the skin using heat from the heat source. The method also includes coupling a processing unit to the first network of heat pipes and the second network of heat pipes. The processing unit is configured to selectively deactivate the first network of heat pipes and the second network of heat pipes. 
     In another aspect, a thermal management system for cooling a heat source onboard an aircraft having a frame and a skin coupled to the frame is provided. The thermal management system includes a network of heat pipes configured to transfer heat from the heat source to the skin by conductive heat transfer between adjacent heat pipes. 
    
    
     
       DRAWINGS 
       These and other features, aspects, and advantages of the present disclosure will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein: 
         FIG. 1  is a perspective view of an exemplary aircraft; 
         FIG. 2  is a partially cut out view of a wing of the aircraft shown in  FIG. 1 ; 
         FIG. 3  is a block diagram of an exemplary thermal management system for cooling avionics onboard the aircraft shown in  FIG. 1 ; 
         FIG. 4  is a schematic side view of a heat pipe that may be used in the thermal management system shown in  FIG. 3 ; 
         FIG. 5  is schematic cross-sectional view of the heat pipe shown in  FIG. 4  taken along plane  5 - 5  of  FIG. 4 ; and 
         FIG. 6  is a schematic plan view of the aircraft shown in  FIG. 1  including an embodiment of the thermal management system shown in  FIG. 3 . 
     
    
    
     Unless otherwise indicated, the drawings provided herein are meant to illustrate features of embodiments of the disclosure. These features are believed to be applicable in a wide variety of systems comprising one or more embodiments of the disclosure. As such, the drawings are not meant to include all conventional features known by those of ordinary skill in the art to be required for the practice of the embodiments disclosed herein. 
     DETAILED DESCRIPTION 
     In the following specification and the claims, reference will be made to a number of terms, which shall be defined to have the following meanings. 
     The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise. 
     “Optional” or “optionally” means that the subsequently described event or circumstance may or may not occur, and that the description includes instances where the event occurs and instances where it does not. 
     Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about” and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and/or interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. 
     The embodiments disclosed herein facilitate effectively cooling heat sources onboard an aircraft. The embodiments further facilitate cooling the electronic systems of an aircraft, thereby enabling the use of more powerful electronic systems onboard the aircraft. The devices, systems, and methods disclosed herein also facilitate increasing the reliability of functioning heat sources (such as electronic systems) onboard an aircraft by preventing them from overheating. The devices, systems, and methods further facilitate cooling heat sources onboard an aircraft using less electricity, thereby being less of an electrical load on the power supply of the aircraft. Additionally, the embodiments disclosed herein facilitate utilizing the undesired heat produced by existing heat sources onboard an aircraft to alter the infrared signature of the aircraft by redistributing the undesired heat to exterior surfaces of the aircraft prior to dissipating the heat from the aircraft, thereby heating the exterior surfaces of the aircraft using less centerline electrical power than would electrically powered heating elements dedicated for use in heating the exterior surfaces of the aircraft. The embodiments also facilitate providing a thermal management system that weighs less, e.g., in the absence of associated ductwork, pumps, larger quantities of liquid coolant, etc., thereby enabling the aircraft to weigh less and be more fuel efficient. The devices, systems, and methods disclosed herein further facilitate improving the fuel efficiency of an aircraft by transferring heat to the fuel in order to preheat the fuel for improving combustion performance. 
       FIG. 1  is a perspective view of an aircraft  100 . In the exemplary embodiment, aircraft  100  is an unmanned aerial vehicle (UAV) that has a fuselage  102 , a first wing  104 , a second wing  106 , and a tail  108 . In other embodiments, aircraft  100  may be a manned aircraft without departing from the scope of this invention. 
       FIG. 2  is a partially cut out view of first wing  104 . While first wing  104  is described in more detail below, it should be noted that second wing  106  is constructed in the same manner as first wing  104 . In the exemplary embodiment, aircraft  100  has a frame  110  and a skin  112  that are both fabricated from a metallic material, e.g., without limitation, an aluminum material. On first wing  104 , frame  110  includes an arrangement of ribs  114  and spars  116  that support skin  112  such that skin  112  takes on an airfoil-type shape. The airfoil-type shape defines a leading edge (or surface)  118 , a trailing edge (or surface)  120 , a pressure side  122  extending from leading edge  118  to trailing edge  120 , and a suction side  124  extending from leading edge  118  to trailing edge  120  opposite pressure side  122 . In this manner, first wing  104  has a spanwise dimension S and a chordwise dimension C. 
     Ribs  114  extend in the chordwise dimension C and are spaced apart from one another in the spanwise dimension S, and spars  116  extend in the spanwise dimension S and are spaced apart from one another in the chordwise dimension C. During flight, air flows over skin  112  of first wing  104  to provide lift for aircraft  100 . Notably, other embodiments of frame  110  may have any suitable arrangement of structural members associated with first wing  104 , and other embodiments of skin  112  may define any suitable shape of first wing  104 . Moreover, it should be noted that frame  110  also supports skin  112  of fuselage  102 , tail  108 , and other suitable regions of aircraft  100 , e.g., frame  110  supports skin  112  such that fuselage  102  takes on a generally tubular shape having a top  126  and a bottom  128 , which are shown in  FIG. 1 . 
       FIG. 3  is a block diagram of a thermal management system  200  for cooling avionics  300  (i.e., an electronic heat dissipating source) onboard aircraft  100 . Thermal management system  200  includes a thermal rail  202  and a plurality of heat pipes  204  coupled in conductive heat transfer with thermal rail  202 , as set forth in more detail below. In the exemplary embodiment, thermal rail  202  is a metal plate or rod disposed within fuselage  102 , and avionics  300  are coupled in conductive heat transfer with thermal rail  202  such that heat is transferrable from avionics  300  to thermal rail  202  via conduction. Notably, thermal rail  202  may be any suitable shape, may be fabricated from any suitable conductive material, and may be coupled to avionics  300  in any suitable manner that enables thermal rail  202  to function as described herein. 
     In the exemplary embodiment, heat pipes  204  are arranged as follows. A first network  206  of heat pipes  204  is coupled in conductive heat transfer with a first segment  130  of skin  112  (e.g., suction side  124  of first wing  104 ). A second network  208  of heat pipes  204  is coupled in conductive heat transfer with a second segment  132  of skin  112  (e.g., pressure side  122  of first wing  104 ). A third network  210  of heat pipes  204  is coupled in conductive heat transfer with a third segment  134  of skin  112  (e.g., suction side  124  of second wing  106 ). A fourth network  212  of heat pipes  204  is coupled in conductive heat transfer with a fourth segment  136  of skin  112  (e.g., pressure side  122  of second wing  106 ). A fifth network  214  of heat pipes  204  is coupled in conductive heat transfer with a fifth segment  138  of skin  112  (e.g., top  126  of fuselage  102 ). A sixth network  216  of heat pipes  204  is coupled in conductive heat transfer with a sixth segment  140  of skin  112  (e.g., bottom  128  of fuselage  102 ). A seventh network  218  of heat pipes  204  is coupled in conductive heat transfer with fuel  144  contained in a fuel tank, a fuel line, or other suitable fuel-containing structure of aircraft  100 . 
     Notably, each network  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  may be made up of any suitable number of heat pipes  204 , i.e., one or more heat pipes  204 , arranged in any suitable manner. Additionally, thermal management system  200  may include any suitable number of networks coupled to any suitable heat source(s) of aircraft  100  in any suitable manner that facilitates dissipating heat from the heat source(s) as described herein. In one alternative embodiment, heat pipes  204  may be coupled directly to avionics  300 , thereby eliminating thermal rail  202  from thermal management system  200 . Moreover, in other embodiments, thermal management system  200  may include any suitable heat moving structures that function in lieu of, or in conjunction with, heat pipes  204  to facilitate enabling the dissipation of heat from heat source(s) onboard aircraft  100  as described herein. 
       FIGS. 4 and 5  are schematic side and cross-sectional views of an exemplary heat pipe  204  of thermal management system  200 . In the exemplary embodiment, heat pipe  204  is a two-phase, capillary device having a sealed first end  220 , a sealed second end  222 , and a hollow body  224  extending from first end  220  to second end  222  along a lengthwise dimension L of heat pipe  204 . A permeable membrane  226  lines an interior surface  228  of body  224  to define an elongate central void  230  extending between first end  220  and second end  222 . A working liquid  232  is disposed within body  224  in a quantity sufficient to permeate at least a portion of membrane  226  without preventing gas from flowing lengthwise along central void  230 . 
     In the exemplary embodiment, heat pipe  204  is fabricated from an aluminum material. In other embodiments, however, heat pipe  204  may be fabricated from any suitable materials such as, for example, a copper material or a titanium material. Additionally, while working liquid  232  may be any suitable liquid, working liquid  232  is a butane liquid or an ammonia liquid in the exemplary embodiment. Moreover, in some suitable embodiments, heat pipe  204  may be configured to function as a structural member of aircraft  100 . For example, in the exemplary embodiment of aircraft  100 , frame  110  of first wing  104 , second wing  106 , and/or fuselage  102  includes structural members, e.g., ribs  114  and spars  116 , fabricated from an extruded aluminum material such that each structural member has a hollow core. In this manner, at least one of the structural members may be utilized as heat pipe  204  by adding membrane  226  and working liquid  232  into the hollow core of the structural member. In other embodiments, however, heat pipe  204  may not be configured to function as a structural member of aircraft  100 . 
     In one exemplary function of heat pipe  204 , heat pipe  204  is heated near first end  220 , and working liquid  232  permeating membrane  226  within heat pipe  204  near first end  220  vaporizes such that a vapor  234  travels toward second end  222  along central void  230 . Because heat pipe  204  is cooler near second end  222 , vapor  234  condenses near second end  222 , thereby warming body  224  near second end  222 . The condensate, i.e., working liquid  232 , that results near second end  222  is then wicked toward first end  220  via capillaries in membrane  226  such that the working liquid  232  is again heated near first end  220 . In this manner, heat pipe  204  maintains a cycle of vaporizing and condensing working liquid  232  to transfer heat along its length L, thereby providing heat spreading performance for its weight. Notably, in other modes of operation, heat pipe  204  may be heated at any suitable location along its length L to transfer heat to any other suitable location along its length L, as opposed to being heated near first end  220  for transferring heat toward second end  222 . 
     Because the size of the particles that make up membrane  226  affects the size of the capillaries within membrane  226  and, therefore, the wicking strength of membrane  226 , the particles of membrane  226  in the exemplary embodiment are made small enough that, when working liquid  232  in membrane  226  is under the influence of an external force (such as gravity) that opposes the intended lengthwise direction of wicking, membrane  226  continues to effectively wick working liquid  232  toward the portion of heat pipe  204  that is being heated. In that regard, heat pipe  204  is configured to wick working liquid  232  against multiple times the force of gravity such that only a negligible difference in heat transfer capability is experienced as a result of changes to the orientation or altitude of heat pipe  204 . More specifically, in the exemplary embodiment, the capillary structure of membrane  226  is made small enough to wick working liquid  232  against forces as strong as five to ten times the force of gravity, thereby enabling heat pipe  204  to effectively transfer heat along its length L in any orientation of heat pipe  204  and at altitudes of between 45,000 and 65,000 feet above sea level (or any other suitable altitude outside of such range). 
       FIG. 6  is a schematic plan view of aircraft  100  having an embodiment of thermal management system  200 . In the embodiment of  FIG. 6 , thermal rail  202  (and, therefore, avionics  300 ) is located within fuselage  102 , and networks  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  of heat pipes  204  extend from thermal rail  202  throughout aircraft  100 , as set forth in more detail below. More specifically, first network  206  is arranged in a herringbone-type configuration as follows. A first primary heat pipe  236  is coupled in conductive heat transfer with thermal rail  202  and extends into first wing  104 . A plurality of first secondary heat pipes  238  extend obliquely from first primary heat pipe  236  within first wing  104 . A plurality of first tertiary heat pipes  240  extend obliquely from each first secondary heat pipe  238  within first wing  104 . First primary heat pipe  236 , first secondary heat pipes  238 , and/or first tertiary heat pipes  240  are coupled in conductive heat transfer with first segment  130  of skin  112  on suction side  124  of first wing  104 . Notably, this embodiment of thermal management system  200  may also include second network  208  of heat pipes  204  (as shown in  FIG. 3 ), and second network  208  of heat pipes  204  may be arranged in a herringbone-type configuration extending from thermal rail  202  into first wing  104  in a manner similar to that of first network  206 , with at least one heat pipe  204  of second network  208  being coupled in conductive heat transfer with second segment  132  of skin  112  on pressure side  122  of first wing  104 . 
     Similarly, third network  210  is also arranged in a herringbone-type configuration as follows. A third primary heat pipe  242  is coupled in conductive heat transfer with thermal rail  202  and extends into second wing  106 . A plurality of third secondary heat pipes  244  extend obliquely from third primary heat pipe  242  within second wing  106 . A plurality of third tertiary heat pipes  246  extend obliquely from each third secondary heat pipe  244  within second wing  106 . Third primary heat pipe  242 , third secondary heat pipes  244 , and/or third tertiary heat pipes  246  are coupled in conductive heat transfer with third segment  134  of skin  112  on suction side  124  of second wing  106 . Notably, this embodiment of thermal management system  200  may also include fourth network  212  of heat pipes  204  (as shown in  FIG. 3 ), and fourth network  212  of heat pipes  204  may be arranged in a herringbone-type configuration extending from thermal rail  202  into second wing  106  in a manner similar to that of third network  210 , with at least one heat pipe  204  of fourth network  212  being coupled in conductive heat transfer with skin  112  on pressure side  122  of second wing  106 . 
     Additionally, fifth network  214  is also arranged in a herringbone-type configuration as follows. A fifth primary heat pipe  248  is coupled in conductive heat transfer with thermal rail  202  within fuselage  102 . A plurality of fifth secondary heat pipes  250  extend obliquely from fifth primary heat pipe  248  within fuselage  102 . A plurality of fifth tertiary heat pipes  252  extend obliquely from each fifth secondary heat pipe  250  within fuselage  102 . Fifth primary heat pipe  248 , fifth secondary heat pipes  250 , and/or fifth tertiary heat pipes  252  are coupled in conductive heat transfer with skin  112  on top  126  of fuselage  102 . Notably, this embodiment of thermal management system  200  may also include sixth network  216  of heat pipes  204  (as shown in  FIG. 3 ), and sixth network  216  of heat pipes  204  may be arranged in a herringbone-type configuration extending from thermal rail  202  within fuselage  102  in a manner similar to that of fifth network  214 , with at least one heat pipe  204  of sixth network  216  being coupled in conductive heat transfer with skin  112  on bottom  128  of fuselage  102 . 
     Alternatively, thermal rail  202  and avionics  300  may be located in any suitable area of aircraft  100  such as, for example, an area other than fuselage  102 . Moreover, heat pipes  204  of each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  may be arranged in any suitable manner, e.g., heat pipes  204  of each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  may extend from thermal rail  202  in parallel to one another, may be oriented perpendicular to one another, or may extend along curvilinear paths. In some embodiments, each heat pipe  204  of each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  may be sealed with respect to other heat pipes  204  within the same network  206 ,  208 ,  210 ,  212 ,  214 ,  216 , i.e., each heat pipe  204  within each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  may utilize its own distinct quantity of working liquid  232  and may be coupled (in conductive heat transfer) to other heat pipes  204  within the same network  206 ,  208 ,  210 ,  212 ,  214 ,  216  such that heat is transferred in series from one heat pipe  204  to another heat pipe  204  within each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  via conduction between adjacent heat pipes  204 . In other embodiments, heat pipes  204  within each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  may be coupled in fluid communication with one another such that heat pipes  204  within each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  share a common quantity of working liquid  232 , effectively providing a single, branched heat pipe  204  for each network  206 ,  208 ,  210 ,  212 ,  214 ,  216 . Notably, in the embodiment of  FIG. 6  and other suitable embodiments, each heat pipe  204  (and, therefore, each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  of heat pipes  204 ) has distal end(s)  254  and is not a continuous loop, thereby relying on two-phase, bidirectional fluid flow in order to transfer heat from the heat source, e.g., avionics  300 , to the heat sink, e.g., skin  112 . 
     During operation of aircraft  100  (which is a UAV in the exemplary embodiment), avionics  300  onboard aircraft  100  may operate a navigation system, a remote sensing system, and a propulsion system, among others systems. As such, avionics  300  are a significant electrical load for the power source onboard aircraft  100 , thereby producing a significant amount of heat. However, when flying at higher altitudes, e.g., 45,000-65,000 feet above sea level, the air density is significantly less than that at sea level, meaning that the convective heat transfer coefficient that permits heat transfer to the ambient air becomes more difficult to manage for removing heat from aircraft  100 . 
     In that regard, thermal management system  200  facilitates effectively removing heat from avionics  300  (and aircraft  100 ) at higher altitudes. More specifically, the heat from avionics  300  is transferred to thermal rail  202  via conduction, and thermal rail  202  transfers the heat via conduction to one or more heat pipes  204  within each network  206 ,  208 ,  210 ,  212 ,  214 ,  216 . The heat is then transferred along each heat pipe  204  (or network  206 ,  208 ,  210 ,  212 ,  214 ,  216  of heat pipes  204 ) as set forth above. The heat is then transferred from at least one heat pipe  204  of each network  206 ,  208 ,  210 ,  212 ,  214 ,  216  to an associated segment  130 ,  132 ,  134 ,  136 ,  138 ,  140  of skin  112  via conduction, thereby heating skin  112  such that the heat generated by avionics  300  is dissipated from skin  112  by air flowing over skin  112  while aircraft  100  is in flight. In this manner, thermal management system  200  facilitates cooling avionics  300  by effectively moving the heat generated by avionics  300  to skin  112  (or other suitable exterior surface) of aircraft  100 , enabling removal of the heat from aircraft  100  by ambient air flowing over skin  112  during flight. 
     By transferring the heat from avionics  300  to larger exterior surfaces of aircraft  100  (such as skin  112 ), the heat is efficiently dissipated into the ambient despite the low air density that may exist at altitude. As a result, the junction temperatures of avionics  300  are better managed, and more powerful avionics  300  may be utilized onboard aircraft  100 . Moreover, because each heat pipe  204  is configured to provide effective heat transfer in any orientation of heat pipe  204  and at higher altitudes, thermal management system  200  is less affected by changes to the orientation and altitude of aircraft  100  during flight, e.g., thermal management system  200  is able to effectively remove heat from avionics  300  despite aircraft  100  flying in a banked orientation for annular patterning at 45,000-65,000 feet above sea level. 
     Furthermore, it may be desirable in some instances to render aircraft  100  less detectable during flight. Noting that aircraft  100  is susceptible to detection from below (via ground or sea-based systems) and from above (via satellite or spacecraft-based systems), the temperature of the land, sea, and sky typically provide thermal backdrops for detecting aircraft  100  when viewing aircraft  100  using a thermal imagery system. More specifically, because the various operating systems onboard aircraft  100  generate heat, aircraft  100  naturally takes on an infrared signature (or profile) that is different, i.e., warmer or cooler, than its land backdrop, sea backdrop, or sky backdrop, thereby making it possible to detect aircraft  100  from a thermal imagery standpoint. As such, it may be desirable to alter the infrared signature of aircraft  100  to minimize being detected. For example, in some instances, it may be desirable to heat up or cool down the exterior surfaces, e.g., skin  112 , of aircraft  100  to better match the temperature of a selected backdrop of land, sea, and/or sky, effectively camouflaging aircraft  100 . Moreover, in other instances, it may be desirable to render the infrared signature of aircraft  100  indiscernible or to eliminate the infrared signature of aircraft  100  altogether. 
     To implement these anti-detection techniques, thermal management system  200  may be further configured to utilize the heat generated by avionics  300  to alter the infrared signature of aircraft  100 . Referring again to  FIG. 3 , thermal management system  200  may further include a processing unit  400  configured to selectively (and independently) deactivate the various networks  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  of heat pipes  204  from their associated heat source(s), e.g., avionics  300 . As used herein, the term “deactivating” a network or any variation thereof refers to actively reducing heat transfer along the network. 
     To this end, in some embodiments, at least one heat pipe  204  of each network  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  may be a variable conductance heat pipe that is operable to deactivate its associated network  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  of heat pipes  204 . The variable conductance heat pipe  204  would be provided with a reservoir  256  of gas, e.g., argon or helium gas, in flow communication with central void  230  such that, when the gas is heated within reservoir  256  and is permitted to exit reservoir  256 , the gas expands within reservoir  256  and exits reservoir  256  to fill central void  230 , thereby preventing (or inhibiting) vapor  234  from traveling along central void  230  and, therefore, deactivating the associated network  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  of heat pipes  204 . To the contrary, when the gas within reservoir  256  is subsequently cooled, the gas contracts and reenters reservoir  256  to no longer prevent (or inhibit) vapor  234  from traveling through central void  230 . Notably, reservoir  256  may be heated in any suitable manner, such as, for example, by an electric heater  258  coupled to reservoir  256 . 
     As such, processing unit  400  may be configured to operate electric heaters  258  (or other suitable heating devices) in order to selectively (and independently) heat the gas within each reservoir  256  in order to deactivate any one or more network(s)  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218 . In other embodiments, thermal management system  200  may include a thermal switch or a bimetallic cantilever associated with each network  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  of heat pipes  204  to facilitate deactivating the various networks  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218 . Alternatively, thermal management system  200  may include any suitable mechanism that acts as a thermal valve of sorts for each heat pipe  204  or each network  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  of heat pipes  204 . 
     Because thermal management system  200  may be configured to permit selective and independent deactivation of networks  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218 , thermal management system  200  provides the added benefit of being able to select (via processing unit  400 ) which exterior segments of aircraft  100 , e.g., skin  112 , are to be heated using the heat generated by avionics  300 . As such, thermal management system  200  enables continuous (or spontaneous) manipulation of the infrared signature of aircraft  100  during flight by selecting to deactivate some network(s)  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  without selecting to deactivate other network(s)  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218 . For example, in one possible sequence, processing unit  400  may deactivate second network  208 , fourth network  212 , sixth network  216 , and seventh network  218  such that the heat from avionics  300  is transferred only to first network  206 , third network  210 , and fifth network  214 , thereby heating only suction side  124  of first wing  104 , suction side  124  of second wing  106 , and top  126  of fuselage  102  in order to facilitate increasing the temperature of the infrared signature of aircraft  100  when viewed from above and decreasing the temperature of the infrared signature of aircraft  100  when viewed from below, thereby camouflaging aircraft  100  such that aircraft  100  has an infrared signature that is more in line with a desired lower-altitude backdrop, e.g., land and/or sea. 
     Alternatively, in another possible sequence, processing unit  400  may repeatedly deactivate one or more networks  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  on a periodic or random time interval in order to render the infrared signature less discernible as being that of an aircraft. For example, processing unit  400  may repeatedly alternate between the following states on a random time interval: (1) simultaneously deactivating third network  210 , fourth network  212 , fifth network  214 , and sixth network  216 ; and (2) simultaneously deactivating first network  206 , second network  208 , fifth network  214 , and sixth network  216 . In this example, the infrared signature of aircraft  100  during flight would fluctuate randomly between the first state (during which only pressure side  122  and suction side  124  of first wing  104  would be heated using heat from avionics  300 ) and the second state (during which only pressure side  122  and suction side  124  of second wing  106  would be heated using heat from avionics  300 ). As such, it would be more difficult to detect the infrared signature as being that of an aircraft. 
     Being that seventh network  218  is coupled in conductive heat transfer with fuel  144  inside of a fuel tank, a fuel line, or other suitable fuel-containing structure of aircraft  100 , at least some heat generated by avionics  300  may also be selectively transferrable to fuel  144  using processing unit  400 . In this manner, fuel  144  may be selectively preheated to enhance engine combustion, thereby rendering aircraft  100  more fuel efficient. In one embodiment, seventh network  218  may be coupled to fuel  144  by virtue of at least one heat pipe  204  of seventh network  218  being inserted directly into fuel  144  within a fuel tank and including miniature fins at the location of interfacing with fuel  144  in order to more efficiently transfer heat to fuel  144 . In another embodiment, a heat exchanger may be provided along a fuel line to exchange heat between fuel  144  and at least one heat pipe  204  of seventh network  218 . Alternatively, at least one heat pipe  204  of seventh network  218  may be integrated into walls of a fuel tank, a fuel line, or other fuel-containing structure to surround fuel  144 , rather than being inserted directly into fuel  144 . Moreover, if a heat pipe  204  of seventh network  218  was to be inserted directly into fuel  144 , the size of heat pipe  204  may be reduced in order to minimize the volume of the fuel-containing structure that is being occupied by heat pipe  204 , thereby enabling the fuel-containing structure to house a larger quantity of fuel  144 . Notably, thermal management system  200  may also be configured to remove heat from fuel  144  if desired, thereby transferring heat from fuel  144  toward thermal rail  202  along seventh network  218  in order to transfer the heat from fuel  144  to any desirable region of aircraft  100 , much like the transfer of heat from avionics  300  described above. 
     Optionally, in another embodiment, a thermal energy storage element  260  such as, for example, a capacitor or other energy storage buffer may be provided along one or more of networks  206 ,  208 ,  210 ,  212 ,  214 ,  216 ,  218  (or along an eighth network that would be dedicated to transferring heat from the heat source, e.g., avionics  300 , to a larger energy storage element) for selectively (and simultaneously) deactivating all of networks  206 ,  208 ,  210 ,  212 ,  214 ,  216  and possibly seventh network  218 . In this manner, the infrared signature of aircraft  100  may be eliminated altogether. For example, in one sequence, the radiation of heat from aircraft  100  may cease entirely (or at least be reduced) by simultaneously deactivating first network  206 , second network  208 , third network  210 , fourth network  212 , fifth network  214 , and sixth network  216 , diverting all of the heat normally associated with those networks  206 ,  208 ,  210 ,  212 ,  214 ,  216  to at least one corresponding thermal energy storage element  260 . 
     In the exemplary embodiment, processing unit  400  suitably includes at least one processor, a memory device coupled to the processor, and at least one input/output (I/O) conduit, wherein the conduit includes at least one I/O channel. As used herein, the term processor is not limited to just those integrated circuits referred to in the art as a computer, but broadly refers to a microcontroller, a microcomputer, a programmable logic controller (PLC), an application specific integrated circuit, and other programmable circuits, and these terms are used interchangeably herein. In the embodiments described herein, the memory device may include, but is not limited to, a computer-readable medium, such as a random access memory (RAM), and a computer-readable non-volatile medium, such as flash memory. Alternatively, a floppy disk, a compact disc-read only memory (CD-ROM), a magneto-optical disk (MOD), and/or a digital versatile disc (DVD) may also be used. 
     In the embodiments described herein, I/O channels may be associated with, but are not limited to, computer peripherals associated with an operator interface such as a mouse and a keyboard. Alternatively, other computer peripherals may also be used that may include, for example, but not be limited to, a control stick for use by an operator in controlling aircraft  100 . Furthermore, in the exemplary embodiment, additional I/O channels may be associated with, but not be limited to, an operator interface monitor or a communications link for remotely controlling aircraft  100  and/or thermal management system  200 . Moreover, the processor may process information transmitted from a plurality of electronic devices onboard aircraft  100 , including, without limitation, electric heaters  258  (and/or other suitable network deactivation mechanisms) or temperature sensors suitably dispersed throughout aircraft  100 . The memory device and the storage devices store and transfer information and instructions to be executed by the processor. The memory device and the storage devices can also be used to store and provide temporary variables, static, i.e., non-volatile and non-changing, information and instructions, or other intermediate information to the processor during execution of instructions by the processor. Instructions that are executed include, but are not limited to, analysis of signals transmitted from electric heaters  258  and/or other suitable network deactivation mechanisms. The execution of sequences of instructions is not limited to any specific combination of hardware circuitry and software instructions. 
     The above-described embodiments facilitate effectively cooling heat sources onboard an aircraft. The devices, systems, and methods disclosed herein further facilitate cooling the electronic systems of an aircraft, thereby enabling the use of more powerful electronic systems onboard the aircraft. The devices, systems, and methods also facilitate increasing the reliability of functioning heat sources (such as electronic systems) onboard an aircraft by better preventing them from overheating. The embodiments disclosed herein further facilitate cooling heat sources onboard an aircraft using less electricity, thereby being less of an electrical load on the power supply of the aircraft. Additionally, the embodiments facilitate utilizing the undesired heat produced by existing heat sources onboard an aircraft to alter the infrared signature of the aircraft by redistributing the undesired heat to exterior surfaces of the aircraft prior to dissipating the heat from the aircraft, thereby heating the exterior surfaces of the aircraft using less centerline electrical power than would electrically powered heating elements dedicated for use in heating the exterior surfaces of the aircraft. The above-described embodiments also facilitate providing a thermal management system that weighs less, e.g., in the absence of associated ductwork, pumps, larger quantities of liquid coolant, etc., thereby enabling the aircraft to weigh less and be more fuel efficient. The devices, systems, and methods further facilitate improving the fuel efficiency of an aircraft by transferring heat to the fuel in order to preheat the fuel for improving combustion performance. 
     An exemplary technical effect of the methods, systems, and apparatus described herein includes at least one of: (a) effectively cooling heat sources onboard an aircraft; (b) cooling electronic systems of an aircraft, thereby enabling the use of more powerful electronics onboard the aircraft; (c) increasing the reliability of functioning heat sources (such as electronic systems) onboard an aircraft by better preventing the heat sources from overheating; (d) cooling heat sources onboard an aircraft using less electricity, thereby being less of an electrical load on the power supply of the aircraft; (e) utilizing heat produced by existing heat sources onboard an aircraft to alter an infrared signature of the aircraft by redistributing the heat to exterior surfaces of the aircraft prior to dissipating the heat from the aircraft; (f) providing a thermal management system that weighs less, e.g., in the absence of associated ductwork, pumps, larger quantities of liquid coolant, etc., thereby enabling the aircraft to weigh less and be more fuel efficient; and (g) improving the fuel efficiency of an aircraft by preheating the fuel to facilitate better combustion performance. 
     Exemplary embodiments of thermal management systems and methods of assembling the same are described above in detail. The systems and methods are not limited to the specific embodiments described herein, but rather, components of the systems and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein. For example, the methods may also be used in combination with other systems and methods, and are not limited to practice with only the aircraft-related systems and methods described herein. Rather, the embodiments may be implemented and utilized in connection with many other applications outside of aviation. 
     Although specific features of various embodiments of the invention may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the invention, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing. 
     This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.