Abstract:
A thermal management system for a vehicle includes a tank, a heat exchanger, a pump, and a valve located on a conduit. The heat exchanger is located downstream of the tank, the pump is located between the tank and the heat exchanger, and the valve is located downstream of the heat exchanger. The heat exchanger places the coolant in a heat exchange relationship with a heat load from the vehicle such that the coolant vaporizes. The valve regulates pressure within the heat exchanger and controls exhaustion of the vaporized coolant from the vehicle. Water, used as a coolant, can be replenished in flight by condensing a portion of the heat exchanger exhaust or condensing water as part of the environmental control system function or by condensing a portion of the engine exhaust.

Description:
BACKGROUND 
     The present disclosure relates generally to thermal management, and more particularly to thermal management for electronic systems aboard vehicles. 
     Thermal Management Systems (TMS) and Environmental Control Systems (ECS) for aircrafts are commonly sized for the worst case coincident design conditions of thermal load, heat sink temperature and availability, and power availability. Mission segments that may result in an unbalance between the load, heat sink and power include hot day ground operation at engine idle, hot day low altitude high speed dash, and high altitude top of descent. Depending on aircraft requirements and system architectures, the TMS and/or ECS may operate at significantly less than design capacity during the majority of the mission. This results in larger and heavier systems than would be required if the TMS and/or ECS were sized for mission average conditions. 
     In aircrafts, thermal loads are becoming increasing less steady-state and de-coupled from available power. One trend is the replacement of hydraulic flight controls with electronic flight controls. Thermal loads from electronic flight controls will peak on approach to landing when engine power settings and speed are reduced. Radar and other thermal loads from electronic systems may be high at any one part of the flight, regardless of engine power setting. 
     SUMMARY 
     A thermal management system for a vehicle includes a tank, a heat exchanger, a pump, and a first valve. The tank is located onboard the vehicle and stores coolant. The heat exchanger is located downstream of the tank on a first conduit. The heat exchanger places the coolant in a heat exchange relationship with a heat load from the vehicle such that the coolant vaporizes. The pump, which is located between the tank and the heat exchanger on the first conduit, pumps the coolant from the tank to the heat exchanger. The first valve, which is located downstream of the heat exchanger on the first conduit, regulates pressure within the heat exchanger and controls exhaustion of the vaporized coolant from the vehicle. 
     In one embodiment of the thermal management system, the vehicle is an aircraft, the coolant is water, the heat exchanger is a boiler, and the valve is a backpressure valve. Within the boiler, a heat load from an electronics system vaporizes the water into steam. The backpressure valve regulates pressure within the boiler and controls exhaustion of the steam from the aircraft. 
     A method for managing heat for an electronics system onboard an aircraft includes storing a liquid coolant on the aircraft, and pumping the liquid coolant to a heat exchanger. The method further includes vaporizing the liquid coolant into a gas coolant in the heat exchanger by absorbing a heat load from the electronics system, and exhausting the gas coolant from the aircraft. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a perspective view of an aircraft having a heat management system for an electronics system in accordance with the present invention. 
         FIG. 2  is a block diagram showing a first embodiment of the heat management system. 
         FIG. 3  is a block diagram showing a second embodiment of the heat management system having an ejector. 
         FIG. 4  is a block diagram showing a third embodiment of the heat management system having a compressor and a coolant replenishment system. 
         FIGS. 5-7  are block diagrams showing other embodiments of the coolant replenishment system. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  is a perspective view of aircraft  10  having heat management system  12  for electronic system  14 . In the depicted embodiment, aircraft  10  is a commercial airplane, although other types of airplanes (e.g. private or military), aircrafts (e.g. helicopters), and land vehicles are also possible. Heat management system  12  and electronic system  14  are both located onboard aircraft  10 . Electronic system  14  can be any electronic system for aircraft  10  (e.g. flight controls, landing gear, radar, or weapons). Electronic system  14  generates heat load H, which varies between a minimum and a maximum or “peak” load. Heat management system  12  has a heat exchange relationship with electronic system  14  such that heat management system  12  absorbs heat load H from electronic system  14 . Heat management system  12  rejects heat load H outside of aircraft  10  via an expendable coolant. Heat management system  12  provides a capacity for cooling heat load H from electronic system  14 . The details of heat management system  12  are discussed below with respect to  FIGS. 2-7 . 
       FIG. 2  is a block diagram showing a first embodiment of heat management system  12 A for absorbing heat load H from electronic system  14 . Heat management system  12 A includes tank  16 , pump  18 , heat exchanger  20 , and valve  22  located on first conduit  24 . Tank  16  includes a volume of coolant  26  and drain  28 . Coolant  26  changes state from liquid coolant  26 L to vapor coolant  26 V as it absorbs heat load H within heat exchanger  20 . 
     Tank  16 , pump  18 , heat exchanger  20 , and valve  22  are located in flow series along first conduit  24 . Tank  16  contains and stores a volume of coolant  26  onboard aircraft  10  and therefore, tank  16  acts as a thermal accumulator. Tank  16  can be insulated to prevent coolant  26  from freezing due to cold ambient temperatures either during flight or on the ground. In addition to insulation, tank  16  can also include an electric heating element. If extra cooling capacity is found on aircraft  10 , tank  16  can benefit from cooling to reduce the temperature of coolant  26 . Drain  28  is included on tank  16  to remove coolant  26  from aircraft  10  during periods of inactivity. In the depicted embodiment, coolant  26  is stored in tank  16  as liquid coolant  26 L. Pump  18  is located downstream of tank  16  on first conduit  24  and pumps liquid coolant  26 L from tank  16  to heat exchanger  20 . If the amount of liquid coolant  26 L flowing through first conduit  24  is relatively small, pump  18  can be a vane-type pump. If the amount of liquid coolant  26 L flowing through first conduit  24  is relatively large, pump  18  can be a centrifugal-type pump. Heat exchanger  20  is located downstream of pump  18  on first conduit  24  and receives liquid coolant  26 L from pump  18 . 
     Liquid coolant  26 L enters a first side of heat exchanger  20  in order to absorb heat load H from electronic system  14 . Within heat exchanger  20 , heat load H is rejected into liquid coolant  26 L, which vaporizes and becomes vapor coolant  26 V. Vapor coolant  26 V, which contains heat load H, exits a second side of heat exchanger  20 . Heat load H can be carried from electronic system  14  to heat exchanger  20  by either a liquid (e.g. propylene glycol/water, PAO oil) or a gas (e.g. air). Accordingly, heat exchanger  20  is either a liquid-to-liquid heat exchanger, a liquid-to-gas heat exchanger, or a direct contact heat exchanger. Heat load H can also have indirect thermal contact with heat exchanger  20 . In the depicted embodiment, heat exchanger  20  is a boiler, liquid coolant  26 L is water, and vapor coolant  26 V is steam. Located downstream of heat exchanger  20  on first conduit  24  is valve  22 . Valve  22  is an exhaust backpressure valve used to regulate pressure within heat exchanger  20  and thereby, control the temperature of vaporization for coolant  26 . If valve  22  is mostly open, vapor coolant  26 V is freely vented to the environment outside of aircraft  10 . If valve  22  is partially closed, exhaustion of vapor coolant  26 V to the environment is reduced and the temperature of vaporization for coolant  26  within heat exchanger  20  is increased. For example, if coolant  26  is water, the boiling point within heat exchanger  20  could be increased from about 212° F. (100° C.) at low altitude to about 300° F. (149° C.). Valve  22 , therefore, allows for temperature management of coolant  26  regardless of the altitude of aircraft  10 . Coolant  26  absorbs heat load H through latent heat of evaporation, and coolant  26  is expendable in that vapor coolant  26 V is rejected to the atmosphere. 
       FIG. 3  is a block diagram showing a second embodiment of heat management system  12 B having ejector  30 . Heat management system  12 B includes tank  16 , pump  18 , heat exchanger  20 , valve  22 , and ejector  30  located on first conduit  24 . Tank  16  includes a volume of coolant  26  and drain  28 . Heat management similar  12 B is similar to heat management system  12 A described above with reference to  FIG. 2 , and like reference numerals indicate like components. 
     As described above, coolant  26  changes state from liquid coolant  26 L to vapor coolant  26 V as it absorbs heat load H within heat exchanger  20 . Ejector  30  is located just downstream of heat exchanger  20  and upstream of valve  22  on first conduit  24 . Ejector  30  pumps compressed air (e.g. bleed air from the engine) into first conduit  24  where it mixes with vapor coolant  26 V. The injection of compressed air into first conduit  24  reduces pressure within heat exchanger  20  through suction or creation of a partial vacuum. A reduction of pressure within heat exchanger  20  also reduces the temperature of vaporization for coolant  26 . For example, if coolant  26  is water, the boiling temperature could be reduced from about 212° F. (100° C.) to as low as 150° F. (66° C.) at low altitude by inclusion of ejector  30  downstream of heat exchanger  20 . Accordingly, heat management system  12 B allows increased control over coolant  26  temperature regardless of ambient pressure. A mechanical suction pump can be used instead of ejector  30  if energy efficiency takes precedence over weight for aircraft  10 . 
       FIG. 4  is a block diagram showing a third embodiment of heat management system  12 C having compressor  32  and a first embodiment of coolant replenishment system  34 A. Heat management system  12 C includes tank  16 , pump  18 , heat exchanger  20 , valve  22 , and compressor  32  located on first conduit  24 . Tank  16  includes a volume of coolant  26  and drain  28 . Coolant replenishment system  34 A includes shutoff valve  36 , condenser  38 , and expansion valve  40  located on second conduit  42 . Heat management system  12 C is similar to heat management system  12 B described above with reference to  FIG. 3 , and like reference numerals indicate like components. 
     As described above, coolant  26  changes state from liquid coolant  26 L to vapor coolant  26 V as it absorbs heat load H within heat exchanger  20 . Compressor  32  is located just downstream of heat exchanger  20  and upstream of valve  22  on first conduit  24 . Compressor  32  functions similarly to ejector  30  described above with reference to  FIG. 3 . Compressor  32  pumps compressed air into first conduit  24  where it mixes with vapor coolant  26 V. The injection of compressed air into first conduit  24  reduces pressure within heat exchanger  20  through suction or creation of a partial vacuum. A reduction in pressure within heat exchanger  20  also reduces the temperature of vaporization for coolant  26 . Compressor  32  can be driven by a motor or a turbine (e.g. turbo-compressor) such that the cool air can be used by another subsystem in aircraft  10 . A turbine-driven compressor is typically more efficient than an ejector, but is also more expensive, heavier, and assumes a higher volume. An assessment of efficiency versus weight can determine whether heat management system  12 B having ejector  30  or heat management system  12 C having compressor  32  is a better fit for aircraft  10 . 
     In  FIG. 4 , heat management system  12 C is associated with a first embodiment of coolant replenishment system  34 A. Since coolant  26  is expendable, it is advantageous to have a means for generating or recycling coolant  26  onboard aircraft  10 . Coolant replenishment system  34 A includes shutoff valve  36 , condenser  38 , and expansion valve  40  located in flow series along second conduit  42 . A first end or inlet of second conduit  42  is attached to first conduit  24  at a location between compressor  32  and valve  22 . A second end or outlet of second conduit  42  is attached to first conduit  24  at a location between pump  18  and heat exchanger  20 . 
     During periods of peak heat load H from electronic system  14 , shutoff valve  36  on second conduit  42  is partially or totally closed such that coolant replenishment system  34 A is partially or totally closed off from heat management system  12 C. Accordingly, vapor coolant  26 V (and heat load H) continues straight from compressor  32  along first conduit  24  and is exhausted through valve  22  to the environment. In other words, coolant  26  is partially or totally expendable during periods of peak heat load H. During period of less than peak heat load H from electronic system  14 , shutoff valve  36  is partially or fully open, thereby allowing vapor coolant  26 V to flow from heat exchanger  20  into second conduit  42 . Second conduit  42  sends vapor coolant  26 V to condenser  38 . Within condenser  38 , vapor coolant  26 V is condensed into pressurized liquid coolant  26 PL and heat load H is rejected into ram air flow or the fuel system of aircraft  10 . Pressurized liquid coolant  26 PL exits condenser  38 , travels along second conduit  42  and enters expansion valve  40 . Within expansion valve  40 , pressurized liquid coolant  26 PL is expanded into liquid coolant  26 L. Depending on temperatures, pressures and the fluid, liquid coolant  26 L exiting expansion valve  30  may be combination of vapor and liquid. Liquid coolant  26 L then flows from expansion valve  40  along second conduit  42  into first conduit  24 . Liquid coolant  26 L is introduced back into heat management system  12 C upstream of heat exchanger  20 . Accordingly, coolant  26  for heat management system  12 C is recycled through coolant replenishment system  34 A during periods of less than peak heat load H. The addition of coolant replenishment system  34 A is advantageous for aircraft  10  where tank  16  cannot provide all coolant  26  necessary for an entire mission. 
       FIG. 5  is a block diagram showing coolant replenishment system  34 B including engine  44 , heat exchanger  46 , and water extractor  48  located in flow series on third conduit  50 . Also depicted are tank  16 , coolant  26 , drain  28 , bleed air  52 , cooling source  54 , and ECS  56 . In coolant replenishment system  34 B, water is extracted from engine  44  bleed air  52  to replenish a volume of coolant  26  in tank  16 . Coolant replenishment system  34 B is an optional addition to any of heat management systems  12 A- 12 C described above, and is most advantageous at low altitudes and high humidity. 
     During flight, bleed air  52  is extracted from engine  44  of aircraft  10  and flows into third conduit  50 . Bleed air  52  flows along third conduit  50  to heat exchanger  46 . Within heat exchanger  46 , bleed air  52  rejects heat into (i.e. is cooled by) ram air, cockpit exhaust, or heat load exhaust flowing into heat exchanger  46  from cooling source  54 . Cooled bleed air  52  exits heat exchanger  46  and flows along third conduit  50  to water extractor  48 . Within water extractor  48 , condensation from cooled bleed air  52  is collected and directed to tank  26  as liquid coolant  26 L for storage and use in one of heat management systems  12 A- 12 C. Dehumidified bleed air  52  exits water extractor  48  and is directed for use by another system of aircraft  10  that can benefit from reduced humidity bleed air, such as ECS  56 . Inclusion of coolant replenishment system  34 B on aircraft  10  allows for in-flight replenishment of coolant  26  in tank  16 . 
       FIG. 6  is a block diagram showing the coolant replenishment system  34 C including engine  44  and ECS  56 . Also depicted are tank  16 , coolant  26 , drain  28 , and bleed air  52 . In coolant replenishment system  34 C, water is extracted from ECS  56  to replenish coolant  26  in tank  16 . Coolant replenishment system  34 C is an optional addition to any of heat management systems  12 A- 12 C. 
     During flight, bleed air  52  is extracted from engine  44  of aircraft  10  for use by ECS  56 . When operating in humid environments, ECS  56  creates excess water, which is commonly sprayed on ECS  56  heat exchangers. Water from ECS  56 , however, can be collected and directed into tank  26  as liquid coolant  26 L for storage and use by one of heat management systems  12 A- 12 C. Bleed air  52  exits ECS  56  and is directed for use by another system of aircraft  10  that can benefit from reduced humidity bleed air  52  (e.g. cockpit, heat loads). Inclusion of coolant replenishment system  34 C on aircraft  10  allows for in-flight replenishment of coolant  26  in tank  16 . 
       FIG. 7  is a block diagram showing the coolant replenishment system  36 D including engine  44 , heat exchanger  58 , and filter  60 . Also depicted are tank  16 , coolant  26 , drain  28 , cooling source  54 , and exhaust air  62 . In coolant replenishment system  34 C, water is extracted from exhaust air  62  of engine  44  to replenish coolant  26  in tank  16 . Coolant replenishment system  34 D is an optional addition to any of heat management systems  12 A- 12 C. 
     During flight, core exhaust air  62  is collected from engine  44  after the low pressure turbine. Exhaust air  62  is ducted to heat exchanger  58  for cooling. Within heat exchanger  58 , exhaust air  62  rejects heat into (i.e. is cooled by) ram air, or cockpit exhaust, or heat load exhaust supplied from cooling source  54  to heat exchanger  58 . While exhaust air  62  is being cooled in heat exchanger  58 , water is condensed and extracted from exhaust air  62 . The extracted water is sent through filter  60  to remove unwanted combustion byproducts. The filtered water is then sent to tank  16  as liquid coolant  26 L for storage and use by one of heat management systems  12 A- 12 C. Dehumidified exhaust air  62  exits heat exchanger  58  and is sent overboard. Inclusion of coolant replenishment system  34 D on aircraft  10  allows for in-flight replenishment of coolant  26  in tank  16 . 
     While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.