Patent Abstract:
An all electric air conditioning system for an aircraft, wherein the aircraft defines an interior volume having conditioned air at a first pressure. A compressor is provided and is operable to compress supply air to a second pressure. The compressor being operated in response to an electrical drive motor. A passage fluidly couples the compressor and the interior volume of the aircraft. A heat dissipating device, such as a heat exchanger, is positioned in the passage to extract heat energy from the supply air. This arrangement permits conditioning of air within the aircraft without using bleed air from the engines. The use of bleed air results in a significant amount of fuel burn. An optional conditioned air recovery system may be coupled to the interior volume of the fuselage to direct at least a portion of the conditioned air from the interior volume back for further conditioning and use.

Full Description:
FIELD OF THE INVENTION  
         [0001]    The present invention generally relates to air conditioning systems and, more particularly, to an electrically driven air conditioning system for an aircraft that does not rely on engine bleed air.  
         BACKGROUND OF THE INVENTION  
         [0002]    Many air conditioning systems employed in modern commercial aircraft utilize the air-to-air thermodynamic cycle to provide cooling and/or heating air to the various compartments on the aircraft, such as the passenger cabin, cargo holds, and the like. Air from the compressor stages of the main jet propulsion engines, also known as “bleed air,” is generally output at high temperature and pressure (i.e.  610 OF and  60  psi). Conventionally, this bleed air is then conditioned through conditioning packs before passing into the pressurized fuselage for cabin temperature control, ventilation, and pressurization. This conditioned air within the fuselage is then discharged to the outside ambient air through various overboard valves, overflow valves, and cabin leaks.  
           [0003]    This known method of conditioning air for use with the various aircraft systems is inefficient. That is, during a typical steady state cruise operation, more energy than is necessary for the primary requirements of the conditioning system (e.g. cabin temperature control, ventilation, and pressurization) is added into the conditioning system at the engines in the form of additional fuel. Much of this excess energy is wasted in the form of heat and pressure drop through ductwork, valves, and various other components of the conditioning system. Moreover, extracting work from the engines in the form of bleed air is inefficient relative to other extraction methods. Consequently, the use of bleed air from the engines reduces the efficiency of the engines and, thus, increases the fuel consumption and load on the engines. By eliminating or at least minimizing the use of bleed air in the various aircraft systems, it is believed that more efficient jet engines may be developed. Moreover, it is believed that alternative air conditioning systems may lead to a reduction in aircraft weight, assembly complexity, and fuel consumption.  
           [0004]    Accordingly, there exists a need in the relevant art to provide an air conditioning system for an aircraft that does not rely on jet engine bleed air for operation. Furthermore, there exists a need in the relevant art to provide an air conditioning system for an aircraft that is capable of reducing the aircraft weight, assembly complexity, and fuel consumption. Still further, there exists a need in the relevant art to provide an air conditioning system for an aircraft driven by electrical energy. Moreover, there exists a need in the relevant art to provide an air conditioning system for an aircraft that overcomes the disadvantages of the prior art.  
         SUMMARY OF THE INVENTION  
         [0005]    An all electric air conditioning system for an aircraft, wherein the aircraft defines an interior volume having conditioned air at a first pressure, is provided having an advantageous construction. A compressor is provided and is operable to compress supply air to a second pressure. The compressor being operated in response to an electrical drive motor. A passage fluidly couples the compressor and the interior volume of the aircraft. A heat-dissipating device, such as a heat exchanger, is positioned in the passage to extract heat energy from the supply air. This arrangement permits conditioning of air within the aircraft without using bleed air from the engines. The use of bleed air results in a significant amount of fuel burn. An optional conditioned air recovery system may be coupled to the interior volume of the fuselage to direct at least a portion of the conditioned air from the interior volume back for further conditioning and use.  
           [0006]    Further areas of applicability of the present invention will become apparent from the detailed description provided hereinafter. It should be understood that the detailed description and specific examples, while indicating the preferred embodiment of the invention, are intended for purposes of illustration only and are not intended to limit the scope of the invention. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0007]    The present invention will become more fully understood from the detailed description and the accompanying drawings, wherein:  
         [0008]    [0008]FIG. 1 is a circuit diagram illustrating a first embodiment of the present invention in a ground or low altitude operation configuration;  
         [0009]    [0009]FIG. 2 is a circuit diagram illustrating the first embodiment of the present invention in a cruise operation configuration;  
         [0010]    [0010]FIG. 3 is a circuit diagram illustrating a second embodiment of the present invention in a ground or low altitude operation configuration;  
         [0011]    [0011]FIG. 4 is a circuit diagram illustrating the second embodiment of the present invention in a cruise operation configuration;  
         [0012]    [0012]FIG. 5 is a circuit diagram illustrating a third embodiment of the present invention in a ground or low altitude operation configuration;  
         [0013]    [0013]FIG. 6 is a circuit diagram illustrating the third embodiment of the present invention in a cruise operation configuration; and  
         [0014]    [0014]FIG. 7 is a circuit diagram illustrating various alternative modifications of the present invention. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
       [0015]    The following description of the preferred embodiments is merely exemplary in nature and is in no way intended to limit the invention, its application, or uses. That is, the electrical air conditioning system of the present invention may find utility in other applications, which commonly use bleed air to drive an air conditioning system, such as in tanks and the like.  
         [0016]    According to a first embodiment of the present invention, an air conditioning system  10  is illustrated that is capable of eliminating the use of bleed air. Air conditioning system  10  comprises a first compressor  12 , a second compressor  14 , a first turbine  16 , a second turbine  18 , a heat exchanger assembly  20 , a reheater  22 , a condenser  24 , and a high-pressure water separator  26 . With particular reference to FIG. 1, a ram air fan  28  provides cold side air for heat exchanger assembly  20  during ground operation. With particular reference to FIG. 2, a ram air inlet scoop (not shown) provides cold side air for heat exchanger assembly  20  during in-flight operation. Ram air, generally indicated as  100 , is provided to first compressor  12  via a passage  30  and further to heat exchanger assembly  20 . A valve  62  controls the amount of ram air  100  that is directed to heat exchanger assembly  20 . Alternatively, engine fan air could be used to provide cool side air to first compressor  12  in place of ram air  100 .  
         [0017]    First compressor  12  is fluidly coupled to a primary heat exchanger  32  of heat exchanger assembly  20  via a passage  34 . Primary heat exchanger  32  in turn is fluidly coupled to second compressor  14  via a passage  36 . Second compressor  14  in turn is fluidly coupled to a secondary heat exchanger  38  via a passage  40 . Secondary heat exchanger  38  is fluidly separate from primary heat exchanger  32 . Hence, it should be understood that primary heat exchanger  32  and secondary heat exchanger  38  might be configured as separate units or a single unit having multiple discrete chambers.  
         [0018]    A compressor bypass valve  42  may fluidly interconnect passage  36  and passage  40  so as to permit bypassing of second compressor  14 . Furthermore, an ozone converter  44  may be positioned in series within passage  40  to permit proper conversion of ambient air during a cruise phase of flight.  
         [0019]    As seen in FIGS. 1 and 2, heat exchanger assembly  20  also employs ram air  100  acting as a heat sink to remove excess heat from the air upon exit from first compressor  12  and again upon exit from second compressor  14 . Trim air  46  may be extracted from passage  40  for use in individual compartment temperature control or for use in other aircraft systems.  
         [0020]    Secondary heat exchanger  38  is fluidly coupled to reheater  22  via a passage  48 . The cold outlet of reheater  22  is directed to first turbine  16  through a passage  50  so as to be expanded and reduced in temperature therein. This air is then directed into the cold inlet of condenser  24  via a passage  52 . It should be appreciated that the cold inlet side of condenser  24  is maintained above freezing to prevent ice formation. The air from condenser  24  is then directed to second turbine  18  via a passage  54  for final expansion. Finally, air exits second turbine  18  via passage  56  and is directed to a mix manifold (not shown) for distribution into the aircraft cabin.  
         [0021]    Still referring to FIGS. 1 and 2, air conditioning system  10  further includes a passage  58  fluidly interconnecting condenser  24  to water collector  26  and a passage  60  fluidly interconnecting water collector  26  to reheater  22 . A passage  61  is further provided that fluidly interconnects reheater  22  and condenser  24 . Air conditioning system  10  still further includes a ram air modulator valve/actuator  62  used for controlling the flow of ram air  100 . A turbine bypass valve  64  fluidly interconnects passage  54  and passage  56  so as to permit bypassing of second turbine  18 . A first motor  66  is operably coupled between first compressor  12  and second turbine  18  and a second motor  68  is operably coupled between second compressor  14  and first turbine  16 .  
         [0022]    With particular reference to FIG. 2, it can be seen that air conditioning system  10  further includes an altitude valve  70  fluidly interconnecting passage  48  and passage  54 . Altitude valve  70  permits bypassing of a condensing loop  72  above a predetermined altitude. Condensing loop  72  generally includes reheater  22 , passage  50 , first turbine  16 , passage  52 , condenser  24 , passage  58 , water collector  26 , passage  60 , and passage  61 . The opening of altitude valve  70  bypasses condensing loop  72  such that primary cooling of the air occurs in heat exchanger assembly  20  and second turbine  18 . This arrangement at altitude enables the overall pressure drop in the system to be minimized so as to provide sufficient flow to the passenger cabin at lower power consumption levels.  
         [0023]    Referring to FIGS. 1 and 2, during operation, first compressor  12  receives ambient air  100  from ram air fan  28 . This air is compressed within first compressor  12  and is passed through primary heat exchanger  32  of heat exchanger assembly  20  to second compressor  14 . Primary heat exchanger  32  removes heat from the air using ram air  100  as a heat sink. The air is then compressed within second compressor  14  and passed through secondary heat exchanger  38  of heat exchanger assembly  20 . Second compressor  14  may be bypassed using compressor bypass valve  42 .  
         [0024]    During ground or low altitude operation, air then exits secondary heat exchanger  38  and is directed to reheater  22 . The cold outlet of reheater  22  directs air to first turbine  16  where the temperature and pressure are reduced. The air is then directed to condenser  24  to remove excess water from the air. The cold outlet of condenser  24  directs the air to second turbine  18  where the temperature and pressure are further reduced. Lastly, the air is then directed to the mixing manifold and distributed to the aircraft cabin.  
         [0025]    Air and water from condenser  24  flows to water collector  26  through passage  58 , where water is collected by water collector  26 .  
         [0026]    During high altitude operation, air from secondary heat exchanger  38  of heat exchanger assembly  20  is directed through altitude valve  70  so as to completely bypass condensing loop  72 . Accordingly, air flows from secondary heat exchanger  38  directly to second turbine  18  so as to minimize the pressure drop within system  10  during high altitude cruise. Therefore, power consumption is minimized.  
         [0027]    According to a second embodiment of the present invention, an air conditioning system  10 ′ is illustrated that is capable of eliminating the use of bleed air and further capable of utilizing the potential energy of pressurized air leaving the aircraft cabin during high altitude flight.  
         [0028]    With particular reference to FIGS. 3 and 4, in addition to those elements described in reference to FIGS. 1 and 2, air conditioning system  10 ′ further includes an outflow turbine  110 . Outflow turbine  110  is illustrated as being operably coupled to motor  68  and first turbine  16 . However, it must be understood that outflow turbine  110  may be alternatively coupled to first compressor  12  or second turbine  18 . Outflow turbine  110  receives previously conditioned air from the cabin of the aircraft through a cabin recovery valve  112 . Cabin recovery valve  112  is actuated to provide flow of conditioned air through outflow turbine  110 . It should be understood that cabin recovery valve  112  or outflow turbine  110  might include an integral anti-depressurization valve to guard against inadvertent depressurization of the aircraft cabin. That is, should a duct burst or other failure to occur, anti-depressurization valve will close to prevent further depressurization of the aircraft cabin. The anti-depressurization valve may be a conventional aerodynamic valve that closes upon sensing too much air flow.  
         [0029]    Cabin air  114  is directed through outflow turbine  110  where it is quickly expanded. This expansion of cabin air  114  causes a rapid temperature drop of cabin air  114 , which is directed through passage  116  to heat exchanger assembly  20 . This cooled air serves to supplement ram air  100 , thereby reducing the drag associated with the ram air system by not requiring as much outside ambient air for heat exchanger assembly  20  cooling. Moreover, the power generated by outflow turbine  110  serves to reduce the work required by motor  68  when driving second compressor  14 . A significant electrical power and ram air drag saving is achieved as the cruise phase is the majority of the entire flight.  
         [0030]    Still referring to FIGS. 3 and 4, during ground or low altitude operation, air conditioning system  10 ′ works identically to air conditioning system  10 . However, during high altitude operation, as described above, cabin air  114  is expanded and cooled in outflow turbine  110  and is passed to heat exchanger assembly  20  for cooling. Like air conditioning system  10 , condensing loop  72  is bypassed using altitude valve  70 . The opening of altitude valve  70  bypasses condensing loop  72  such that primary cooling of the air occurs in heat exchanger assembly  20 , supplementing with expanded cabin air  114 , and second turbine  18 . This arrangement, at altitude, enables the overall pressure drop in system  10 ′ to be minimized so as to provide sufficient flow to the passenger cabin at lower power consumption levels.  
         [0031]    According to a third embodiment of the present invention, an air conditioning system  10 ″ is illustrated that is capable of eliminating the use of bleed air and further capable of utilizing the potential energy of pressurized air leaving the aircraft cabin during high altitude flight. However, unlike the second embodiment of the present invention, air conditioning system  10 ″ employs a series of control valves such that first turbine  16  acts similar to outflow turbine  110  of the second embodiment.  
         [0032]    More particularly, as best seen in FIGS. 5 and 6, in addition to those elements described in reference to FIGS. 1 and 2, air conditioning system  10 ″ further includes a first cabin recovery valve  210 . First turbine  16  receives previously conditioned air  114  from the cabin of the aircraft through first cabin recovery valve  210 . First cabin recovery valve  210  is variably actuated to control the preferred flow of conditioned air  114  into passage  50 . Conditioned air  114  joins air flow within passage  50  and is directed to first turbine  16  where it is expanded and cooled. It should be understood that first cabin recovery valve  210  or first turbine  16  may include an integral anti-depressurization valve to guard against inadvertent depressurization of the aircraft cabin. That is, should a duct burst or other failure to occur, anti-depressurization valve will close to prevent further depressurization of the aircraft cabin. The anti-depressurization valve may be a conventional aerodynamic valve that closes upon sensing too much air flow.  
         [0033]    Air conditioning system  10 ″ further includes a second cabin recovery valve  212  disposed within passage  50  upstream from the inflow of cabin air  114 . Second cabin recovery valve  212  is selectively actuated to prohibit air flow from reheater  22  to first turbine  16  and backflow of cabin air  114  to reheater  22 . A third cabin recovery valve  214  is disposed within a passage  216  interconnecting passage  52  and heat exchanger assembly  20 . A check valve  218  is further disposed in passage  52  downstream from the interconnection with passage  216 . Check valve  218  prevents backflow of air from condenser  24  in the event of a failure of third cabin recovery valve  214 .  
         [0034]    Still referring to FIGS. 5 and 6, during ground or low altitude operation, air conditioning system  10 ″ works identically to air conditioning system  10 . However, during high altitude operation, cabin air  114  is expanded and cooled in first turbine  16  and is passed to heat exchanger assembly  20  for cooling. Like air conditioning system  10 , condensing loop  72  is bypassed using altitude valve  70  and the bypass valves are actuated to direct cabin air  114  to first turbine  16  and heat exchanger assembly  20 . Specifically, first cabin recovery valve  210  is opened to allow flow of cabin air  114  into a passage  220 . Cabin air  114  is then directed to first turbine  16  via passage  50  by closing second cabin recovery valve  212 . Cabin air  114  is then expanded and cooled and used to supplement ram air  100  in heat exchanger assembly  20 . Check valve  218  prevents flow through a failed-open valve  214  to the ram system. The opening of altitude valve  70  bypasses condensing loop  72  such that primary cooling of the air occurs in heat exchanger assembly  20 , supplementing with expanded cabin air  114 , and second turbine  18 . This arrangement, at altitude, enables the overall pressure drop in system  10 ″ to be minimized so as to provide sufficient flow to the passenger cabin at lower power consumption levels.  
         [0035]    In addition to the above embodiments described in detail, there are numerous modifications that are anticipated to further tailor the air conditioning system of the present invention. However, it must be understood that each of the following modifications, although described together, is individually applicable to the above described embodiments. That is, each modification may be employed separately from the remaining modifications, if desired. They are simply being described together here in the interest of brevity.  
         [0036]    Referring to FIG. 7, it should be understood that ram air fan  28  may alternatively be coupled to second compressor  14 , generally indicated at  28 ′. Ram air fan  28 ′ would thus supply ram air to second compressor  14 . Still referring to FIG. 7, primary heat exchanger  32  may be eliminated if it is determined that a two-stage heat exchanger system is not required, thereby generally designated as  20 ′. Similarly, motor  68  may be eliminated if added mechanical input is not required between second compressor  14  and first turbine  16 . Likewise, second turbine  18  may be eliminated if the necessary temperature and pressure are achieved depending on the equipment used and the aircraft requirements. However, it is preferable that if second turbine  18  is eliminated, then turbine bypass valve  64  be similarly eliminated since its use is now defeated. Alternatively, turbine bypass valve  64  may be repositioned between passage  50  and passage  52 , thereby serving to selectively bypass first turbine  16 .  
         [0037]    Existing aircraft require the use of bleed air to operate the aircraft air conditioning system. However, bleed air requires a significant amount of fuel burn where a significant amount of energy is wasted by the processing of the bleed air. Hence, there is a need in modern designs to alleviate the use of bleed air in air conditioning systems. According to the principles of the present invention, an all electrical air conditioning system is provided that eliminates the need for bleed air. Moreover, the present invention enables much of the energy of the conditioned air within the cabin to be recovered, thereby reducing electrical power consumption. The elimination of the use of bleed air enables aircraft engines to be more efficiency designed, thereby reducing the use of fuel. It should be appreciated that extracting electricity from jet engines is much more efficient than extracting bleed air. Still further, the present invention provides a method of reducing the weight and maintenance requirements of the aircraft since engine pneumatic ducting, APU ducting, and pneumatic components are eliminated. Duct leaks may be eliminated or at least reduce while overheat detection systems may no longer be necessary. Additionally, air conditioning systems may be modularized, since they no longer need to be sized relative to APU/Engine pneumatic operation performance.  
         [0038]    The description of the invention is merely exemplary in nature and, thus, variations that do not depart from the gist of the invention are intended to be within the scope of the invention. Such variations are not to be regarded as a departure from the spirit and scope of the invention.

Technology Classification (CPC): 1