Abstract:
An aircraft pressurization system, includes an auxiliary compressor for further compressing compressed air received from a low pressure compressor section of a gas turbine engine while the compressed air is below a predetermined pressure level; a bleed passage for fluidically connecting the auxiliary compressor to the low pressure compressor section; and an environmental control system coupled to an output of the auxiliary compressor for conditioning the compressed air to a predetermined level.

Description:
FIELD OF INVENTION 
       [0001]    The present invention relates to gas turbine engine bleed air, and in particular to the use of low-pressure compressor bleed air for an aircraft pressurization system that is extracted from a gas turbine engine compressor and augmented by an auxiliary compressor. 
       DESCRIPTION OF RELATED ART 
       [0002]    In a typical gas turbine engine, a compressor compresses air and passes that air along a primary flow path to a combustor where it is mixed with fuel and combusted. The combusted mixture expands and is passed to a turbine, which is forced to rotate. When used on an aircraft, the primary purpose of this system is to provide propulsive force for the aircraft. 
         [0003]    In some gas turbine engines, a portion of the air compressed by the compressor is diverted from the primary flow path to a bleed inlet of a bleed air system. This bleed air can be used for a variety of purposes, such as to de-ice a wing or to provide pressurized air to a cabin of the aircraft. Because the bleed air is often at an undesirably high temperature, a heat exchanger is used to cool the bleed air. Bleeding off and cooling compressed air typically does not generate thrust or useful work, thus reducing the efficiency of the compressor and the entire gas turbine engine. Moreover, the heat exchanger takes up a relatively large amount of space and can increase the overall weight of the bleed air system. 
       BRIEF SUMMARY 
       [0004]    According to one aspect of the invention, an aircraft pressurization system includes an auxiliary compressor for further compressing compressed air received from a low pressure compressor section of a gas turbine engine while the compressed air is below a predetermined pressure level; a bleed passage for fluidically connecting the auxiliary compressor to the low pressure compressor section; and an environmental control system coupled to an output of the auxiliary compressor for conditioning the compressed air to a predetermined level. 
         [0005]    According to another aspect of the invention, a method for pressurizing an aircraft includes receiving air compressed to a first pressure via a low pressure compressor section of a gas turbine engine; compressing, via an auxiliary compressor, the compressed air to a second pressure while the compressed air is below a predetermined pressure level; fluidically connecting, via a bleed passage, the auxiliary compressor to the low pressure compressor section; and conditioning the compressed air to a predetermined level via an environmental control system coupled to the auxiliary compressor. 
         [0006]    Other aspects, features, and techniques of the invention will become more apparent from the following description taken in conjunction with the drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
         [0007]    Referring now to the drawings wherein like elements are numbered alike in the FIGURES: 
           [0008]      FIG. 1  illustrates a schematic view of a gas turbine engine having a low pressure compressor exit bleed system according to an embodiment of the invention; and 
           [0009]      FIG. 2  illustrates a schematic view of a gas turbine engine having a gearbox assembly according to an embodiment of the invention. 
       
    
    
     DETAILED DESCRIPTION 
       [0010]    Embodiments of an aircraft pressurization system include a bleed energy system for extracting bleed air from a single bleed port at a low pressure compressor (“LPC”) during all but the descent segment of an aircraft&#39;s flight and an auxiliary compressor for augmenting the aircraft pressurization system during the descent segment of the flight. Further embodiments are discussed below in detail. In one embodiment, the LPC bleed air provides adequate pressurization during the cruising segment while the auxiliary compressor conditions the LPC bleed air for adequate cabin pressurization during the descent segment. 
         [0011]    Referring now to  FIG. 1  an example of a gas turbine engine  10  coupled to a bleed energy system  12  is illustrated. The gas turbine engine  10  includes a main compressor section  14 , a main combustor section  16 , and a main turbine section  18  arranged in a serial, axial flow relationship. The main compressor section  14  creates and provides compressed air that passes into the combustor section  16  where fuel is introduced and the mixture of fuel and compressed air is burned, generating hot combustion gases. The hot combustion gases are discharged to the main turbine section  18  where they are expanded to extract energy therefrom. Further, the gas turbine engine  10  includes a low pressure spool  20  including a low pressure compressor (“LPC”)  22  and low pressure turbine  24  connected by low pressure shaft  26 , and a high pressure spool  28  having a high pressure compressor  30  and high pressure turbine  32  connected by high pressure shaft  34 , each extending from main compressor section  14  to main turbine section  18 . Air flows from the main compressor section  14  to the main turbine section  18  along main flow path  36 . The engine  10  incorporates a bleed energy system  12  for extracting compressed bleed air from a bleed port  44  connected to the LPC  22  in order to supply LPC bleed air to a cabin. In one embodiment, the LPC bleed air is used by an environmental control system (ECS)  38  to pressurize the cabin of an aircraft. In other embodiments, the LPC bleed air may be used for anti-icing or deicing, heating or cooling, and/or operating pneumatic equipment. It is to be appreciated that a plurality of bleed ports, such as bleed port  44 , may be connect to the LPC  22  in order to supply LPC bleed air to the ECS  38  or to other components if needed. 
         [0012]    Also shown in  FIG. 1 , the bleed energy system  12  includes a bleed passage  40  coupled to a shut-off valve  46  and an auxiliary compressor  42 . Compressed air from LPC  22  is extracted from bleed valve  44 , passes through bleed passage  40  and to auxiliary compressor  42  through shut-off valve  46 . In one embodiment, shut-off valve  46  is selectively opened or closed to control the bleed air flow rate to the auxiliary compressor  42 . In some situations, passing bleed air through auxiliary compressor  42  would reduce its temperature and pressure below desirable levels, such as when the engine  10  is operating at relatively low speeds and in particularly cold environments. In these situations, some or all of the bleed air can be diverted at shut-off valve  46 , passed through bleed passage  60  and returned back to bleed passage  52 . The bleed passage  40  provides a source of high-pressure engine air for pressurized air that is ultimately delivered to the ECS  38  by bleed passage  52  during the cruising segment of the aircraft&#39;s flight so as to provide pressurization during the longest segment of the flight when engine  10  efficiency is critical. 
         [0013]    As illustrated, the auxiliary compressor  42  is mechanically connected to a motor  48  via shaft  50 . The auxiliary compressor  42 , powered by the aircraft electricity source (not shown), augments the compressed LPC bleed air when the bleed air cannot provide adequate pressurization to the ECS  38 . It is to be appreciated that air entering bleed passage  40  is at a pressure and temperature substantially higher than what is needed by ECS  38 . In one embodiment, the minimum bleed air pressure is at 20 psi (137.9 kPa) in order for ECS  38  to maintain cabin pressure to 11.8 psi (81.4 kPa) and provide fresh air at 0.55 Pounds Mass/Minute/Person. 
         [0014]    In operation, LPC bleed air is extracted through bleed valve  44  and fluidically communicated to auxiliary compressor  42  through bleed passage  40  to provide aircraft pressurization during all segments of flight. Shut-off valve  46  may be selectively opened or closed to control the bleed air flow rate to the auxiliary compressor  42 . The LPC bleed air enters into an inlet of auxiliary compressor  42 , and passes out an outlet of auxiliary compressor  42  into bleed passage  52  and into ECS  38 . According to one embodiment, during the cruising segment of the flight, the engine  10  provides all of the LPC compressed air for pressurization of the aircraft&#39;s cabin. In this case, the LPC bleed air is extracted from low pressure compressor  22  and flows through the auxiliary compressor  42  without substantial change to its pressure or temperature. In another embodiment, the auxiliary compressor  42  adds energy to the LPC bleed air to increase pressure and temperature to suitable levels below a certain threshold before passing the conditioned LPC bleed air to the ECS  38 . In one or more embodiments, heat exchangers may be positioned along bleed passage  40  or bleed passage  52  in order to lower the temperature (i.e., remove energy) from the LPC bleed air. 
         [0015]    During the descent segment of flight, the auxiliary compressor  42  augments the compressed LPC bleed air when the LPC bleed air cannot provide adequate compressed bleed air for pressurization by the ECS  38 . In particular, LPC bleed air from the low pressure compressor  22  is further compressed with the auxiliary compressor  42  in order to condition the LPC bleed air to minimum levels before communicating the compressed air to the ECS  38 . The auxiliary compressor  42  is mechanically connected to and is driven by motor  48  in order to compress the extracted air from the low-pressure compressor  22 . In other embodiments, the motor  48  is powered by electricity from the aircraft, or may be coupled to a gear box ( FIG. 2 ) that is connected to and driven by the high pressure spool  28 , or by a bleed powered boost compressor. 
         [0016]    In an embodiment, illustrated in  FIG. 2 , the auxiliary compressor  42  is mechanically connected to a gearbox  54 . Particularly, the auxiliary compressor  42  is mechanically connected to and driven by a gearbox  54  via a shaft  58  in order to compress the extracted air from the low-pressure compressor  22 , while all other aspects remain substantially the same as those of gas turbine engine  10  and bleed energy system  12  shown and illustrated in  FIG. 1 . The gearbox  54  is connected to a high pressure spool  28 . The rotating spool  28  correspondingly controls gearbox  54  via bleed line  56  and causes the gearbox  54  to control the rotation of the shaft  58  and drive the auxiliary compressor  42  in order to compress the extracted air from the low pressure compressor  22 . Also, bleed passage  62  is provided to divert some or all of the bleed air from auxiliary compressor  42  into bleed passage  62  and returned back to bleed passage  64  when passing bleed air through auxiliary compressor  42  would reduce its temperature and pressure below desirable levels. 
         [0017]    The technical effects and benefits of exemplary embodiments include an aircraft pressurization system with only one engine bleed port located at the exit of the low pressure compressor for providing adequate pressurization during all segments of flight except descent. For the descent flight segment, LPC bleed air is compressed to the required pressure by an electric, gearbox mounted, or bleed powered boost compressor. 
         [0018]    The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention. While the description of the present invention has been presented for purposes of illustration and description, it is not intended to be exhaustive or limited to the invention in the form disclosed. Many modifications, variations, alterations, substitutions, or equivalent arrangement not hereto described will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the invention. Additionally, while various embodiment of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.