Patent Publication Number: US-2016237907-A1

Title: Intercooled cooling air with auxiliary compressor control

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application is a continuation-in-part of U.S. patent application Ser. No. 14/695,578 (filed on Apr. 24, 2015 and entitled “Intercooled Cooling Air”) and claims priority to U.S. Provisional Patent Application No. 62/115578, filed 12 Feb. 2015. 
    
    
     BACKGROUND 
     This application relates to improvements in providing cooling air from a compressor section to a turbine section in a gas turbine engine. 
     Gas turbine engines are known and typically include a fan delivering air into a bypass duct as propulsion air. Further, the fan delivers air into a compressor section where it is compressed. The compressed air passes into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate. 
     It is known to provide cooling air from the compressor to the turbine section to lower the operating temperatures in the turbine section and improve overall engine operation. Typically, air from the high compressor discharge has been tapped, passed through a heat exchanger, which may sit in the bypass duct and then delivered into the turbine section. The air from the downstream most end of the compressor section is at elevated temperatures. 
     SUMMARY 
     In a featured embodiment, a gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passing the tapped air through a heat exchanger and then to a cooling compressor, which compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. The cooling compressor rotates at a speed proportional to a speed of at least one rotor in the turbine section. The cooling compressor is allowed to rotate at a speed that is not proportional to a speed of the at least one rotor under certain conditions. 
     In another embodiment according to the previous embodiment, an input shaft to the cooling compressor is driven at a speed proportional to the at least one turbine rotor. A clutch disconnects the input shaft from a rotor of the cooling compressor should the input shaft experience a decrease in speed. 
     In another embodiment according to any of the previous embodiments, a one-way ratchet is positioned between the input shaft and the rotor. 
     In another embodiment according to any of the previous embodiments, a flywheel rotates with the rotor, such that the compressor maintains a rotational speed after the clutch opens and allows the rotor to rotate independently of the input shaft. 
     In another embodiment according to any of the previous embodiments, a main fan delivers bypass air into a bypass duct and into the main compressor section. The heat exchanger is positioned within the bypass duct to be cooled by bypass air. 
     In another embodiment according to any of the previous embodiments, the cooling compressor includes a centrifugal compressor impeller. 
     In another embodiment according to any of the previous embodiments, air temperatures at the downstream most location of the high pressure compressor are greater than or equal to 1350° F. 
     In another embodiment according to any of the previous embodiments, the turbine section drives a bull gear, which further drives an impeller of the cooling compressor. 
     In another embodiment according to any of the previous embodiments, the bull gear also drives an accessory gearbox. 
     In another embodiment according to any of the previous embodiments, a gear ratio multiplier is included such that the impeller rotates at a faster speed than the tower shaft. 
     In another embodiment according to any of the previous embodiments, the impeller is a centrifugal compressor impeller. 
     In another embodiment according to any of the previous embodiments, the impeller is a centrifugal compressor impeller. 
     In another embodiment according to any of the previous embodiments, an auxiliary fan is positioned upstream of the heat exchanger. 
     In another embodiment according to any of the previous embodiments, the auxiliary fan operates at a variable speed. 
     In another featured embodiment, an intercooling system for a gas turbine engine comprises a heat exchanger for cooling air drawn from a portion of a main compressor section at a first temperature and pressure for cooling the air to a second temperature cooler than the first temperature. A cooling compressor compresses air communicated from the heat exchanger to a second pressure greater than the first pressure and communicates the compressed air to a portion of a turbine section. The cooling compressor rotates at a speed proportional to a speed of at least one rotor in the turbine section. The cooling compressor is allowed to rotate at a speed that is not proportional to a speed of the at least one rotor under certain conditions. 
     In another embodiment according to the previous embodiment, an input shaft to the cooling compressor is driven at a speed proportional to the at least one turbine rotor. A clutch disconnects the input shaft from a rotor of the cooling compressor should the input shaft experience a decrease in speed. 
     In another embodiment according to any of the previous embodiments, a one-way ratchet is positioned between the input shaft and the rotor. 
     In another embodiment according to any of the previous embodiments, a flywheel rotates with the rotor, such that the compressor maintains a rotational speed after the clutch opens and allows the rotor to rotate independently of the input shaft. 
     In another embodiment according to any of the previous embodiments, a bull gear drives an impeller of the cooling compressor. 
     In another featured embodiment, a gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A low pressure compressor provides some of the more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and delivers air into the high pressure turbine. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  schematically shows an embodiment of a gas turbine engine. 
         FIG. 2  shows a prior art engine. 
         FIG. 3  shows one example engine. 
         FIG. 4  is a graph illustrating increasing temperatures of a tapped air against the work required. 
         FIG. 5  shows a detail of an example of an engine. 
         FIG. 6  shows a further detail of the example engine of  FIG. 5 . 
         FIG. 7  shows a further feature of the example engine of  FIG. 5 . 
         FIG. 8  shows a graph associated with the example engine of  FIG. 5 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
     Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
     The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that connects a fan  42  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan  42  through a speed change device, such as a geared architecture  48 , to drive the fan  42  at a lower speed than the low speed spool  30 . The high-speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
     A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . In another example, the high pressure turbine  54  includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The example low pressure turbine  46  has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. 
     A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the low pressure turbine  46 . 
     Airflow through the core airflow path C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes vanes  60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  58 . Reducing or eliminating the number of vanes in the low pressure turbine  46  shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
     The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3. 
     In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point. 
     “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. 
     “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
     The example gas turbine engine includes the fan  42  that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section  22  includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about 6 turbine rotors schematically indicated at  34 . In another non-limiting example embodiment the low pressure turbine  46  includes about 3 turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     Gas turbine engines designs are seeking to increase overall efficiency by generating higher overall pressure ratios. By achieving higher overall pressure ratios, increased levels of performance and efficiency may be achieved. However, challenges are raised in that the parts and components associated with a high pressure turbine require additional cooling air as the overall pressure ratio increases. 
     The example engine  20  utilizes air bleed  80  from an upstream portion of the compressor section  24  for use in cooling portions of the turbine section  28 . The air bleed is from a location upstream of the discharge  82  of the compressor section  24 . The bleed air passes through a heat exchanger  84  to further cool the cooling air provided to the turbine section  28 . The air passing through heat exchanger  84  is cooled by the bypass air B. That is, heat exchanger  84  is positioned in the path of bypass air B. 
     A prior art approach to providing cooling air is illustrated in  FIG. 2 . An engine  90  incorporates a high pressure compressor  92  downstream of the low pressure compressor  94 . As known, a fan  96  delivers air into a bypass duct  98  and into the low pressure compressor  94 . A downstream most point, or discharge  82  of the high pressure compressor  92  provides bleed air into a heat exchanger  93 . The heat exchanger is in the path of the bypass air in bypass duct  98 , and is cooled. This high pressure high temperature air from location  82  is delivered into a high pressure turbine  102 . 
     The downstream most point  82  of the high pressure compressor  82  is known as station  3 . The temperature T3 and pressure P3 are both very high. 
     In future engines, T3 levels are expected to approach greater than or equal to 1350° F. Current heat exchanger technology is becoming a limiting factor as they are made of materials, manufacturing, and design capability which have difficulty receiving such high temperature and pressure levels. 
       FIG. 3  shows an engine  100  coming within the scope of this disclosure. A fan  104  may deliver air B into a bypass duct  105  and into a low pressure compressor  106 . High pressure compressor  108  is positioned downstream of the low pressure compressor  106 . A bleed  110  taps air from a location upstream of the downstream most end  82  of the high pressure compressor  108 . This air is at temperatures and pressures which are much lower than T3/P3. The air tapped at  110  passes through a heat exchanger  112  which sits in the bypass duct  105  receiving air B. Further, the air from the heat exchanger  112  passes through a compressor  114 , and then into a conduit  115  leading to a high turbine  117 . This structure is all shown schematically. 
     Since the air tapped at point  110  is at much lower pressures and temperatures than the  FIG. 2  prior art, currently available heat exchanger materials and technology may be utilized. This air is then compressed by compressor  114  to a higher pressure level such that it will be able to flow into the high pressure turbine  117 . 
     An auxiliary fan  116  may be positioned upstream of the heat exchanger  112  as illustrated. The main fan  104  may not provide sufficient pressure to drive sufficient air across the heat exchanger  112 . The auxiliary fan will ensure there is adequate air flow in the circumferential location of the heat exchanger  112 . 
     In one embodiment, the auxiliary fan may be variable speed, with the speed of the fan varied to control the temperature of the air downstream of the heat exchanger  112 . As an example, the speed of the auxiliary fan may be varied based upon the operating power of the overall engine. 
     Referring to  FIG. 4 , a temperature/entropy diagram illustrates that a lower level of energy is spent to compress air of a lower temperature to the desired P3 pressure level. Cooler air requires less work to compress when compared to warmer air. Accordingly, the work required to raise the pressure of the air drawn from an early stage of the compressor section is less than if the air were compressed to the desired pressure within the compressor section. Therefore, high pressure air at P3 levels or higher can be obtained at significantly lower temperatures than T3. As shown in  FIG. 4 , to reach a particular pressure ratio, 50 for example, the prior system would move from point 2 to point 3, with a dramatic increase in temperature. However, the disclosed or new system moves from point  2  to point 5 through the heat exchanger, and the cooling compressor then compresses the air up to point 6. As can be appreciated, point 6 is at a much lower temperature. 
       FIG. 5  shows a detail of compressor  114  having an outlet into conduit  115 . A primary tower shaft  120  drives an accessory gearbox  121 . The shaft  126  drives a compressor rotor within the compressor  114 . The shafts  120  and  126  may be driven by a bull gear  125  driven by a turbine rotor, and in one example, with a high pressure compressor rotor. 
       FIG. 6  shows an example wherein a gear  128  is driven by the shaft  126  to, in turn, drive a gear  130  which drives a compressor impeller  129 . An input  132  to the compressor impeller  129  supplies the air from the tap  110 . The air is compressed and delivered into the outlet conduit  115 . 
     By providing a gear ratio multiplier between the compressor impeller  129  and the high spool bull gear  125 , the compressor impeller may be driven to operate an optimum speed. As an example, the gear ratio increase may be in a range of 5:1-8:1, and in one embodiment, 6:1. 
     Details of the engine, as set forth above, may be found in co-pending U.S. patent application Ser. No. 14/695,578, which is incorporated herein by reference in its entirety. 
       FIG. 7  shows an embodiment that addresses a challenge with the above-referenced embodiment. Since the cooling or auxiliary compressor  114  is driven at a speed proportional to the speed of at least one turbine section, it will rotate at a speed proportional to the associated compressor section also. 
     At times, the overall gas turbine engine may have a very immediate deceleration in speed. If the auxiliary compressor  114  also dramatically reduces speed, there can be stability issues, and the auxiliary compressor could experience a phenomenon called surge. This would be undesirable. 
     In  FIG. 7 , an embodiment  150  includes an input shaft  152 , which could be the shaft— 130  as shown in  FIG. 6 . Shaft  152  drives a one-way clutch, such as ratchet clutch  154 . The compressor  156  rotates with a flywheel  158 . 
     Now, consider a point of operation where the system  150  has been driven at a first speed for a period of time. The compressor rotor  156  will be rotating at the speed, and the flywheel  158  will also be rotating at the speed. 
     Flywheels are known, and typically include a mass that will continue to rotate the rotor even if the input shaft slows. Essentially, a flywheel serves as a means of storing rotational energy. 
     Should the input  152  now suddenly decelerate, the one-way ratchet  154  will allow the compressor rotor  156  and flywheel  158  to rotate independently of the input  152 . Essentially, the ratchet will slip relative to the input  152 , and the rotor  156  and flywheel  158  can rotate at a higher speed than the input  152  for a period of time. This will allow the auxiliary compressor to avoid a surge condition, and more gradually approach the speed of the remainder of the engine. 
     This is illustrated in  FIG. 8 . An engine high spool is shown with a sudden deceleration in speed. However, the auxiliary compressor will have its speed deteriorate on a much slower slope. 
     While a one-way ratchet and flywheel are shown, more generally, it could be said that the auxiliary compressor is driven at a speed proportional to a turbine speed, and there is a method of allowing the auxiliary compressor to move away from a speed proportional to the speed of the turbine under certain conditions. In embodiments, that would be the one-way ratchet  154  and flywheel  158 , however, other ways of achieving such a disconnect could come within the scope of this invention. 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.