Patent Publication Number: US-2020291856-A1

Title: Apparatus for selective delivery of pressurised gas

Description:
CROSS-REFERENCE TO RELATED PATENT APPLICATIONS 
     This application is based upon and claims the benefit of priority from UK Patent Application No. GB 1808659.5, filed on May 25, 2018, the entire contents of which are incorporated by reference. 
     BACKGROUND 
     Technical Field 
     The present disclosure relates to apparatus for selective delivery of pressurised gas, particularly although not exclusively in relation to selective delivery of pressurised gas obtained from the core of a gas turbine engine. 
     Description of the Related Art 
     A gas turbine engine may have one or more bleed valves comprised in a system or arrangement which allows high pressure air to be obtained from the core of the engine as needed in order to operate systems requiring compressed or pressurised air input. In the case of an aero engine, compressed air is needed to operate auxiliary systems of an aircraft comprising the aero engine. 
     Conventionally, a bleed valve is located outside the core (for example it may be attached to the exterior of the core) and is coupled to the interior of the core by a connecting duct or stub pipe coupling an output port of the casing of the core to the input port of the bleed valve. The output port in the core casing may for example be located immediately downstream of one of the compressor stages of the engine. Bleed valves are kept closed when an engine is operated at high thrust (for example during aircraft take-off) in order to protect the auxiliary systems from excessively high pressures and temperatures generated in the engine core. When a bleed valve is closed, the pressure within a connecting duct or stub pipe coupling the engine core to the bleed valve, and within the valve itself up to the sealing line of the valve, is the same as the pressure within the engine core at the position of the output port in the core casing. The bleed valves are opened during phases of the engine cycle when the engine is operated at moderate or low thrust. 
     Such an arrangement is vulnerable to failure by rupturing of the connecting duct or stub pipe and by failure of an interface between one end of the duct or pipe and either the input port of the bleed valve or the output port of the core casing, when the engine is operated at high thrust conditions. In the event of such a failure, hot, high-pressure air escapes from the engine core into the space between the engine core casing and the core cowling, thereby damaging components located in that space. 
     BRIEF SUMMARY 
     According to a first aspect, apparatus for selective delivery of pressurised gas comprises a containment wall which at least partially bounds of a volume of pressurised gas during use of the apparatus, the apparatus further comprising a valve having a valve body containing a valve mechanism, the valve body being directly coupled to an output port of the containment wall such that the containment wall and the valve retain the pressurised gas when the valve is closed. 
     The output port may comprise a portion of the containment wall which extends outwardly from the remainder of the containment wall with respect to the volume of pressurised gas. The output port may further comprise an additional hollow element coupled to the outwardly-extending portion of the containment wall. The cross-sectional area of the output port may increase as a function of distance from a part of the containment wall not comprised in the output port. 
     Alternatively, the output port may comprise a portion of the containment wall which extends inwardly from the remainder of the containment wall with respect to the volume of pressurised gas. 
     Alternatively, the output port may be an aperture in the containment wall. 
     The valve may be located at least partially within the output port. The sealing line of the valve mechanism may be located within the output port. 
     The apparatus may comprise an output duct coupled to the valve body on a side thereof remote from the output port. 
     The containment wall may be a compressor casing wall or a turbine casing wall of a gas turbine engine. 
     According to a second aspect, a gas turbine engine comprises apparatus according to the first aspect. 
     According to a third aspect, an aircraft comprises a gas turbine engine according to the second aspect. 
     Except where mutually exclusive, a feature described in relation to any one of the above aspects may be applied mutatis mutandis to any other aspect. Furthermore except where mutually exclusive any feature described herein may be applied to any aspect and/or combined with any other feature described herein. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       Embodiments are described below by way of example only, with reference to the figures in which: 
         FIG. 1  is a longitudinal section of a known, geared, gas turbine engine; 
         FIG. 2  is a detailed longitudinal section of an upstream portion of the  FIG. 1  engine; 
         FIG. 3  shows a side view of known apparatus comprised in the  FIG. 1  engine for selective delivery of compressed air; 
         FIG. 4  shows a longitudinal section of the  FIG. 3  apparatus; 
         FIG. 5  shows a side view of a first example apparatus for selective delivery of compressed air; 
         FIG. 6  shows a longitudinal section of the  FIG. 5  apparatus; 
         FIG. 7  shows a second example apparatus for selective delivery of compressed air; 
         FIG. 8  shows a third example apparatus for selective delivery of compressed air; 
         FIG. 9  shows a fourth example apparatus for selective delivery of compressed air; 
         FIG. 10  shows a fifth example apparatus for selective delivery of compressed air; 
         FIG. 11  shows a sixth example apparatus for selective delivery of compressed air; and 
         FIG. 12  shows a seventh example apparatus for selective delivery of compressed air. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates a gas turbine engine  10  having a principal rotational axis  9 . The engine  10  comprises an air intake  12  and a propulsive fan  23  that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine  10  comprises a core  11  that receives the core airflow A. The engine core  11  comprises, in axial flow series, a low pressure compressor  14 , a high-pressure compressor  15 , combustion equipment  16 , a high-pressure turbine  17 , a low pressure turbine  19  and a core exhaust nozzle  20 . A nacelle  21  surrounds the gas turbine engine  10  and defines a bypass duct  22  and a bypass exhaust nozzle  18 . The bypass airflow B flows through the bypass duct  22 . The fan  23  is attached to and driven by the low pressure turbine  19  via a shaft  26  and an epicyclic gearbox  30 . 
     In use, the core airflow A is accelerated and compressed by the low pressure compressor  14  and directed into the high pressure compressor  15  where further compression takes place. The compressed air exhausted from the high pressure compressor  15  is directed into the combustion equipment  16  where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines  17 ,  19  before being exhausted through the nozzle  20  to provide some propulsive thrust. The high pressure turbine  17  drives the high pressure compressor  15  by a suitable interconnecting shaft  27 . The fan  23  generally provides the majority of the propulsive thrust. The epicyclic gearbox  30  is a reduction gearbox. 
     An exemplary arrangement for a geared fan gas turbine engine  10  is shown in  FIG. 2 . The low pressure turbine  19  (see  FIG. 1 ) drives the shaft  26 , which is coupled to a sun wheel, or sun gear,  28  of the epicyclic gear arrangement  30 . Radially outwardly of the sun gear  28  and intermeshing therewith is a plurality of planet gears  32  that are coupled together by a planet carrier  34 . The planet carrier  34  constrains the planet gears  32  to process around the sun gear  28  in synchronicity whilst enabling each planet gear  32  to rotate about its own axis. The planet carrier  34  is coupled via linkages  36  to the fan  23  in order to drive its rotation about the engine axis  9 . Radially outwardly of the planet gears  32  and intermeshing therewith is an annulus or ring gear  38  that is coupled, via linkages  40 , to a stationary supporting structure  24 . 
     Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan  23 ) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft  26  with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan  23 ). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan  23  may be referred to as a first, or lowest pressure, compression stage. 
     It will be appreciated that the arrangement shown in  FIGS. 1 and 2  is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox  30  in the engine  10  and/or for connecting the gearbox  30  to the engine  10 . By way of further example, the connections (such as the linkages  36 ,  40  in the  FIG. 2  example) between the gearbox  30  and other parts of the engine  10  (such as the input shaft  26 , the output shaft and the fixed structure  24 ) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of  FIG. 2 . For example, where the gearbox  30  has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in  FIG. 2 . 
     Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations. Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor). The present disclosure also extends to gas turbine engines without a gearbox, for example an engine such as the engine  10  of  FIGS. 1 and 2  in which gearbox  30  is omitted and fan  23  is driven directly by shaft  26 . 
     Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in  FIG. 1  has a split flow nozzle  20 ,  22  meaning that the flow through the bypass duct  22  has its own nozzle that is separate to and radially outside the core engine nozzle  20 . However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct  22  and the flow through the core  11  are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. 
     Referring to both  FIGS. 1 and 2 , the core  11  of the engine  10  has a core cowling  25 , the outer surface of which forms the inner surface of the bypass duct  22 . The gas annulus  35  of the core  11  in the portion of the engine  10  depicted in  FIG. 2  is defined by outer and inner compressor casing walls  29 ,  31  between which blades of the low pressure compressor  14  are located. Compressed (pressurised) air within the gas annulus  35  is retained by the outer and inner compressor casing walls  29 ,  31  during operation of the engine  10 . The outer and inner compressor casing walls  29 ,  31  therefore act as containment walls for a volume of pressurised air between them. A space  27  defined by the outer compressor casing wall  29  and the core cowling  25  houses various other engine components. An output port in the outer compressor casing wall  29  is provided at a location  33  downstream of the low pressure compressor  14  for the purpose of bleeding off compressed air selectively, that is as required or necessary (under human or automatic control) to operate certain auxiliary systems of an aircraft comprising the engine  10 . 
       FIGS. 3 and 4  show conventional apparatus  50  comprised in the engine  10  for selective delivery of compressed air derived from the downstream end of the low pressure compressor  14  via the output port  36  in the compressor casing outer wall  29  at the location  33 . The output port  36  is an aperture in the compressor casing outer wall  29 . A stub pipe  42  is attached to the exterior of the compressor casing outer wall  29  at the output port  36  by means of bolts  37 . The stub pipe  42  is coupled to the input of a valve  44  by an attachment feature  45 . The output of valve  44  is coupled to an output duct  46  by an attachment feature  47 . When the valve  44  is open, compressed within the gas annulus  35  passes to the output duct  46 , providing compressed air input to one or more auxiliary aircraft systems (not shown). When the valve  44  is closed, air within the stub pipe  42  and within the valve  44  up to its sealing line  49  (i.e. the sealing position of a valve mechanism  43  contained within the valve body of the valve  44 ) is at the same pressure as air within the gas annulus  35 . The valve  44  may have only two states, i.e. fully open and fully closed, or alternatively it may allow the rate of throughput of air to be controlled when it is operated. The valve  44  is located radially outwardly of the output port  36  with respect to the axis  9 . 
     Under certain engine operating conditions, for example at aircraft take-off, the pressure and temperature of air in the gas annulus  35  at the position  33  may be very high. With the valve  44  in its closed configuration, the stub pipe  42  may rupture under such conditions, or the interface between the stub pipe  42  and the compressor casing outer wall  29  may fail, or the interface between the stub pipe  42  and the input of the valve  44  may fail. A combination of two or more of these failure modes may occur simultaneously. One or more such failures will result in the escape of hot, high-pressure air from the gas annulus  35  into the space  27 , damaging engine components located therein. Systems may be provided in the engine  10  to detect such failures and to cut off the engine&#39;s fuel supply in response; however such systems interfere with engine operation and add complexity to the engine  10 . 
       FIGS. 5 and 6  show apparatus  150  according to an example for selective delivery of pressurised air obtained from the gas annulus  135  of an engine having a compressor with an outer compressor casing wall  129  which partly defines the gas annulus  135 . A valve  144  comprising a valve body containing a valve mechanism  143  with a sealing line  149  is coupled directly to the exterior of the compressor casing outer wall  129  substantially within an output port  136  thereof which is formed by a portion of the compressor casing outer wall  129  which extends radially outwardly from the remainder of the compressor casing outer wall, with respect to the principal rotation axis  109  of the engine. The output port  136  thus extends outwardly with respect to a volume of pressurised air in the gas annulus  135  retained by outer and inner compressor casing walls of the engine. The output port  136  may for example be located at the downstream end of the compressor. The valve  144  is substantially contained within the output port  136 . An output duct arrangement  146  is coupled to the output of the valve  144 . A pair of bolts  139  attaches the duct arrangement  146  to the valve  144 , and the valve  144  to the compressor casing outer wall  129 . The sealing line  149  of the valve mechanism  143  of the valve  144  is located within the output port  136 . The input of the valve  144  is thus in direct fluid communication with the gas annulus  135 , obviating the need for a connecting duct coupling the output port  136  to the valve  144 , as is the case in the apparatus  50  of  FIGS. 3 and 4 . The possibility of rupturing of such a duct is therefore removed. In addition, a single interface between the valve  144  and the compressor casing outer wall  129  replaces two interfaces in the arrangement of  FIGS. 3 and 4 ; a first between a first end of connecting duct  42  and the valve  44  and a second between a second end of the duct  42  and the compressor casing outer wall  29 . 
       FIG. 7  shows a second example apparatus  250  for selective delivery of compressed air, in which a valve  244  is coupled directly to an output port  236  in the compressor casing outer wall  229  of a gas turbine engine. The output port  236  is formed by a portion of the wall  229  which extends radially inwardly with respect to the principal rotational axis of the engine (not shown), and forms a recessed step for accommodating the valve  244 . Bolts  139 A fix the valve  244  to the compressor casing outer wall  229 . An output duct  246  is attached to the exterior of the compressor casing outer wall  229  by bolts  139 B. The valve  244  is located substantially entirely within the output port  236 , and the sealing line  249  is therefore also located within the output port  236 . When the valve  244  is opened, compressed air from the gas annulus  235  passes into the output duct  246  via the valve  244 . The output port  236  thus extends inwardly with respect to a volume of pressurised air produced in the gas annulus  135 , the volume being partly bounded and retained by the compressor casing outer wall  229 . 
       FIG. 8  shows a third example apparatus  350  in which an output port  336  in a compressor casing outer wall  329  of a gas turbine engine takes the form of a simple aperture which is occupied by a valve  349 . An output duct  346  is coupled to the exterior of the compressor casing outer wall  329  to provide a seal between the valve  344  and the compressor casing outer wall  329  at the position of the output port  336 . 
       FIGS. 9 to 12  show portions of fourth to seventh example apparatus  450 ,  550 ,  650 ,  750  respectively for selective delivery of compressed air. Parts of the apparatus  450 ,  550 ,  650 ,  750  of  FIGS. 9 to 12  are labelled with reference numerals differing by  300 ,  400 ,  500  and  600  respectively from reference numerals labelling corresponding parts in  FIGS. 5 and 6 . 
     Wth reference to  FIG. 9 , an output port  436  in a compressor casing outer wall  429  is formed by a radially outwardly extending portion of the wall  429  and an additional, separate, hollow cylindrical component  429 A which is attached to the outwardly-extending portion of wall  429  by bolts  441  and to which a valve  444  is attached by means of bolts  443 . An output duct (not shown) may be attached to the valve  444  on a side thereof remote from the output port  436 . The hollow cylindrical component  429 A is an extension of the outwardly-extending portion of the wall  429  and may be permanently attached to that portion, for example by welding, so that it is integrated with the wall  429 . 
       FIG. 10  shows apparatus  550  similar to the apparatus  450  of  FIG. 5 , in which a radially outwardly-extending output port  536  is formed entirely by a radially-outwardly extending portion of a compressor casing outer wall  529 . 
     The apparatus  450  of  FIG. 9  may be more economical to manufacture than the apparatus  550  of  FIG. 10  for a given output port length because the compressor casing outer wall  429  may be cheaper and easier to forge due to the shorter radially outwardly extending portion of the wall  429  compared to that of the wall  529 . 
       FIGS. 11 and 12  show apparatus  650 ,  750  similar to the apparatus  450 ,  550  of  FIGS. 9 and 10  respectively, except that the output ports  636 ,  736  each increase in diameter and cross-sectional area in the radial direction away from gas annuli  635 ,  735  (i.e. as a function of distance from parts of the compressor casing outer walls  629 ,  729  which partly define gas annuli  635 ,  735 ). Specifically, the output ports  636 ,  736  each exhibit a stepwise increase in diameter and cross-sectional area at a particular distance from the exterior of the main (non-port) portions of the compressor casing outer walls  629 ,  729 . This allows the effective cross-sectional area for gas passing through the output ports  636 ,  736  and valves  636 ,  736  to remain constant or substantially constant as a function of radial position. 
     The valves  144 ,  244 ,  344 ,  444 ,  544 ,  644 ,  744  in  FIGS. 5 to 12  are all modular; each comprises valve body housing a valve mechanism. In the event of failure of a valve the entire valve is replaced, providing for simple maintenance of apparatus comprising the valve. 
     It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.