Abstract:
A bypass duct has a support unit comprising a pair of aerofoils arranged as an “A” frame. A bleed duct assembly is provided on the radially inner wall of the bypass duct annulus and the aerofoils project from the surface and extend across the annulus between the inner wall and an outer wall of the annulus. The aerofoils lean at an acute angle to the surface with the first flank facing toward the inner wall and adjoining a bleed duct opening. The bleed duct having a bleed duct passage and a submerged scoop.

Description:
TECHNICAL FIELD OF INVENTION 
       [0001]    The present invention relates to bleed ducts in a gas turbine engine and in particular bleed ducts adjoining an aerofoil extending across a bypass duct of the engine 
       BACKGROUND OF INVENTION 
       [0002]    Modern gas turbines are provided with an engine core comprising a compressor, combustor and turbine section and a surrounding annular bypass duct through which an air flow is guided by a fan. The bypass duct is limited by a radially inner wall and by a radially outer wall. Between the inner wall and the outer wall of the bypass duct, a support unit is provided that includes strut-like support elements connected at one end to the inner wall and at the other end to the outer wall. 
         [0003]      FIG. 1  depicts a ducted fan gas turbine engine generally indicated at  10  which comprises, in axial flow series, an air intake  1 , a propulsive fan  12 , an intermediate pressure compressor  7 , a high pressure compressor  9 , combustion equipment  115 , a high pressure turbine  116 , an intermediate pressure turbine  117 , a low pressure turbine  118  and an exhaust nozzle  119 . 
         [0004]    Air entering the air intake  8  is accelerated by the fan  12  to produce two air flows, a first air flow  11  into the intermediate pressure compressor  7  and a second air flow  10  that passes over the outer surface of the engine casing  12  and through a bypass duct  2  which provides propulsive thrust. The intermediate pressure compressor  7  compresses the air flow directed into it before delivering the air to the high pressure compressor  9  where further compression takes place. 
         [0005]    Compressed air exhausted from the high pressure compressor  9  is directed into the combustion equipment  115 , where it is mixed with fuel that is injected from a fuel injector and the mixture combusted. The resultant hot combustion products expand through and thereby drive the high  116 , intermediate  117  and low pressure  118  turbines before being exhausted through the nozzle  119  to provide additional propulsive thrust. The high, intermediate and low pressure turbines respectively drive the high and intermediate pressure compressors and the fan by suitable interconnecting shafts. 
         [0006]    In the bypass duct  2 , several fan outlet guide vanes  13  are arranged downstream of the fan  9  which reduce or remove a twist in the flow of the bypass flow  10 . In addition, supports  14   a  or  14   b  are provided downstream of the fan outlet guide vanes  13 , bracing the inner wall  3  and the outer wall  4  against one another. In addition to the supports  14 , the engine can have further support structures called bifurcations  15  through which lines are routed for supplying the jet engine device  1  or an airframe of an aircraft provided with the jet engine device  1 . The position of the support  14   a  is the position typically used within large or medium civil gas turbines. The position of the support  14   b  is that typically used in smaller, business jet type applications. The general structure and form of the supports is the same for each and will be discussed in common as reference  14  in the rest of the specification. 
         [0007]    Depending on the specific application, it is also possible for the bifurcations  15  to be arranged in the same cross-sectional plane as the supports  14 . 
         [0008]    The support unit  14  includes strut-like aerofoil support elements  17  to  20  shown in more detail in  FIG. 2  and connected at one end to the inner wall  3  and at the other end to the outer wall  4 . The aerofoils  17  to  20  of each support element each describe an acute angle  17 E,  18 E,  19 E and  20 E between the circumferentially outward facing flanks  17 C,  18 C,  19 C and  20 C and the radially outer wall of the bypass duct  4 . The aerofoils  17  to  20  of each support element also each describe an acute angle  17 F,  18 F,  19 F and  20 F between the circumferentially inward facing flanks  17 D,  18 D,  19 D and  20 D and the radially inner wall of the bypass duct  3 . 
         [0009]    Each support unit comprises two aerofoils  17  and  18  or  19  and  20 , respectively that are connected in a manner forming an A-arrangement to the inner wall  3  and to the outer wall  4  in the manner shown in more detail in  FIG. 2  and form so-called A-frames of the support unit  14 . The “A” tapering as the support unit progresses radially outwards. The facing flanks of each aerofoil pair describing an acute angle with the radially inner wall of the bypass duct. 
         [0010]    The aerofoils  17  and  18  or  19  and  20 , respectively, representing A-frames level in the bypass duct  2  with a duct height H in order to transmit engine loads acting in the area of the engine core  12  outwards in the direction of the outer wall  4 . The aerofoils  17  and  18  or  19  and  20 , respectively, assigned to one another in pairs form the support units and are, depending on the specific application, arranged relative to one another at a defined acute angle  17 F,  18 F or  19 F,  20 F, respectively, and at a distance D defined in the circumferential direction. 
         [0011]    The acute angle of the aerofoil flanks to the radially inner and outer walls causes an increase in the velocity of the air which subsequently interacts with the main flow boundary layer and causes wakes to form which add to the pressure loss through the bypass duct and can take energy from the bulk flow. The lost energy reduces the overall efficiency of the gas turbine engine and reduces the engine performance. 
         [0012]    In many conventional engines an offtake is provided in the by-pass duct to supply cool air for proper functioning of the engine and its units. The offtakes are separated from the aerofoils and also generate wakes as shown in  FIG. 3  and therefore also generate further pressure loss in the by-pass duct. 
         [0013]      FIG. 4  depicts the static pressure of the bypass flow  10  for the arrangement of  FIG. 3 . The bulk pressure has a region  30  where the pressure is relatively constant. At the leading edge of the support unit there are regions of lower static pressure  32  and similar regions of lower static pressure  34 ,  36  can be seen at boundary layers of the radially inner wall  3  and the radially outer wall  4  respectively. 
         [0014]    The regions of lowest static pressure  38 ,  40  are found where the aerofoils of the A frame form an acute angle with the radially inner  3  and the radially outer  4  wall of the bypass duct. These regions can be significant and can cause significant wakes to form that reduce the efficiency of the gas turbine engine. The wakes are formed in part by the acute angle that the aerofoil forms to the duct wall and which causes an increase in the velocity of the air that goes on to interact with the main flow boundary layer. The wakes generate an area of increased loss in the by-pass duct flow and a reduction in engine performance. 
         [0015]    It is an object of the present invention to seek to provide an arrangement having an improved efficiency. 
       STATEMENTS OF INVENTION 
       [0016]    According to the present invention there is provided a bleed duct assembly for a gas turbine bypass duct comprising a circumferentially extending surface defining an inner wall of an annulus onto which a bleed duct opens at a bleed duct opening and an aerofoil projecting from the surface and extending across the annulus between the inner wall and an outer wall of the annulus, the aerofoil having a leading edge, a trailing edge and first and second flanks connecting the leading edge and the trailing edge, wherein the aerofoil is leant at an acute angle to the surface with the first flank facing toward the inner wall and adjoining the bleed duct opening. 
         [0017]    Advantageously, the invention allows generated wakes from the bleed duct and the aerofoils to be combined such that the resulting wake of the combined features has smaller pressure drop than the sum of the wakes of the separate features. 
         [0018]    Advantageously, the lean of the aerofoil can assist in directing the flow radially inwards towards the duct opening. This turning of the flow increases the flow through the opening and allows a smaller opening to be used to generate a desired duct flow than the opening required to generate the same flow if the duct opening is positioned away from the aerofoil flank. 
         [0019]    Preferably the aerofoil provides an axially extending edge to the bleed duct. 
         [0020]    The bleed duct opening may have a leading edge at a first axial location and a trailing edge at a second axial location axially rearward of the first axial location 
         [0021]    The first axial location is preferably at or before the axial location of the leading edge of the aerofoil. 
         [0022]    The second axial location is preferably rearward of the axial location of the leading edge of the aerofoil and at or aft of the axial location of the trailing edge of the aerofoil. 
         [0023]    The second axial location may be rearward of the axial location of the trailing edge of the aerofoil. 
         [0024]    The leading edge of the bleed duct opening and the trailing edge of the bleed duct opening may extend substantially circumferentially, wherein the circumferential length of the leading edge is less than the circumferential length of the trailing edge. 
         [0025]    The leading edge of the bleed duct opening may be angled such that it is orthogonal to the boundary layer flow direction. 
         [0026]    The bleed duct may have a bleed duct passage and a scoop recessed radially inside the inner wall. 
         [0027]    The recessed scoop preferably has a side wall, the first flank projecting radially inside the surface and providing at least part of the scoop side wall. 
         [0028]    According to a second aspect of the invention there is provided a support unit for a gas turbine engine, the support unit having a pair of aerofoils with a first aerofoil having the aerofoil having a leading edge, a trailing edge and first and second flanks connecting the leading edge and the trailing edge, the aerofoil leaning at an acute angle to the surface with the first flank facing toward the inner wall and a second aerofoil having a leading edge, a trailing edge and first and second flanks connecting the leading edge and the trailing edge, the aerofoil leaning at an acute angle to the surface with the first flank facing toward the inner wall, wherein the first flank of the first aerofoil and the first flank of the second aerofoil face towards each other; characterised in that the first aerofoil is the aerofoil in a bleed duct assembly according to any of the preceding 12 paragraphs. 
         [0029]    The second aerofoil may also be the aerofoil in a bleed duct assembly according to any of the preceding claims. There may be a duct opening for each of the aerofoils in the support unit 
         [0030]    According to a third aspect of the invention there is provided a gas turbine engine having a support unit according to the preceding two paragraphs. 
     
    
     
       DESCRIPTION OF DRAWINGS 
         [0031]      FIG. 1  depicts a conventional ducted fan gas turbine engine. 
           [0032]      FIG. 2  depicts another view of a conventional ducted fan gas turbine engine. 
           [0033]      FIG. 3  depicts wakes generated when offtakes are separated from the aerofoils. 
           [0034]      FIG. 4  depicts the static pressure of the bypass flow for the arrangement of  FIG. 3 . 
           [0035]      FIG. 5  depicts a perspective view of the “A-frames” arrangement in accordance with the invention; 
           [0036]      FIG. 6  depicts an enlarged view of the cross-section through aerofoil  20  of  FIG. 3 ; 
           [0037]      FIG. 7  shows a schematic perspective view of a bleed duct arranged relative to an aerofoil assembly; 
           [0038]      FIG. 8  depicts a rearward looking view of the bleed duct/aerofoil of  FIG. 6 ; 
           [0039]      FIG. 9  depicts a forward looking view of the bleed duct/aerofoil of  FIG. 6 ; 
           [0040]      FIG. 10  shows a schematic forward looking perspective view of the bleed duct/aerofoil of  FIG. 7 ; 
           [0041]      FIG. 11  shows a simplified rear view of the arrangement of  FIG. 10  that depicts the location of an actuator to the bleed duct. 
           [0042]      FIG. 12  shows the combined wakes of the aerofoils and associated bleed duct. 
           [0043]      FIG. 13  shows the pressure distribution in the bypass duct along adjoining the aerofoil having a bleed duct in accordance with the invention. 
       
    
    
     DETAILED DESCRIPTION OF INVENTION 
       [0044]    A perspective view of the “A frames” is shown in  FIG. 5 . The aerofoils  17 ,  18  and  19 ,  20  making up the frames extend through the inner wall  3  of the bypass duct and extend radially across the duct. A bifurcation  15  is also shown. Each aerofoil  17  to  20  is designed with aerodynamically shaped cross-sectional profiles which when radially stacked one above the other determine the shape of the A-frame or support unit  14 . The aerofoils  17  to  20  of the support unit  14  here have no curvature relative to the engine axis  6  and are designed relative to a central longitudinal plane  21  with a thickness distribution forming a desired profile e.g. an elliptical and hence aerodynamically optimized cross-sectional profile in a cross-sectional plane  22  perpendicular to the central longitudinal plane  21 . An exemplary cross-section is shown in  FIG. 6  which is an enlarged view of the end of the aerofoil  20  in  FIG. 3 . 
         [0045]    To address the problems with the regions of low pressure and wake generation a bleed duct  42  is located adjacent to the aerofoil flank  18 D which forms the acute angle with the radially inner wall  3  of the bypass duct. As depicted in  FIG. 7 , which is a schematic perspective view, and  FIG. 8 , which is a view looking rearwardly along the a line parallel to the axis of the engine, and  FIG. 9 , which is a view looking forward along a line parallel to the axis of the engine the duct is preferably of the submerged inlet type having an opening  42  that is flush with the surface  3  and which feeds a duct passage  44  via a scoop  46 . 
         [0046]    The opening and scoop is of the NACA type which allows air flow into the duct passage  44  with a minimal disturbance to the main flow  10 . The scoop consists of a shallow ramp with walls that are recessed below the radially inner wall  3 . The opening and scoop both flare from a relatively narrow upstream edge to a wider downstream edge. The side edges of the opening and scoop have a curved profile to minimise detrimental vortices being shed therefrom. 
         [0047]    The combination of the shallow ramp angle and the side walls create counter rotating vortices  47  which deflect the boundary layer away from the inlet to draw in the faster moving are whilst avoiding the drag and flow separation that can occur with protruding scoop designs. 
         [0048]    The scoop leading edge  48  is preferably located axially upstream of the leading edge  18   a  of the aerofoil  18  whilst the trailing edge of the scoop is positioned axially rearward of the trailing edge  18   b  of the aerofoil. Circumferentially in the engine it is desirable for the scoop leading edge  48  to be positioned in line with the leading edge of the aerofoil, or slightly circumferentially offset from the leading edge and curved such that the flank of the aerofoil forms one of the side edges of the scoop. 
         [0049]    The streamline flow is depicted as lines  49  and is directed around the aerofoil. The aerofoil angle  17 F,  18 F,  19 F,  20 F can assist in directing the flow radially inwards towards the opening and the scoop. This turning of the flow increases the flow into the scoop and allows a smaller opening to be used to generate a desired duct flow than the opening required to generate the same flow if the duct is positioned away from the aerofoil flank. At the rearmost edge of the scoop a raised lip  50  is blended with the trailing edge of the aerofoil and the surface  3  to provide an aerodynamic surface which helps to scoop the turned air from the “A frame” aerofoil  18  into the duct  46 . The raised lip slopes radially inwardly from the trailing edge of the respective aerofoil towards surface  3  as it extends circumferentially across the scoop  42 . The lip may also extend axially either from a forward location at the trailing edge of the aerofoil to the surface or from the surface to a rearward location at the trailing edge.  FIG. 10  is a perspective view looking axially forward towards the front of the engine of the scoop arrangement of  FIGS. 7 to 9 . 
         [0050]      FIG. 10  also shows the position of an electric valve actuator  52  that can be used to control the amount of air flowing through the bleed duct preferably by sliding the raised lip  50  fore and aft. Advantageously, the actuator is located within in a cold environment exemplified by arrow  56  in the engine and separated from a hot environment exemplified by arrow  58  by a fire-shield  54 . The actuator is preferably located circumferentially adjacent to the bleed duct passage  46  as shown in  FIG. 11 . 
         [0051]    One of the advantages of locating the bleed duct adjacent to the aerofoil is shown in  FIG. 12  as the wakes generated by the aerofoil and wakes generated by the bleed duct are combined into a single wake group  58  such that their combined value is less than the sum of their individual values. 
         [0052]    As shown in  FIG. 13 , which is equivalent to  FIG. 3  , the static pressure of the bypass flow  10  along cross-section II-II of  FIG. 1  is shown with a scoop  42  located adjacent the A frame aerofoil. The bulk pressure has a region  30  where the pressure is relatively constant. As can be noted, the region of lowest static pressure which were found where the aerofoils of the A frame form an acute angle with the radially inner wall  3  of the bypass duct has been significantly removed by locating the scoop opening adjacent the aerofoil flank. Advantageously, wake generation is reduced giving rise to an overall improvement in the efficiency of the gas turbine engine. 
         [0053]    It will be appreciated that modifications may be made without departing from the invention described herein. For example the axial location of the leading edge of the scoop may be moved in-line or aft of the leading edge of the aerofoil. 
         [0054]    Scoops may be provided for one or more of the aerofoils in the support units adjacent to the flank forming the acute angle with the surface having the duct opening. 
         [0055]    The air within the duct passage may be utilised for a cooling flow within the combustor or turbine section of the engine or to cool auxiliary components within the inner core fairing.