Patent Publication Number: US-11391164-B2

Title: Compressor aerofoil

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the US National Stage of International Application No. PCT/EP2018/065820 filed 14 Jun. 2018, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP17177900 filed 26 Jun. 2017. All of the applications are incorporated by reference herein in their entirety. 
     FIELD OF INVENTION 
     The present invention relates to a compressor aerofoil. 
     In particular it relates to a compressor aerofoil rotor blade and/or compressor aerofoil stator vane for a turbine engine, and/or a compressor rotor assembly. 
     BACKGROUND 
     A compressor of a gas turbine engine comprises rotor components, including rotor blades and a rotor drum, and stator components, including stator vanes and a stator casing. The compressor is arranged about a rotational axis with a number of alternating rotor blade and stator vane stages, and each stage comprises an aerofoil. 
     The efficiency of the compressor is influenced by the running clearances or radial tip gap between its rotor and stator components. The radial gap or clearance between the rotor blades and stator casing and between the stator vanes and the rotor drum is set to be as small as possible to minimise over tip leakage of working gases, but sufficiently large to avoid significant rubbing that can damage components. The pressure difference between a pressure side and a suction side of the aerofoil causes the working gas to leak through the tip gap. This flow of working gas or over-tip leakage generates aerodynamic losses due to its viscous interaction within the tip gap and with the mainstream working gas flow particularly on exit from the tip gap. This viscous interaction causes loss of efficiency of the compressor stage and subsequently reduces the efficiency of the gas turbine engine. 
     Two main components to the over tip leakage flow have been identified, which is illustrated in  FIG. 1 , which shows an end on view of a tip  1  of an aerofoil  2  in situ in a compressor, thus showing a tip gap region. A first leakage component “A” originates near a leading edge  3  of the aerofoil at the tip  1  and which forms a tip leakage vortex  4 , and a second component  5  that is created by leakage flow passing over the tip  1  from the pressure side  6  to the suction side  7 . This second component  5  exits the tip gap and feeds into the tip leakage vortex  4  thereby creating still further aerodynamic losses. 
     Hence an aerofoil design which can reduce either or both tip leakage components is highly desirable. 
     SUMMARY 
     According to the present disclosure there is provided apparatus as set forth in the appended claims. Other features of the invention will be apparent from the dependent claims, and the description which follows. 
     Accordingly there may be provided a compressor aerofoil ( 70 ) for a turbine engine, the compressor aerofoil ( 70 ) comprising: a root portion ( 72 ) spaced apart from a tip portion ( 100 ) by a main body portion ( 102 ). The main body portion ( 102 ) may be defined by: a suction surface wall ( 88 ) having a suction surface ( 89 ), a pressure surface wall ( 90 ) having a pressure surface ( 91 ), whereby the suction surface wall ( 88 ) and the pressure surface wall ( 90 ) meet at a leading edge ( 76 ) and a trailing edge ( 78 ). The tip portion ( 100 ) may comprise: a shoulder ( 104 ) provided on the pressure surface wall ( 90 ) between the leading edge ( 76 ) and the trailing edge ( 78 ); a tip wall ( 106 ) which extends from the aerofoil leading edge ( 76 ) to the aerofoil trailing edge ( 78 ); a transition region ( 108 ) of the pressure surface wall ( 90 ) which tapers from the shoulder ( 104 ) in a direction towards the tip wall ( 106 ). The tip wall ( 106 ) may comprise: a squealer ( 110 ) defined by a first tip wall region ( 112 ) which extends from the trailing edge ( 78 ) to a winglet ( 114 ) defined by a second tip wall region ( 116 ) which increases in width relative to the first tip wall region ( 112 ) to a tip wall widest point (A-A), and then reduces in width towards the leading edge ( 76 ). 
     The first tip wall region ( 112 ) which defines the squealer ( 110 ) may have a substantially constant width w1B along its extent. The first tip wall region ( 112 ) which defines the squealer ( 110 ) may have a substantially constant width w1B along at least part of its extent. 
     The distance between pressure surface ( 91 ) and the suction surface ( 89 ) of the main body ( 102 ) along the extent of the squealer is wbB, wherein the squealer width w1B may have a value of at least 0.1 wbB but no more than 0.2 wbB. 
     A chord line from the leading edge ( 76 ) to the trailing edge ( 78 ) has a length L; and the winglet ( 114 ) extends from the leading edge ( 76 ) towards the trailing edge ( 78 ) by a distance L1, where L1 may have a value of at least 0.25 L but no more than 0.65 L. 
     The widest point (A-A) of the winglet ( 114 ) is at a distance of L2 from the leading edge ( 76 ), where L2 may have a value of at least 0.4 L1 but no more than 0.6 L1. 
     Along the length of the winglet ( 114 ), the winglet ( 114 ) may be narrower than a distance wbA between the pressure surface ( 91 ) and the suction surface ( 89 ) in the corresponding region of winglet ( 114 ). 
     Along the length of the winglet ( 114 ), the winglet ( 114 ) may be recessed beneath the pressure surface ( 91 ). 
     The widest point (A-A) of the winglet ( 114 ) may have a width w3A of at least 0.8 wbA but no more than 0.95 wbA. 
     The tip wall ( 106 ) may define a tip surface ( 118 ) which extends from the aerofoil leading edge ( 76 ) to the aerofoil trailing edge ( 78 ). At the widest point (A-A) of the winglet ( 114 ): the transition region ( 108 ) of the pressure surface wall ( 90 ) may extend from the shoulder ( 104 ) in a direction towards the suction surface ( 89 ), and at an inflexion point ( 120 ) the transition region ( 108 ) may curve to extend in a direction away from the suction surface ( 89 ) toward the tip surface ( 118 ). 
     The tip portion ( 100 ) may further comprise an inflexion line ( 122 ) defined by a change in curvature on the pressure surface ( 91 ); the inflexion point ( 120 ) being provided on the inflexion line ( 122 ). The inflexion line ( 122 ) may extend between the leading edge ( 76 ) and the trailing edge ( 78 ). 
     The inflexion line ( 122 ) is provided a distance h2A, h2B from the tip surface ( 118 ); and the shoulder ( 104 ) is provided a distance h1A, h1B from the tip surface ( 118 ); where distance h1A, h1B may have a value of at least 1.5 h2A but no more than 2.7 h2A. 
     The inflexion line ( 122 ) at the widest point of the winglet ( 114 ) is provided a distance w2A from the suction surface ( 89 ); wherein w2A may have a value of at least 0.8 w3A but no more than 0.95 w3A. 
     The pressure surface ( 91 ) and the suction surface ( 89 ) are spaced apart by a distance wbA, wbB. The distance wbA, wbB may decrease in value between the main body widest point (A-A) and the leading edge ( 76 ). The distance wbA, wbB may decrease in value between the main body widest point (A-A) and the trailing edge ( 78 ). 
     There may also be provided a compressor rotor assembly for a turbine engine, the compressor rotor assembly comprising a casing and a compressor aerofoil according to the present disclosure, wherein the casing and the compressor aerofoil ( 70 ) define a tip gap hg defined between the tip surface ( 118 ) and the casing ( 50 ). 
     The distance h2A, h2B from the inflexion line ( 122 ) to the tip surface ( 118 ) may have a value of at least 1.5 hg but no more than 3.5 hg. 
     Hence there is provided an aerofoil for a compressor which is reduced in thickness towards its tip to form a squealer on the suction (i.e. convex) side of the aerofoil. In addition a winglet type extension is provided on the pressure (i.e. concave) side near the leading edge. Together, these features reduce the tip leakage mass flow thus diminishing the strength of the interaction between the leakage flow and the main stream flow which in turn reduces loss in efficiency relative to examples of the related art. 
     Hence the compressor aerofoil of the present disclosure provides a means of controlling losses by reducing the tip leakage flow. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Examples of the present disclosure will now be described with reference to the accompanying drawings, in which: 
         FIG. 1  shows an example aerofoil tip, as discussed in the background section; 
         FIG. 2  shows part of a turbine engine in a sectional view and in which an aerofoil of the present disclosure may be provided; 
         FIG. 3  shows an enlarged view of part of a compressor of the turbine engine of  FIG. 2 ; 
         FIG. 4  shows part of a main body and a tip region of an aerofoil according to the present disclosure; 
         FIGS. 5 a , 5 b    show sectional views of the aerofoil as indicated at A-A and B-B in  FIG. 4 ; 
         FIG. 6  shows an end on view of a part of the tip region of the aerofoil shown in  FIG. 4 ; and 
         FIG. 7  is a table of relative dimensions of the features shown in  FIGS. 5 a , 5 b   ,  6 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 2  shows an example of a gas turbine engine  10  in a sectional view which may comprise an aerofoil and compressor rotor assembly of the present disclosure. 
     The gas turbine engine  10  comprises, in flow series, an inlet  12 , a compressor section  14 , a combustor section  16  and a turbine section  18  which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis  20 . The gas turbine engine  10  further comprises a shaft  22  which is rotatable about the rotational axis  20  and which extends longitudinally through the gas turbine engine  10 . The shaft  22  drivingly connects the turbine section  18  to the compressor section  14 . 
     In operation of the gas turbine engine  10 , air  24 , which is taken in through the air inlet  12  is compressed by the compressor section  14  and delivered to the combustion section or burner section  16 . The burner section  16  comprises a burner plenum  26 , one or more combustion chambers  28  and at least one burner  30  fixed to each combustion chamber  28 . 
     The combustion chambers  28  and the burners  30  are located inside the burner plenum  26 . The compressed air passing through the compressor  14  enters a diffuser  32  and is discharged from the diffuser  32  into the burner plenum  26  from where a portion of the air enters the burner  30  and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the resulting combustion gas  34  or working gas from the combustion is channelled through the combustion chamber  28  to the turbine section  18 . 
     The turbine section  18  comprises a number of blade carrying discs  36  attached to the shaft  22 . In addition, guiding vanes  40 , which are fixed to a stator  42  of the gas turbine engine  10 , are disposed between the stages of annular arrays of turbine blades  38 . Between the exit of the combustion chamber  28  and the leading turbine blades  38 , inlet guiding vanes  44  are provided and turn the flow of working gas onto the turbine blades  38 . 
     The combustion gas from the combustion chamber  28  enters the turbine section  18  and drives the turbine blades  38  which in turn rotate the shaft  22 . The guiding vanes  40 ,  44  serve to optimise the angle of the combustion or working gas on the turbine blades  38 . 
     Compressor aerofoils (that is to say, compressor rotor blades and compressor stator vanes) have a smaller aspect ratio than turbine aerofoils (that is to say, turbine rotor blades and turbine stator vanes), where aspect ratio is defined as the ratio of the span (i.e. width) of the aerofoil to the mean chord (i.e. straight line distance from the leading edge to the trailing edge) of the aerofoil. Turbine aerofoils have a relatively large aspect ratio because they are necessary broader (i.e. wider) to accommodate cooling passages and cavities, whereas compressor aerofoils, which do not require cooling, are relatively narrow. 
     Compressor aerofoils also differ from turbine aerofoils by function. For example compressor rotor blades are configured to do work on the air that passes over them, whereas turbine rotor blades have work done on them by exhaust gas which passes over them. Hence compressor aerofoils differ from turbine aerofoils by geometry, function and the working fluid which they are exposed to. Consequently aerodynamic and/or fluid dynamic features and considerations of compressor aerofoils and turbine aerofoils tend to be different as they must be configured for their different applications and locations in the device in which they are provided. 
     The turbine section  18  drives the compressor section  14 . The compressor section  14  comprises an axial series of vane stages  46  and rotor blade stages  48 . The rotor blade stages  48  comprise a rotor disc supporting an annular array of blades. The compressor section  14  also comprises a casing  50  that surrounds the rotor stages and supports the vane stages  46 . The guide vane stages include an annular array of radially extending vanes that are mounted to the casing  50 . The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions. 
     The casing  50  defines a radially outer surface  52  of the passage  56  of the compressor  14 . A radially inner surface  54  of the passage  56  is at least partly defined by a rotor drum  53  of the rotor which is partly defined by the annular array of blades  48  and will be described in more detail below. 
     The aerofoil of the present disclosure is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the aerofoil of the present disclosure is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications. The term rotor or rotor assembly is intended to include rotating (i.e. rotatable) components, including rotor blades and a rotor drum. The term stator or stator assembly is intended to include stationary or non-rotating components, including stator vanes and a stator casing. Conversely the term rotor is intended to relate a rotating component, to a stationary component such as a rotating blade and stationary casing or a rotating casing and a stationary blade or vane. The rotating component can be radially inward or radially outward of the stationary component. The term aerofoil is intended to mean the aerofoil portion of a rotating blade or stationary vane. 
     The terms axial, radial and circumferential are made with reference to the rotational axis  20  of the engine. 
     Referring to  FIG. 3 , the compressor  14  of the turbine engine  10  includes alternating rows of stator guide vanes  46  and rotatable rotor blades  48  which each extend in a generally radial direction into or across the passage  56 . 
     The rotor blade stages  49  comprise rotor discs  68  supporting an annular array of blades. The rotor blades  48  are mounted between adjacent discs  68 , but each annular array of rotor blades  48  could otherwise be mounted on a single disc  68 . In each case the blades  48  comprise a mounting foot or root portion  72 , a platform  74  mounted on the foot portion  72  and an aerofoil  70  having a leading edge  76 , a trailing edge  78  and a blade tip  80 . The aerofoil  70  is mounted on the platform  74  and extends radially outwardly therefrom towards the surface  52  of the casing  50  to define a blade tip gap, hg (which may also be termed a blade clearance  82 ). 
     The radially inner surface  54  of the passage  56  is at least partly defined by the platforms  74  of the blades  48  and compressor discs  68 . In the alternative arrangement mentioned above, where the compressor blades  48  are mounted into a single disc the axial space between adjacent discs may be bridged by a ring  84 , which may be annular or circumferentially segmented. The rings  84  are clamped between axially adjacent blade rows  48  and are facing the tip  80  of the guide vanes  46 . In addition as a further alternative arrangement a separate segment or ring can be attached outside the compressor disc shown here as engaging a radially inward surface of the platforms. 
       FIG. 3  shows two different types of guide vanes, variable geometry guide vanes  46 V and fixed geometry guide vanes  46 F. The variable geometry guide vanes  46 V are mounted to the casing  50  or stator via conventional rotatable mountings  60 . The guide vanes comprise an aerofoil  62 , a leading edge  64 , a trailing edge  66  and a tip  80 . The rotatable mounting  60  is well known in the art as is the operation of the variable stator vanes and therefore no further description is required. The guide vanes  46  extend radially inwardly from the casing  50  towards the radially inner surface  54  of the passage  56  to define a vane tip gap or vane clearance  83  therebetween. 
     Collectively, the blade tip gap or blade clearance  82  and the vane tip gap or vane clearance  83  are referred to herein as the ‘tip gap hg’. The term ‘tip gap’ is used herein to refer to a distance, usually a radial distance, between the tip&#39;s surface of the aerofoil portion and the rotor drum surface or stator casing surface. 
     Although the aerofoil of the present disclosure is described with reference to the compressor blade and its tip, the aerofoil may also be provided as a compressor stator vane, for example akin to vanes  46 V and  46 F. 
     The present disclosure may relate to an un-shrouded compressor aerofoil and in particular may relate to a configuration of a tip of the compressor aerofoil to minimise aerodynamic losses. 
     The compressor aerofoil  70  comprises a suction surface wall  88  and a pressure surface wall  90  which meet at the leading edge  76  and the trailing edge  78 . The suction surface wall  88  has a suction surface  89  and the pressure surface wall  90  has a pressure surface  91 . 
     As shown in  FIG. 3 , the compressor aerofoil  70  comprises a root portion  72  spaced apart from a tip portion  100  by a main body portion  102 . 
       FIG. 4  shows an enlarged view of part of a compressor aerofoil  70  according to the present disclosure.  FIGS. 5 a , 5 b    show sectional views of the aerofoil at points A-A and B-B as indicated in  FIG. 4 .  FIG. 6  shows an end on view of a part of the tip region of the aerofoil  70 , and  FIG. 7  summarises the relationship between various dimensions as indicated in  FIGS. 5 a , 5 b   ,  6 . 
     The main body portion  102  is defined by the convex suction surface wall  88  having the suction surface  89  and the concave pressure surface wall  90  having the pressure surface  91 . The suction surface wall  88  and the pressure surface wall  90  meet at the leading edge  76  and the trailing edge  78 . 
     As shown in  FIGS. 5 a , 5 b   , the pressure surface  91  and the suction surface  89  are spaced apart by a distance wb, identified as wbA, wbB at sections A-A and B-B respectively. The distance between the pressure surface  91  and the suction surface  89  (i.e. value wb, wbA, wbB) decreases in value between the main body widest point and the leading edge  76 . The distance between the pressure surface  91  and the suction surface  89  (i.e. the value wb, wbA, wbB) also decreases in value between the main body widest point and the trailing edge  78 . 
     The suction surface wall  88  and pressure surface wall  90  each extend from the root portion  72  to the tip portion  100 . 
     The tip portion  100  comprises a shoulder  104  provided on the pressure surface wall  90  between the leading edge  76  and the trailing edge  78 . The shoulder  104  extends at least part of the way between the leading edge  76  and the trailing edge  78 . The shoulder  104  may extend substantially the whole way between the leading edge  76  and the trailing edge  78 . 
     The tip portion  100  further comprises a tip wall  106  which extends from the aerofoil leading edge  76  to the aerofoil trailing edge  78 . The tip portion  100  also comprises a transition region  108  of the pressure surface wall  90  which tapers from the shoulder  104  in a direction towards the tip wall  106  such that the compressor aerofoil  70  is narrower at the tip wall  106  than between the pressure surface  91  and the suction surface  89  along the length of the shoulder  104 . 
     The shoulder  104  and the transition region  108  are each defined in the cross-sectional view of  FIGS. 5 a , 5 b    and each extends along at least a part of the tip portion  100  between the leading edge and the trailing edge. 
     On the suction surface wall  88 , the suction surface  89  of the tip portion  100  extends without interruption to the tip wall  106 . That is to say, the profile of the suction surface wall  89  continues into and through the tip portion  100  to the tip wall  106 . Put another way, in the tip portion  100 , the suction surface  89  extends in the same direction (i.e. with the same curvature) towards the tip wall  106  as it does in the main body portion  102 . That is to say, in the tip portion  100 , the suction surface  89  extends from the main body portion  102  without transition and/or change of direction towards the tip wall  106 . 
     The tip wall  106  comprises a squealer  110  defined by a first tip wall region  112  which extends from the trailing edge  78  to a winglet  114  defined by a second tip wall region  116  which increases in width relative to the first tip wall region  112  to a tip wall widest point (for example at A-A), and then reduces in width towards the leading edge  76 . 
     In one example, the first tip wall region  112  which defines the squealer  110  has a substantially constant width w1B along its extent. 
     In a further example, the first tip wall region  112  which defines the squealer  110  has a width w1B which varies along its extent, tapering towards the trailing edge  78 . 
     In another example, the squealer width w1B may have a value of at least about 0.1, but no more than about 0.2, of the distance wbB between pressure surface  91  and the suction surface  89  of the main body  102  along the extent of the squealer  110 . The value wbB varies along the length of the tip portion  100 , and hence the value of w1B may vary along the length of the tip portion  100 . 
     Put another way, where the distance between pressure surface  91  and the suction surface  89  of the main body  102  along the extent of the squealer is wbB, the squealer width w1B may have a value of at least about 0.1 wbB but no more than about 0.2 wbB. 
     As indicated in  FIGS. 4, 6 , the winglet  114  may extend from the leading edge  76  towards the trailing edge  78  by a chord distance L1, where L1 may have a value of at least about 0.25, but no more than about 0.65, of the chord length L (i.e. chord line) from the leading edge  76  to the trailing edge  78 . 
     For the avoidance of doubt, the term “chord” refers to an imaginary straight line which joins the leading edge  76  and trailing edge  78  of the aerofoil  70 . Hence the chord length L is the distance between the trailing edge  78  and the point on the leading edge  76  where the chord intersects the leading edge. 
     Hence chord distance L1 above (and L2 below) refer to a sub-section of the chord line L. 
     Put another way, where a chord line from the leading edge  76  to the trailing edge  78  has a length L, the winglet  114  extends from the leading edge  76  towards the trailing edge  78  by a distance L1, where L1 may have a value of at least about 0.25 L but no more than about 0.65 L. 
     The widest point (for example at section A-A) of the winglet  114  may be at a distance L2 of at least about 0.4, but no more than about 0.6, of L1 from the leading edge  76 . 
     Put another way, the widest point (for example at section A-A) of the winglet  114  may be at a chord distance of L2 from the leading edge  76 , where L2 has a value of at least about 0.4 L1 but no more than about 0.6 L1. 
     As shown in  FIG. 5 a   , along the length of the winglet  114 , the winglet  114  is narrower than a distance wbA between the pressure surface  91  and the suction surface  89  in the corresponding region of winglet  114 . That is to say, along the length of the winglet  114 , the winglet is recessed beneath the pressure surface  91 . Put another way, along the length of the winglet  114 , the winglet does not extend beyond the limit of the pressure surface  91 . 
     The widest point (for example at section A-A) of the winglet  114  may have a width w3A of at least about 0.8 wbA but no more than about 0.95 wbA. 
     The tip wall  106  defines a tip surface  118  which extends from the aerofoil leading edge  76  to the aerofoil trailing edge  78 . At the widest point of the winglet  114  the transition region  108  of the pressure surface wall  90  extends from the shoulder  104  in a direction towards the suction surface  89 . As shown in  FIGS. 5 a , 5 b   , at an inflexion point  120  the transition region  108  then curves to extend in a direction away from the suction surface  89  toward the tip surface  118 . Hence the winglet  114  overhangs the transition region  108 . Put another way, in the region of the winglet  114 , the transition region  108  forms a channel. That is to say, in the region of the winglet  114 , the transition region  108  defines a re-entrant feature which defines the overhang of the winglet  114 . 
     The tip portion  100  further comprises an inflexion line  122  defined by a change in curvature on the pressure surface  91  and along with the inflexion point  120  is with respect to the cross-section view of  FIGS. 5 a , 5 b   . The inflexion line  122  extends between the leading edge  76  and the trailing edge  78 . The inflexion points  120  are provided on the inflexion line  122 . Put another way, the inflexion line  122  is defined by a series of curvature inflexion points  120  which extends from the leading edge  76  to the trailing edge  78  on the pressure surface wall  90  in the tip portion  100 . 
     As shown in  FIGS. 5 a , 5 b   , the inflexion line  122  may be provided a distance h2A, h2B from the tip surface, and the shoulder  104  may be provided a distance h1A, h1B of at least about 1.5 times, but no more than about 2.7 times, the distance h2A of the inflexion line  122  from the tip surface  118 . 
     Put another way, as shown in  FIGS. 5 a , 5 b   , the inflexion line  122  may be provided a distance h2A, h2B from the tip surface, and the shoulder  104  may be provided a distance h1A, h1B from the tip surface  118 , where h1A, h1B may have a value of at least about 1.5 h2A but no more than about 2.7 h2A. 
     The inflexion line  122  at the widest point of the winglet  114  may be provided a distance w2A of at least about 0.8, but no more than about 0.95, of w3A from the suction surface  89 . 
     Put another way, the inflexion line  122  at the widest point of the winglet  114  may be provided a distance w2A from the suction surface  89 , wherein w2A may have a value of at least about 0.8 w3A but no more than about 0.95 w3A. 
     With reference to a compressor rotor assembly for a turbine engine comprising a compressor aerofoil according to the present disclosure, and as described above and shown in  FIGS. 5 a , 5 b   , the compressor rotor assembly comprises a casing  50  and a compressor aerofoil  70  wherein the casing  50  and the compressor aerofoil  70  define a tip gap, hg, defined between the tip surface and the casing. 
     In such an example the distance h2A, h2B from the inflexion line  122  to the tip surface has a value of at least about 1.5, but no more than about 3.5, of the tip gap hg. Put another way the distance h2A, h2B from the inflexion line  122  to the tip surface may have a value of at least about 1.5 hg but no more than about 3.5 hg. That is to say, the distance h2A, h2B from the inflexion line  122  to the tip surface may have a value of at least about 1.5 but no more than about 3.5 of a predetermined (i.e. desired) tip clearance gap hg. 
     In operation in a compressor, the geometry of the compressor aerofoil of the present disclosure differs in two ways from arrangements of the related art, for example as shown in  FIG. 1 . 
     The inflexions  120  (i.e. inflexion line  122 ) in the transition region  108  which forms the overhanging winglet  114  inhibits primary flow leakage by virtue of intrusion of the winglet  114  into the air flow directed radially (or with a radial component) along the pressure surface  91  towards the tip portion  100 , and hence the tip flow vortex formed is of lower intensity than those of the related art. 
     The squealer  110 , being narrower than the overall width of the main body  102 , results in the pressure difference across the tip surface  118  being lower than if the tip surface  118  had the same cross section as the main body  102 . Hence secondary flow across the tip surface  118  will be less than in examples of the related art, and the primary flow vortex formed is consequently of lesser intensity as there is less secondary flow feeding it than in examples of the related art. 
     Additionally, since the winglet  114  of the aerofoil  70  is within the boundary of the walls of main body  102  (i.e. as shown in  FIG. 5 a   , is recessed below surface of the main body walls  88 ,  90 , and does not extend beyond the main body walls  88 ,  90 ), the configuration is frictionally less resistant to movement than an example of the related art in which the winglet  114  extends beyond boundary of the walls of the main body  102 . That is to say, since the winglet  114  of the present disclosure has a relatively small surface area, the frictional and aerodynamic forces generated by it with respect to the casing  50  will be less than in examples of the related art. 
     Thus the amount of over tip leakage flow flowing over the tip surface  118  is reduced, as is potential frictional resistance. The reduction in the amount of over tip leakage flow is beneficial because there is then less interaction with (e.g. feeding of) the over tip leakage vortex. 
     Hence there is provided an aerofoil rotor blade and/or stator vane for a compressor for a turbine engine configured to reduce tip leakage flow and hence reduce strength of the interaction between the leakage flow and the main stream flow, which in turn reduces overall loss in efficiency. 
     As described, the aerofoil is reduced in thickness towards its tip to form a squealer on the suction (convex) side of the aerofoil, which reduces the pressure difference across the tip and hence reduces secondary leakage flow. The winglet is provided on the pressure side near the leading edge which acts to diminish primary leakage flow. Together, these features reduce the tip leakage mass flow thus diminishing the strength of the interaction between the leakage flow and the main stream flow which in turn reduces the loss in efficiency. 
     Hence the compressor aerofoil of the present disclosure results in a compressor of greater efficiency compared to known arrangements. 
     Attention is directed to all papers and documents which are filed concurrently with or previous to this specification in connection with this application and which are open to public inspection with this specification, and the contents of all such papers and documents are incorporated herein by reference. 
     All of the features disclosed in this specification (including any accompanying claims, abstract and drawings), and/or all of the steps of any method or process so disclosed, may be combined in any combination, except combinations where at least about some of such features and/or steps are mutually exclusive. 
     Each feature disclosed in this specification (including any accompanying claims, abstract and drawings) may be replaced by alternative features serving the same, equivalent or similar purpose, unless expressly stated otherwise. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features. 
     The invention is not restricted to the details of the foregoing embodiment(s). The invention extends to any novel one, or any novel combination, of the features disclosed in this specification (including any accompanying claims, abstract and drawings), or to any novel one, or any novel combination, of the steps of any method or process so disclosed.