AEROFOIL FOR AN AXIAL FLOW TURBOMACHINE

Abstract: A turbine stator vane (41) for use in an axial flow gas turbine. The vane has an aerofoil the pressure face of which is convex between platform (45) and tip (46) regions in a plane (48) which extends both radially of the turbine and transversely of the general working fluid flow direction between the vanes. The trailing edge (43) of the aerofoil is straight from platform to tip, and the spanwise convex and concave curvatures of the aerofoil pressure and suction surfaces respectively are achieved by rotational displacement of the aerofoil sections about the straight trailing edge. However, the axial width (W) of the aerofoil is substantially constant over substantially all of the aerofoil radial height and the chord line at mid-height aerofoil sections (44) is shorter than the chord lines in aerofoil sections at platform or tip regions. Reducing chord length at the mid-height region in this way lowers aerodynamic profile losses without unduly affecting vane performance. Also disclosed is a turbine rotor blade designed to form a stage pair with the stator vane.

Description: AEROFOIL FOR AN AXIAL FLOW TURBOMACHINE Technical Field of the Invention The invention relates to improved aerofoil shapes for use as stator vanes or rotor blades in turbines of axial flow turbomachmes, such as gas turbine engines Background Turbomachmes are used to add energy to a working fluid and/or to extract energy from it Accordingly, they may comprise compressors and/or turbines For example, gas turbine engines typically comprise three main sections, a compressor section, a combustion section and a turbine section Nr from the atmosphere is drawn into and is compressed by the compressor It is then passed into the combustion section where fuel is added and the mixture ignited so that an energised working fluid is created in the form of a pressurised hot gas The working fluid passes from the combustion section to the turbine section where its energy is extracted by turbine blades and used to turn the compressor via a turbine shaft and do additional work Eventually the working gas, now at much reduced temperature and pressure, is discharged to atmosphere via an exhaust duct system In the present invention, the means used to convert turbine working fluid energy into shaft rotational energy is a system of aerofoils comprising axial flow rotor blades and stator vanes The rotor blades and stator vanes are arranged to intercept the working fluid as a number of axially successive annular rows Each rotor blade is attached to a turbine rotor disc or drum via a blade root portion, the disc or drum being mounted on a rotor shaft, the longitudinal centre line of which defines the rotational axis of the turbine The stator vanes are fixed, e g , to a circumscribing turbine casing or to an inner static drum, and rows of vanes and blades alternate with each other so that each row of blades is paired with a preceding row of stator vanes Each such pair of rows is collectively termed a stage and a turbine will comprise at least one stage Whereas the function of the rotor blade rows is to extract energy from the working fluid and transfer it to a turbine rotor disc or drum and hence to the shaft, the function of the stator vanes is to smooth the flow of the working fluid and then direct it at an optimum outlet angle to the rotor blades so that efficient energy transfer may be achieved there to turn the rotor The efficiency with which both blades and vanes perform their function is of vital importance in determining stage efficiency In the gas turbine engine field, aerofoils of turbine vanes and blades have respective generic types of cross-section profile and may bear a strong visual likeness one to another, notwithstanding scale differences usually dependent upon engine size However, on inspection it is found there are measurable differences of aerofoil profiles not only between engines of different make and type but also between turbine stages of the same engine Further, such differences may have significant effects on turbine efficiency Similarly, there are differences in other aspects of turbine stage design which alone or in combination also have an effect Small differences in such design features, which may appear minimal or unimportant to those unskilled in the art, may in fact have a significant effect on turbine stage performance Hence, vane and blade geometrical shapes, their positional relationships to each other and also to the stream of working fluid have an effect on turbine efficiency and thus on turbomachine efficiency overall In known state-of-the art gas turbine engines, the turbine stage efficiency is currently in the region of 90% and at such high efficiency it is regarded as now very difficult to improve by even parts of 1% Nonetheless, it is an object of the present invention to increase turbine stage efficiency by a significant amount In part, the present invention incorporates and improves upon previous teachings in respect of so called "Controlled Flow" principles by the present inventor and others In particular, see patent GB 2 295 860 B "Turbine Blade", directed particularly at steam turbines Other patents showing similar principles include US 5,326,221 Amyot et al (for steam turbines) and US 4,741,667 Price et al (for gas turbines) Definitions For the purposes of the present invention, it will be understood that the term "vane" refers to the stator blades which precede the rotor blades in turbomachines, including the so- called "nozzle guide vanes" in gas turbine engines, which function to direct the hot gases from the combustor onto the first stage of turbine rotor blades Also, when the word "blade" is used without the qualifying words "stator" or "rotor", it should be taken to mean "rotor blade" The radially innermost extremity of the aerofoil portions of axial flow blades and vanes will be termed their " platform region" (even though the radially innermost portion of a gas turbine rotor blade is usually termed a "root"), and the radially outermost extremities of their aerofoil portions will be termed their "tip region" (despite the fact that blades and vanes can have radially outer shrouds). The "pressure" surface of an aerofoil section shape is its concave side and the "suction" surface is its convex side. A "prismatic" aerofoil is designed such that the notional aerofoil sections of the blade or vane, each considered orthogonal to a radial line from the turbine axis, have the same shape from the aerofoil platform region to the aerofoil tip region, are not skewed, i e , have the same setting angle from the platform region to the tip region, and are "stacked" one on top of another so that their leading edges and their trailing edges collectively form straight lines in the radial direction The outlet angle α of an aerofoil is the angle, relative to the circumferential direction of the rotor, that the working fluid leaves a vane or blade row and is derived from the relationship α = sin ' ^T/P), where T is the throat dimension and P is the pitch dimension Throat dimension T is defined as the shortest line extending from one aerofoil trailing edge normal to the suction surface of the adjacent aerofoil in the same row, whereas pitch dimension P is the circumferential distance from one aerofoil trailing edge to the adjacent aerofoil trailing edge m the same row at a specified radial distance from the platform region of the aerofoil The setting angle β is the angle through which any particular aerofoil section at a station along the height or span of the aerofoil is displaced in its own plane from a predetermined zero datum The datum may, for example, be taken as being where the aerofoil section has the same "stagger angle", 1 e the same oπentation relative to the turbine axis, as a known prismatic aerofoil in a known turbine utilising such aerofoils The "chord line" is the shortest line tangent to leading and trailing edge radii of an aerofoil section The "chord length" is the distance between two lines normal to the chord line and passing through the points where the chord line touches the leading and trailing edges respectively The "axial width" of an aerofoil is the axial distance between its leading and trailing edges, i e , the distance between its leading and trailing edges as measured along the rotational axis of the turbine Summary of the Invention According to a first aspect of the present invention, a turbine stator vane is for use in a ring of similar vanes arranged in an axial flow turbine having an annular path for a turbine working fluid, the vane comprising an aerofoil spanning the annular path and having a radially inner platform region, a radially outer tip region, an axially forward leading edge and an axially rearward trailing edge, the aerofoil having a pressure surface and a suction surface which are respectively convex and concave between the platform region and the tip region in a plane extending both radially of the annular path and transversely of the axial direction, the trailing edge of the aerofoil being straight from the platform region to the tip region and oriented radially c f the annular path, and said convex and concave curvatures of the aerofoil pressure and suction surfaces being achieved by rotational displacement of the aerofoil sections about the straight trailing edge, the axial width of the aerofoil being substantially constant over substantially all of the aerofoil radial height and the chord line at mid-height aerofoil sections being shorter than the chord lines in aerofoil sections at platform or tip regions. In the context of a gas turbine engine, the invention in its first aspect may be applied to the aerofoils of nozzle guide vanes in the first or high pressure stage of the turbine, but also to the stator vanes of succeeding stages. Because the chord line at mid-height aerofoil sections is shorter than the chord lines in aerofoil sections at both the platform regions and the tip regions, the aerofoil exhibits a so-called "compound lean" appearance when viewed on its leading edge, in which the aerofoil is skewed in the same circumferential direction at both radial extremities. In accordance with a second aspect of the invention, if the aerofoil is that of a nozzle guide vane at the entry to a gas turbine, the aerofoil is preferably positioned in relation to the axial length of the turbine such that the trailing edge of the aerofoil is in a divergent part of the gas flow passage, whereby the trailing edge of the aerofoil is substantially longer than its leading edge. In the case of a nozzle guide vane aerofoil, the aerofoil's platform and tip outlet angles are preferably of substantially the same value, for example, not more than about 10 degrees, preferably in the range 8-10 degrees. The aerofoil's outlet angle at mid-height of the aerofoil may be in the range 13-16 degrees, preferably approximately 14 degrees Conveniently, the aerofoil is of approximately constant aerofoil cross-section from its platform region to its tip region In accordance with a further aspect of the invention, a turbine stage comprises a row of stator vanes as described above, and a row of rotor blades in flow sequence with the vanes, in which the blades comprise aerofoils having a radially inner platform region, a radially outer tip region, an axially forward leading edge and an axially rearward trailing edge, each blade aerofoil having a pressure surface and a suction surface which are respectively convex and concave between the platform region and the tip region in a plane extending both radially of the annular path and transversely of the axial direction, said convex and concave curvatures of the aerofoil pressure and suction surfaces being achieved by rotational displacement of the aerofoil sections about a radial line through the aerofoil, each aerofoil having outlet angles which are smaller near its platform and tip regions than at mid-height. From an aerodynamic point of view, each blade aerofoil ideally has a radially oriented straight trailing edge, the rotational displacement of the aerofoil sections being about the straight trailing edge, though this ideal may be compromised by the dynamic design requirements of the blades. To reduce dynamic loading at the root fixings and the platform, the blade aerofoil may taper from its platform region to its tip region, such that its chord length reduces over the blade aerofoil's radial height from a maximum at its platform region to a minimum at its tip region and its leading edge has a backward lean in the axial direction. In yet another aspect, the invention provides a turbine stage comprising a row of nozzle guide vanes having aerofoils as described above, and a row of rotor blades in flow sequence with the vanes, in which the blade aerofoil platform and tip outlet angles are in the range 14- 17 degrees, preferably about 16 degrees. The blade aerofoil outlet angle at mid-height of the aerofoil may be in the range 18-21 degrees, preferably about 19 degrees. The invention is believed applicable whether the aerofoils are shrouded or unshrouded, i.e., whether the aerofoils are joined to a structure forming an outer wall of the passages between adjacent aerofoils, or are not so joined, but are free at their radially outer or tip regions Further aspects of the invention will be apparent from the following description and claims Brief Description of the Drawings Exemplary embodiments of the invention will now be described with reference to the accompanying drawings, in which Fig 1 is a computer generated perspective view of a prior art aerofoil shape utilising the "Controlled Flow" principle, Fig 2 is a sketch of a poor art gas turbine vane aerofoil as viewed from the tip end of the aerofoil towards the platform end, Fig 3 is an axial side view of the vane aerofoil of Fig 2 showing its position in the turbine passage, Fig 4 is a view similar to Fig 2, but of a vane aerofoil shaped according to the present invention, Fig 5 is an axial side view of the vane aerofoil of Fig 4, Fig 6 is a view similar to Fig 5, but of a different embodiment of the invention, Fig 7 is a diagram showing corresponding elemental sections of two adjacent aerofoils to illustrate the concept of outlet angle, which is important in relation to an aspect of the invention, Fig 8 is a computer generated perspective view of an aerofoil of a gas turbine engine nozzle guide vane shaped in accordance with the present invention, and Fig 9 is a computer generated perspective view of an aerofoil of a gas turbine engine rotor blade shaped in accordance with the present invention Detailed Description of Preferred Embodiments Figure 1, extracted from Patent Number GB 2 295 860 B, to which the reader is referred for further details, shows the aerofoil of a steam turbine stator blade or vane which is shaped in accordance with the principles of the invention disclosed in that patent The grid pattern shown on the surface is computer-generated and serves to emphasize the curved formation of the aerofoil. It has a straight trailing edge 25 like previously known aerofoils, but the remainder of the aerofoil, and in particular the leading edge 24, is not straight but is curved in a manner such that the pressure surface 26 of the aerofoil is convex between platform region 35 and tip region 37 in a plane which extends both radially of the turbine and transversely of the general steam flow direction between the aerofoils. One such plane 31 is indicated, the convex curvature in this plane on the pressure surface 26 being obscured but conforming to that at the leading edge 24 More specifically, relative to a prismatic aerofoil, the individual aerofoil sections 33 may be considered as being rotated in their own planes about the trailing edge 25 by a setting angle which is positive in the central part of the radial height, and negative in the platform and tip portions 'Positive' is taken to be a rotation toward the pressure surface 26 and 'negative' is taken to be a rotation toward the suction surface 27 In Fig. 1, the setting angle varies in parabolic manner from about minus 2 5° at the platform and tip regions to plus 2 5° at the centre of the radial height, referred to a datum stagger angle of 48 5° It would to some extent be acceptable to skew the aerofoil sections about some other axis than the trailing edge 25, for example a radial line through the leading edge 24 or some intermediate axis However, the choice of the trailing edge as the axis about which the aerofoil sections are rotated has several advantages It keeps the critical interspace gap between the fixed vanes and the downstream rotor blades constant This gap has an important influence upon the unsteady aerodynamic forces on the moving blade and also on the stage efficiency via boundary layer growth on the radially inner and outer turbine passage walls (termed the "end walls") Secondly, by building the curvature largely into the leading edge a "Compound Lean" effect is incorporated into the leading edge area of the aerofoil where secondary flows are generated These secondary flows comprise vortices in parallel with the nain flow, the vortices being near the end walls between adjacent fixed blades By the us of the compound curved aerofoil of Fig 1, over the inner (1 e lower) half of the aerofoil height the pressure surface points radially inwards, and over the outer half of the aerofoil height the pressure surface points radially outwards The body forces exerted on the flow are counteracted by higher static pressures on the end walls This results in lower velocities near the end walls and hence lower frictional losses Figs 2 and 3 show a pπor art gas turbine vane whose aerofoil 1 is designed on similar principles to that of Fig 1 Dashed line 2 represents the axial centre line of the turbine, 7 and 8 are radially inner and outer walls defining the turbine working fluid passage, 4 is the leading edge at vane mid-height region, 5 and 6 are platform and tip regions respectively, the arrow D indicates the overall direction of flow of the working fluid and the angle which line L makes with the axis 2 represents the prismatic aerofoil datum stagger angle As in Fig 1, the vane aerofoil sections are stacked about a straight, radially oriented trailing edge 3 and are rotated or "skewed" toward the closed position at leading edge platform and tip, I e at leading edge platform and tip the setting angle is at its greatest negative value -β relative to the datum line L and the throat dimension T (see Fig 7) is at a minimum For clarity, Fig 2 shows an exaggerated platform and tip skew However at the mid-height of the aerofoil, the setting angle is at its greatest positive value +β Thus, the leading edge at vane mid-height region, 4, is axially forward of the leading edge at platform and tip regions, 5 and 6, by an amount "X" This means that even though the chord lines of all the aerofoil sections are the same length, the axial width W of the aerofoil (l e , the distance between its leading and trailing edges 4 3 in the axial direction) varies bv X over the radial height of the aerofoil Referring now to Figs 4, 5 and 8 the views in Figs 4 and 5 are similar to Figs 2 and 3 but of a vane aerofoil 41 in accordance with the present invention which is based on a modification of the Controlled Flow principle Fig 8 is a perspective view on the trailing edge of the aerofoil 41, the aerofoil being overlaid by a computer-generated grid as in Fig 1 Coordinates for the computer generated grid are indicated as X and Z, X being the axial direction and Z being the radial direction As in Fig 1, the trailing edge 43 is radially oπented and straight and the pressure face 47 of the aerofoil is convex between platform 45 and tip 46 in a plane 48 which extends both radially of the turbine and transversely of axial centre line 2, this being achieved by rotational displacement of the aerofoil sections 49 about the radial trailing edge However, it can be seen in Figs 4 and 5 that the leading edge 44 at mid-height position is not forward of the platform and tip regions but is substantially in line with them, with respect to the axial direction defined by axis 2 Since the trailing edge 43 is straight, the axial width W of the vane aerofoil is consequently substantially constant over substantially all of the aerofoil radial height and the chord lines at mid-height aerofoil sections are shorter than the chord lines in aerofoil sections at platform or tip regions It is found that reducing chord length at the mid-height region in this way has the effect of advantageously lowering aerodynamic profile losses without unduly affecting vane performance This is because the "wetted" area, and hence friction loss, is reduced Fig 6 illustrates a further embodiment of the invention, applicable to first stage vane aerofoils 61 at the entry to a turbine As in Figs 4, 5 and 8, the pressure face of the aerofoil is convex between platform 65 and tip 66, the leading edge 64 at mid-height position is substantially axially in line with the platform and tip regions, and the radially oπented trailing edge 63 is straight However, it has been found advantageous to posit'on the aerofoil 61 in relation to the axial length of the turbine such that its trailing edge 63 is in a divergent part of the gas flow passage, so causing the trailing edge 63 to be substantially longer than the leading edge 64 Although this is normal for turbine second ana subsequent stages, it is not normal tor a first stage Usually, as shown in F>g 5, ^ ~ s stage vanes have a leading edge longer than, or substantially the same length as, the trailing edge Clear'v, stac ing the vane aerofoil sections as described wit 4 reference *o F'gs, -t + j '-> -in 8, so that they nave smaller outlet angles at the platform and tip regions than at r α-heignt may create flow incidence problems onto the succeeding rotor blade row and it is therefore necessary to apply similar Controlled Flow principles to the rotor blade aerofoils Hence, Fig 9 is a perspective view similar to Fig 8, but of a high pressure turbine rotor blade aerofoil 90 situated axially adjacent to and immediately downstream of the vane aerofoil of Fig 8, i e , together with vane aerofoil 41, blade aerofoil 90 comprises the first stage of a gas turbine Similarly to the vane aerofoil 41, blade aerofoil 90 has a straight trailing edge 91 oriented in the radial direction Referred again to a plane 95 which extends radially of the turbine and transversely of the rotational axis of the turbine, the pressure surface 92 is convex between platform region 93 and tip region 94 and the suction surface 96 is concave As before, the spanwise convex and concave shapes of the pressure and suction surfaces respectively are achieved by rotational displacement of the aerofoil sections 97 about the trailing edge 91 Furthermore, by virtue of the radially convex and concave shapes of the pressure and suction surfaces 92 and 96, the blade aerofoil 90 as a whole is skewed towards the 'throat closed' position at both the tip and platform regions, its outlet angles again being smaller at the platform and tip regions than at mid-height Despite these similarities, rotor blade aerofoil 90 has a somewhat different appearance from nozzle guide vane aerofoil 41 and in particular the leading edge 98 of blade aerofoil 90 has a different appearance from the leading edge 44 of vane aerofoil 41 Unlike the vane aerofoil 41, the blade aerofoil 90 tapers from platform to tip, I e , its axial width, and hence its chord length, reduces over the aerofoil's radial height from a maximum at the platform region 93 to a minimum at the tip region 94 Such tapeπng of the blade aerofoil in the radial direction is intended to reduce the centπfugally induced stresses experienced in the platform region and in the root fixings of the blade duπng operation of the gas turbine, because the mass of the radially outer portion of the blade aerofoil is reduced Since the aerofoil has a radially oπented straight trailing edge 91, its reduction in axial width with radial distance from the platform region means that its leading edge 98 has a backward lean in the axial direction, and this is shown in Fig 9 Note that if necessary for the reduction oϊ eccentric bending stresses generated during rotation of the blade, the radially convex and concave pressure and suction surfaces - 1; respectively of the blade aerofoil can alternatively be achieved by rotating the aerofoil sections about a radial line other than a radial line though the trailing edge - e.g., a line through the centroid of the notional prismatic aerofoil. This would result in a curved trailing edge. With regard to vane aerofoil setting angles and thus outlet angles, Fig 7 shows corresponding elemental sections of two adjacent vane aerofoils to illustrate outlet angle α, T being the throat dimension and P being the blade pitch. Typically, vane aerofoils are designed with setting angles (relative to axial direction) which result in larger outlet angles at the tip region than at the platform region. However, it is found advantageous in the present invention to have vane aerofoil platform and tip outlet angles of substantially the same value. Also, it is surprising that these outlet angles, being not more than about 10 degrees and preferably in the range 8-10 degrees, are less than is suggested in known gas turbines. Similarly, the outlet angle at a mid-height region for a vane aerofoil in accordance with the invention is in the range 13-16 degrees, or approximately 14 degrees, and this is iess than might be expected for "Controlled Fiow " designs in a gas turbine engine. This variation in outlet angle α over the radial height of the aerofoil is not readily- apparent from the perspective of Fig. 8, but can be readily appreciated by reference to Fig 4 As stated, vanes and blade rows co-operate as a stage pair Therefore, amongst othei things, vane and blade aerofoil angles must be matched for best efficiency It is found that suitable outlet angles for blade aerofoils in a turbine stage according to the invention are blade aerofoil platform and tip, α in the range 1 -17°, preferably ~ 16" blade aerofoii at mid-height; α in the range 18-2 ϊ preferably α ~ 19" The design process for Controlled Flow vane and blade profiles considers firstly the vanes and secondly the blades, each separately, then finally together as a matching pair to achieve best overall sta <r s n erformance The are usually designed through an iterativ e process with inputs from physically or mathematically defined design guidelines and intuitive experience, all compiomiscd by requirements for reasonable aerofoil strength, vibration characteristics, accommodation of internal cooling passages, etc. In the present invention the reduced chordal length at mid height is a further complication affecting the detail of profile shapes. In practice each gas turbine engine maker will generally have its own design rules and will settle for profile shapes within those rules. It is a feature of the present invention that a particularly effective set of aerofoil section profiles (from platform to tip) are achieved by adhering to X-Y co-ordinates, within certain dimensional limits of variation of X and Y, as laid down below in Tables 1 to 3 (for vane aerofoil platform region, mid-height, and tip region respectively) and Tables 4 to 6 (for blade aerofoil platform, mid-height and tip respectively). The dimensional limits of variation mentioned are plus or minus 5% of chordal length, e.g. for a chord of 30mm the X and Y dimensions may vary by plus or minus 1.5 mm. For scaling purposes the X-Y co-ordinates of Tables 1 to 6 may be multiplied by a predetermined number or scaling factor to achieve similar aerodynamic performance from either larger or smaller vanes and blades. It will be known to those skilled in the art that simple linear scaling of vanes and blades does not indicate similar linear scaling of, for example, engine power (which would, in comparison, scale to the square). Nevertheless, with appropriate scaling, the aerofoil section profile shapes and angles described in the Tables may be used for any size gas turbine engine Further, it should be noted that the invention is not limited to the particular aerofoil section profile shapes and angles described in the Tables. Table 1 - Vane platform -744137E+00 1 2.06360E+0I -686691 E+00' 366105E+01 -7 ?15_24E+00> 215114E+01; -656406E+00 354220E+01 -8.19843E+00J 2.23826E+01, -6.27495E+Crø 3 ~ 42301E+01 -8.59130E-K)0 2 2496E-01 -59978SE+00< 3.30354E+0I -899422E J -0O " 24 " lll9E+όr 318382E+01 -940751E^00 249694E- t -01 -547391 E+00 306390E+01 -983144E+00 258216E 01 -522446E+00294382E+01 02662E+01 2.66683E+01 -498187E+002.82359E+01 -107120F.+01 27S092E+01 , -4745Ϊ9E+00270 25E+01 ■111690E+01 283442E+01 1 -451358E 00258281E+01 -9.20219E-01S 6.80254E-01; -2.17108E+01; 4.77245E+01 -1.20376E+00I 149785E+00! -207075E+01I 4.82908E+01: -1.49057E+00I 2.31430E+00! -1.95945E+01i4.85871E+0li -1.78042E+00I 3.12968E+00' -1.84424E+01! 4.85930E+01! -2.07308E+00 1 3.94406E+00I ! -1.73198E+01! 4.83316E+0f > -2.36837E+00' 4.75749E+001 -1.6271 lE+Oli 4.78520E+01 -2.66652E+005.56987E+00 -1.53155E+0l| 4.72055E+01 Table 2 - Vane Mid-Height (Continued) Table 3 - Vane Tip -7.42662E+00 2.20050E+01 ; -6.00671E+00; 3.64249E+01 I-7.80735E+00! 2.29398E+0 ] -5.72164E+00i 3.51299E+01 I-8.19838E+00! 2.38703E+01 -5.45003E+00! 3.38319E+01 -8.60013E+00! 2.47963E+01 -5.19008E+00I 3.25316E+01 •9.01301E+00I 2.57173E+01 -4.94019E+00I 3.12294E+01 -9.43738E+00! 2 6633 1E+01 1 -4 69895E+00! 2.99255E+01 ; -9.87354E+00! 2.75434E+01 -4.465 1 lE+θ ! 2.86202E+01 ' 1.03218E+OU 2.84477E+01 -4.23758E+00I 2.73138E+01 -1 07822E+01 , 2.93459E+01 -4.01538E+00 2.60065E+01 ; ■ 1.12551E+01 ! 3.02377E+01 ! -3.79767E+00I 2.46985E+01 Table 4 - Blade platform X-Y (mm) X-Y (mm) ! X-Y (mm) ! X-Y (mm) 1.28295E+01! -2.06567E+01 1 -5.47895E+00I -1.25458E+00 1 1.28472E+01 -2.07835E+01I -6.22214E+00! -1.38526E+00 1.28238E+01 -2.09093E+01 -7.12124E+001 -1.60851E+00 1.27618E+01 -2.10213E+01 -8.00453E+00i -1.88689E+00' 1.26675E+01 -2.11080E+01 -8.88112E+00 -2.18359E+00! 1.25506E+01 -2.11605E+01 -977936E+00 -237509E+00 ; 1.24230E+01 -2.11734E+01 ■1.04283E+01 -1.82815E+00 1.22978E+01 -2.11453E+01 -1.05255E+01I -9.14849E-01 1.21879E+01 -2.10791 E+01 . -1.04622E+01 8.34471E-03 1.21044E+01 -2.09815E+01 : -1.03081E+01I 9.21463E-01 ! 1.19650E+01 -2.06548E+01 1 -1.00877E+0H 1.82088E+00: 1.17433E+01I -2.01168E+01 -9.80728E+00 2.72388E+00 U5198E+01 -1.95794E+01 -9.47649E+00 ' 3.60966E+00 1.12944E+01: -1.90428E+01 -9.09341E+00 1 4.47419E+00 j 1.10670E+01! -1.85071E+01! 1 -8.65471E+00! 5.31188E+00 | 1.08374E+01I -179724E+01 . -8.15726E+00I 6.11594E+00 1.06057E+0F -1.74385E+01 -7.59696E+00 687723E+00 1.03719E+01; -1.69055E+01 _-696847E+00__ 7^58293E+00 1.01360E+01 -1.63735E+01. -627170E+00; 8.22133E+00 j 9.89802E+00J -1.58425E+01 -5.50357E+OOI 8.77210E+00' 9.65799E+00J -1.53123E+OU -4.65844E+00I 9.19564E+00I Table 5 - Blade Mid-Height X (mm) Y (mm) X (mm) Y (mm) 1.23893E+01 ■1.90680E+01 -5.26410E+00 -9.42599E-01 1.24106E+01 -1.91945E+01 -6.07693E+00 -9.99129E-01 1.23908E+01 -1.93213E+01 -6.88482E+00 -1.10519E+00 1.23320E+01 -1.94353E+01 -7.69007E+00 -1.23009E+00 1.22401E+01 -1.95249E+01 -8.50128E+00 -1.19699E+00 1.21246E+01 -1.95809E+O1 -9.15014E+00 -7.30371E-01 1.19974E+01 ■1.95975E+01 -9.39980E+00 4.05600E-02 1.18714E+01 -1.95731E+01 -9.41515E+00 8.54467E-01 1.17597E+01 -1.95101E+01 ■9.31880E+00 1.66339E+00 1.16735E+01 ■1.94150E+01 ■9.14888E+00 2.46020E+00 1.15336E+01 -1.91105E+01 -8.91008E+00 3.25372E+00 1.13138E+01 -1.86153E+01 ■8.60472E+00 4.03865E+00 1.10923E+01 ■1.81208E+01 -8.23789E+00 4.79679E+00 1.08689E+01 •1.76272E+01 ■7.80816E+00 5.521 lOE+00 1.06435E+01 -1.71345E+01 ■7.31397E+O0 6.20301E+00 1.04161E+01 -1.66427E+01 •6.75336E+00 6.83138E+00 1.01867E+01 -1.61519E+01 ■6.12475E+00 7.39157E+00 9.95554E+00 ■1.56618E+01 •5.42703E+00 7.86261 E+00 9.72272E+00 -1.51726E+01 -4.66495E+00 8.21995E+00 9.48835E+O0 -1.46841E+01 •3.85360E+00 8.44356E+O0 9.25257E+00 -1.41963E+01 -3.01568E+00 8.52181E+00 9.13417E+00 -1.39526E+01 -2.59484E+00 8.50446E+00 9.01546E+00 ■1.37091E+01 -2.17720E+00 8.44981E+00 8.77711E+00 ■1.32225E+01 -1.36345E+00 8.23479E+00' 8.53764E+00 -1.27365E+01 -5.93556E-01 7.89434E+00 8.29716E+00 -1.22510E+01 1.21631E-01 7.45004E+00 8.05554E+00 -1.17660E+01 7.79230E-01 6.92404E+00 7.81263E+00 -1.12817E+01 1.38073E+00 6.33459E+00 7.56831E+00 -1.07981E+01 1.93083E+00 5.69680E+00 7.32222E+00 -1.03154E+01 2.43677E+00I 5.02337E+00! 7.07308E+00 -9.83425E+00 2.90625E+00 4.32400E+00 6.82103E+00 -9.35463 E+00 3.34479E+00 3 60480E+00 6.56549E+00 -8.87686E+00 3 75675E+00 2 87004E+00 6.30517E+00 -8 40168E+00 4 14585E+00I 2.12291E+00 6.03889E+00 -7.92982E+001 4.51422E+00I 1.36535E+00- 5.76538E+00! -7.4621 1E+00! 4.86527E+00! 5.99590E-01 5.48326E+00I -6.99954E+00 ! 5.20428E+00I -1.71578E-01 5.19097E+00I -6.54334E+00I 5 53316E+00! -9 47124E-01 4.88691E+00I -6.09490E+00 5.85259E+001 -1.72660E+00 4.56958E+00 -5.65576E+00ι 6 16253E+00! -2.50991E+00 4.23770E+00 -5.22751E+00! 6.46670E+00 -3.29547E+00' 3.89030E+00 -4.81 176E+00 6.76906E+00 -4.08174E+00 Table 6 - Blade Tip X (mm) Y (mm) X (mm) Y (mm) 1.07073E+01 -2.13251E+01 -4.72853E+00 1.30247E+00 1.07213E+01 -2.14512E+01 -5.57007E+00 1.23777E+00I 1.06952E+01 -2.15754E+01 -6.40589E+00 1.12234E+00 1.06316E+01 -2.16852E+01 -7.24159E+00 1.00621 E+00 1.05369E+01 -2.17697E+01 -8.08318E+00 1.00430E+00 1 .04205E+01 -2.18203 E+01 -8.77869E+00 .43708E+00 1.02941E+01 -2.18320E+01 -9.01179E+00 2.241 18E+001 1.01704E+01 -2.18037E+01 -9.01241E+00 3.08439E+00i 1.00618E+01 -2.17381E+01 -8.91320E+00 3.92206E+00 9.97906E+00 -2.16418E+01 -8.73634E+00 4.74696E+00 9.83998E+00 -2. 13174E+01 -8.47843E+00 5.55610E+00 9.62101E+00 -2.07816E+01 -8. 13325E+00 6.33782E+00' 9.40378E+00 -2.02451E+011 -7.69912E+00 7.07379E+00 9.18810E+00I -1.97079E+01 ι -7.17450E+001 7.74815E+001 8.97380E+00 •1.91702E+011 -6.56032E+001 8.34195E+00 8.76071 E+00 -1.86320E+01 -5.86147E+00I 8.83306E+00 8.65456E+00 -1.83627E+01 -5.48341E+00 9.03249E+00 8.54864E+00 -1.80934E+01 I -5.θ " 8888E+OOi 9 19697E+00 8.33742E+001 -1.75545E+01 ! -4.26187E+00! 9.40933E+001 8.12689E+00) -1.70153E+01 ! -3740916E+001 9 " 45098E+00J Table 6 - Blade Ti (Continued)

Claims: CLAΓMS 1. A turbine stator vane (41) for use in a ring of similar vanes arranged in an axial flow turbine having an annular path for a turbine working fluid, the vane comprising an aerofoil spanning the annular path and having a radially inner platform region (45), a radially outer tip region (46), an axially forward leading edge (44) and an axially rearward trailing edge (43), the aerofoil having a pressure surface (47) and a suction surface (42) which are respectively convex and concave between the platform region (45) and the tip region (46) in a plane (48) extending both radially of the annular path and transversely of the axial direction, the trailing edge (43) of the aerofoil being straight from the platform region to the tip region and oriented radially of the annular path, and said convex and concave curvatures of the aerofoil pressure and suction surfaces being achieved by rotational displacement of the aerofoil sections about the straight trailing edge, the axial width (W) of the aerofoil being substantially constant over substantially all of the aerofoil radial height and the chord line at mid-height aerofoil sections (44) being shorter than the chord lines in aerofoil sections at platform or tip regions. 2. A turbine stator vane according to claim 1, comprising a nozzle guide vane aerofoil 3. A turbine stator vane according to claim 1 or claim 2, the aerofoil having platform and tip outlet angles of substantially the same value. 4. A turbine stator vane according to any preceding claim, the aerofoil outlet angles at the platform and tip regions being not more than about 10 degrees. 5 A turbine stator vane according to claim 4, the aerofoil outlet angles at the platform and tip regions being in the range 8- 10 degrees 6 A turbine stator vane according to any of claims 1 to 5, the aerofoil outlet angle at mid-height of the aerofoil being in the range 13-16 degrees 7 A turbine stator vane according to claim 6, the aerofoil outlet angle at mid-height of the aerofoil being approximately 14 degrees 8 A turbine stator vane according to any preceding claim, the aerofoil being of approximately constant aerofoil cross-section from its platform region to its tip region 9 A turbine including nozzle guide vanes according to claim 2, in which the nozzle guide vane aerofoils are positioned in relation to the axial length of the turbine such that the trailing edge of the aerofoils are in a divergent part of the gas flow passage, whereby the trailing edges of the aerofoils are substantially longer than their leading edges 10 A turbine stage comprising a row of stator vanes in accordance with any one of claims 1 to 8, and a row of rotor blades in flow sequence with the vanes, in which the blades comprise aerofoils having a radially inner platform region (93), a radially outer tip region (94), an axially forward leading edge (98) and an axially rearward trailing edge (91), each blade aerofoil having a pressure surface (92) and a suction surface (96) which are respectively convex and concave between the platform region (93) and the tip region (94) in a plane (95) extending both radially of the annular path and transversely oi trie axial direction, said convex and concave curvatures of the aerofoil pressure and suction surfaces being achieved by rotational displacement of the aerofoil sections about a radial line through the aerofoil, each aerofoil having outlet angles which are smaller near its platform and tip regions than at mid-height 1 1 A turbine stage according to claim 10, each blade aerofoils having a radiallv oriented straight trailing edge and the rotational displacement of the aerofoil sections being about the straight trailing edge 12. A turbine stage according to claim 10 or claim 11, in which each blade aerofoil tapers from its platform region to its tip region, such that its chord length reduces over the blade aerofoil's radial height from a maximum at its platform region to a minimum at its tip region and its leading edge has a backward lean in the axial direction. 13. A turbine stage comprising a row of stator vanes in accordance with any one of claims 1 to 8, and a row of rotor blades in flow sequence with the vanes, in which the blade aerofoil platform and tip outlet angles are in the range 14-17 degrees. 14. A turbine stage according to claim 13, in which the blade aerofoil platform and tip outlet angles are about 16 degrees. 15. A turbine stage according to claim 13 or claim 14, in which the blade aerofoil outlet angle at mid-height of the aerofoil is in the range 18-21 degrees. 16. A turbine stage according to claim 15, in which the blade aerofoil outlet angle at mid-height of the aerofoil is about 19 degrees. 17. A stator vane for a gas turbine engine whose aerofoil section profiles in X-Y co- ordinates at the platform, mid-height, and tip regions are substantially as shown in Tables 1 to 3, respectively, within dimensional limits of variation of X and Y of ± 5% of chordal length. 18. A rotor blade for a gas turbine engine whose aerofoil section profiles in X-Y co- ordinates at the platform, mid-height, and tip regions are substantially as shown in Tables 4 to 6, respectively, within dimensional limits of variation of X and Y of ± 5% of chorda! length. 19. A turbine stage comprising a row of vanes in accordance with any one of claims 1 to 8, and a row of blades in flow sequence with the vanes, whose blade aerofoil section profiles in X-Y co-ordinates at the blade platform, blade mid-height, and blade tip regions are substantially as shown in Tables 4 to 6, respectively, within dimensional limits of variation of X and Y of ± 5% of chordal length. 20. A turbine stage comprising a row of vanes in accordance with claim 17, and a row of blades in flow sequence with the vanes, whose blade aerofoil section profiles in X-Y coordinates at the platform, mid-height, and tip are substantially as shown in Tables 4 to 6, respectively, within dimensional limits of variation of X and Y of + 5% of chordal length.

Inventors: HALLER BRIAN ROBERT, HALLER, BRIAN, ROBERT
Applicants: ALSTOM POWER NV, HALLER BRIAN ROBERT, ALSTOM POWER N.V., HALLER, BRIAN, ROBERT
Application year: 2001
Main classification: F01D5
IPC codes: F01D5, F01D9