Turbomachine blade

0001The invention relates to a blade for blade grids of turbomachines according to the preamble of claim 1. Such a blade is disclosed in GB 2 004 599 A.

0002According to the current state of the art, the airfoils are produced from a large number of individual profile cuts, the individual profile cuts being adapted to the radially variable flow effects and conditions. The shape of the airfoil is defined by threading the individual profile cuts.

0003As is known, secondary flows and secondary losses occur in vane grids. A wall boundary layer is formed on the lateral boundary walls of the blade grille in the inflow to the meshes and is deflected in the blade channel. The lower-energy fluid in the pre-boundary layer, however, cannot keep the pressure field built up from the profile suction side to the profile pressure side of an adjacent blade, which is essentially determined by the flow outside the edge zones. As a result, with the formation of more curved flow paths, a cross flow occurs from the pressure side to the suction side of the neighboring blade, which continues in a vortex in the blade channel.

0004The transport of the low-energy material of the pre-boundary layer and the overlay with the suction-side profile boundary layer cause additional losses, the so-called edge losses. The intensity of the overflow is decisive for the amount of the losses.

0005There are additional influences of other vortex systems on the transverse movement. In the wall-side, spatial stowage area in front of the front edge of the blade, the local pressure rises sharply in the direction of flow. As a result, the layers of the fluid close to the wall in the inflow boundary layer diverge in zones of lower pressure transversely to the main flow.

0006A backflow occurs in the plane of symmetry of the spatial accumulation area. Due to the interaction with the outside flow, a vortex is created that runs horseshoe-shaped around the profile. The horseshoe vortex therefore consists of a suction-side and a pressure-side branch. The swirl arm on the suction side runs along the profile suction side, has an opposite direction of rotation with respect to the channel swirl and hinders the movement of the channel swirl to the suction side. The pressure-side branch of the horseshoe vertebra has a sense of rotation within the blade channel that is the same as that of the channel vortex. A complex vortex system is created, with the individual vertebrae interacting strongly with one another.

0007In the generic GB 2 004 599 A a proportionally increasing profile thickness is proposed in the housing-side wall area of ​​guide vane grids to reduce the secondary flow losses. This reduces the radial pressure gradient in the outflow plane on the housing, which then lowers the secondary flow losses. However, this design only leads to an improvement in the flow conditions in the case of guide vane gratings in the wall area on the housing side.

0008Similar, proportionally extending thickenings are disclosed in JP 55-142 909 A.

0009Also in the dissertation "Weiß, Andreas P .: The influence of the profile geometry on the development of the secondary flows in turbine grids, dissertation 1993, University of the Federal Armed Forces Munich, pp. 80-81", the influence of the secondary flow by variation of the profile geometry is examined in order to To minimize grid losses.

0010Proceeding from this, it is an object of the invention to provide a generic airfoil of guide and rotor blade grids, the blade geometry of which influences the secondary flow in the sense of minimizing the blade losses.

0011According to the invention the object is solved by the features of claim 1. Alternative solutions are given in the features of independent claims 5 and 8.

0012The invention has the advantage that the vortex system in the peripheral zones is favorably influenced by the aerodynamic thickening of the airfoil downstream of the front edge in the suction and pressure side region and within the airfoil channel. As a result, the intensity of the secondary vertebrae in the leading edge area (horseshoe / channel vertebrae) and their spatial position are changed, which in the case of axial, radial, diagonal and straight vane grids of turbomachines leads to a reduction in vane lattice losses and to an improvement in step efficiency. Furthermore, depending on the choice of the influencing parameters such as leading edge radius, wedge angle and profile thickness, additional influence can be exerted on the blade load, that is to say the pressure distribution. By forming the thickening similar to a cosine or tangent function or by the dependence of the radial extent of the thickening on the thickness of the side wall boundary layer or The outflow angles on the trailing edge of the airfoil have a targeted effect on the position and size of the secondary vertebrae.

0013The side wall is understood to mean both the hub-side, that is to say the radially inner, and the housing-side, that is to say radially outer, boundary of the ring channel, which can be designed as a platform designed in the circumferential and axial direction or as a cover band or machine housing. The invention can inter alia be used in the stator and rotor blading of compressors, turbines, blowers and pumps, the thickening depending on the influence of the secondary flow being able to be carried out radially inside and / or outside.

0014In the case of simple blade geometries, for example straight turbine grids, a profile cut P is used as the reference profile cut_M to understand which at about 50% of the total blade height h_total can be found. In contrast, airfoils in axial turbomachines have more complex blade geometries since the individual profile cuts P that define the airfoils are adapted to the locally prevailing flow conditions. Because of the radially variable flow angle and conditions, the airfoil generally consists of a large number of individual profile cuts which are adapted to these conditions and change along the height of the airfoil and in particular in its orientation and position.

0015Aerodynamic, mechanical and thermal loads also lead to a change in the profile thickness and the axial dimensions over the height of the airfoil. For example, with guide vanes attached radially on the outside of the housing, the gas pressure load requires an increasing profile thickness and axial dimension in order to withstand the bending stress. In contrast, rotor blades in the hub area have the greatest profile thickness and axial dimension in order to be able to withstand the centrifugal stress. The thickening according to the invention at the edge zones is therefore to be understood as an additive thickening of a basic profile, which takes into account the aforementioned changes in the thickness of the airfoil. At the thickest point X of the airfoil, the thickening can be up to 30% of the chord length of the profile section P.

0016Further advantageous embodiments of the invention result from the features of claims 2 to 4, 6, 7 and 9 to 16.

0017By designing the invention by means of the features according to one of claims 12 or 13, the loss-reducing influence on the secondary flow is possible without inevitably modifying an already optimized profile pressure distribution in an undesired manner by the thickening. Here, the staggering angle b represents a design parameter which, after optimization of the secondary flow by means of the thickening according to the invention, has only a slight influence on the secondary flow, but the pressure distribution over the profile cuts P can still be optimized in the desired manner.

0018In order to compensate for an increase in weight of the airfoil due to the thickening, the airfoil can be made hollow in the region of the thickening. In addition, the strength properties of the airfoil can be improved by the thickening.

0019Preferred embodiments of the invention are explained below with reference to the accompanying drawings. It shows:<dl id="dl0001"><dt>Fig. 1</dt><dd>a longitudinal section through a low-pressure turbine of axial design with curved and straight blades,</dd><dt>Fig. 2a</dt><dd>2 shows a perspective view of a hub-side blade section according to section WW from FIG. 1 with an S-shaped thickening profile,</dd><dt>Fig. 2b</dt><dd>2 shows a fluid-mechanical model of the airfoil section according to FIG. 2a,</dd><dt>Fig. 2c</dt><dd>2 shows a fluid-mechanical model of the airfoil section according to FIG. 2a, showing the profile cuts P,</dd><dt>Fig. 2d</dt><dd>2 a shows a top view of the blade model according to FIG. 2 a with a reference and a profile section close to the edge,</dd><dt>Fig. 2e</dt><dd>2 shows a diagram of the course of the nose radius in the edge region of the airfoil according to FIG. 2a,</dd><dt>Fig. 2f</dt><dd>2 shows a diagram of the profile thickness profile in the edge region of the airfoil according to FIG. 2a,</dd><dt>Fig. 2g</dt><dd>2 shows a diagram of the course of the wedge angle g on the hub side at the front edge of the airfoil according to FIG. 2a,</dd><dt>Fig. 3a</dt><dd>2 shows a perspective view of a hub-side blade section according to section WW from FIG. 1 with an exponential thickening curve,</dd><dt>Fig. 3b</dt><dd>3 shows a fluid-mechanical model of the airfoil section according to FIG. 3a,</dd><dt>Fig. 3c</dt><dd>3 shows a diagram of the course of the nose radius in the edge region of the airfoil according to FIG. 3a,</dd><dt>Fig. 3d</dt><dd>3 shows a diagram of the course of the nose radius in the edge region of the airfoil according to FIG. 3a,</dd><dt>Fig. 3e</dt><dd>3 shows a diagram of the course of the wedge angle g at the front edge of the airfoil, FIG.</dd><dt>Fig. 4a</dt><dd>2 shows a perspective view of a hub-side Schuafelblattababschnitt according to section WW of FIG. 1 with threading the profile cuts in the circumferential direction,</dd><dt>Fig. 4b</dt><dd>4a shows a fluid-mechanical model of the airfoil section according to FIG. 4a,</dd><dt>Fig. 4c</dt><dd>4a shows a top view of the fluid mechanical model according to FIG.</dd><dt>Fig. 5a</dt><dd>View of a straight airfoil of a blade grille,</dd><dt>Fig. 5b</dt><dd>Profile section of the airfoil according to Fig. 5a,</dd><dt>Fig. 5c</dt><dd>Diagram of the course of the outflow angle α at the trailing edge of the airfoil according to FIG. 5a.</dd></dl>

0020An axial-circumferential-radial coordinate system zf -r, which is common in fluid mechanics, is used for direction and reference information. The upper half of a two-stage axial turbine 1 shown schematically in FIG. 1 has guide and rotor blade grids 2 and 3 arranged axially one after the other in pairs, which are equipped with straight or concave guide or rotor blades 4, 5 which are curved in a direction opposite to the direction of flow S. The blades 6 of the guide and rotor blades 4, 5 extend radially in an annular channel 7 arranged concentrically with the machine longitudinal axis A of the axial turbine 1. The axially concentric side walls of the hub and housing form the radially inner and radially outer channel delimiters 8 and 9 of the annular channel 7 and give it a divergent course with respect to the flow direction S.

0021The rotor blade grids 3 are designed in a disk design, ie the rotor blades 5 are each attached to a disk 10 in a grid manner.

00222a to 2c show the hub-side edge region of the blade blade 6 'belonging to the guide vane grille 2 in accordance with the section WW in FIG. 1. As is clear from the model according to FIGS. 2b and 2c, the blade blade 6' is formed in the area near the edge a variety of individual profile cuts P_R, P_1-7 and P_M generated. The shape of the profile sections P is customarily adapted to the radially variable flow conditions. Aerodynamic, mechanical and thermal loads have an influence on the design of the profile cuts P, which lead among other things to a change in the profile thickness d along the blade height h. Typically, the airfoil 6 ′ of the guide vane grille 2 is attached radially on the outside to the outer channel boundary 9, that is to say on the housing, and therefore generally has an increasing profile thickness d and profile depth t due to the gas pressure load. The rotor blade 5 belonging to the rotor blade grille 3, on the other hand, has the greatest profile thickness in the hub region in order to be able to withstand the centrifugal force stress. At the transitions of the blades 6 into the delimiting side walls 8, 9 - in the case of the blades 5, this is the case at the inner channel boundary 8 - they have a radius of curvature r_N or r_G on, which reduces the mechanical loads, such as the notch effect, so that the blades 6 each merge into the side walls 8,9.

0023As can be seen in FIGS. 2a to 2d and shown in the diagram in FIG. 2f in the qualitative course, the profile cuts P close to the edge have_R and P₁ - P₅ compared to a reference profile section P_M a thickening 11 increasing towards the outer channel boundary 9 in order to positively influence the secondary flow in the edge region. Here, the radial extent of the thickening 11 was chosen such that the thickening 11 lies in the region of the side wall boundary layer, which has a radial extent of just under<maths id="math0001" num=""><math display="inline"><mrow><msub><mrow><mtext>h / h</mtext></mrow><mrow><mtext>total</mtext></mrow></msub><mtext> = 12 %</mtext></mrow></math><img file="EP0798447A2_D0001.tif" /></maths> having. As a reference profile section P_M a profile cut P was chosen, which is 12% of the blade height h_total is spaced from the inner channel boundary 8 and has no thickening 11.

00242d shows the extent of the thickening 11 over the entire profile depth 1 compared to the thin reference profile section P._M clear. Starting from the leading edge 12 of the blade, the thickening 11 increases continuously up to a location X, designated as a thickness reserve, at which the maximum profile thickness is reached, and then continuously decreases as far as the trailing edge 13 of the blade. This increase and decrease in the thickening 11 takes place continuously in order to obtain an aerodynamically favorable shape. Due to the different graduation angle b_{R, M} of the two profile sections P_R and P_M fall the skeleton lines K_M and K_R of the two profile sections P_R and P_M not directly on top of each other, so that the thickening 11 is not directly visible on the pressure and suction side 14 or 15 in the trailing edge area.

0025As the diagrams in FIGS. 2e to 2g show, the thickening 11 results from the enlargement of the three profile parameters nose radius R._N on the blade leading edge 12, profile thickness d and wedge angle γ_N at the front edge 12. Starting from the reference profile cut P_M increases up to the profile section P near the edge_R the nose radius R_N by almost 300%, the profile thickness d by about 200% and the wedge angle γ_N by about 20 °. All three courses have in common that the thickening 11 starting from the profile section P near the edge_R up to the relative blade height <maths id="math0002" num=""><math display="inline"><mrow><msub><mrow><mtext>H</mtext></mrow><mrow><mtext>1</mtext></mrow></msub><msub><mrow><mtext>/H</mtext></mrow><mrow><mtext>total</mtext></mrow></msub><mtext> = 3%</mtext></mrow></math><img file="EP0798447A2_D0002.tif" /></maths> remains essentially constant, then within the height h₂, in a transition area T_Ü, which has an approximate tangent function with inflection point W at about 7% of the airfoil height h_total corresponds to running out to a value that corresponds to the reference profile section P_M corresponds. The height h_on with thickening 11 is thus from the height h₁ and h₂ together.

0026As can be seen in FIG. 2a, this results in a thickening 11 that extends up to an airfoil height <maths id="math0003" num=""><math display="inline"><mrow><msub><mrow><mtext>H</mtext></mrow><mrow><mtext>1</mtext></mrow></msub><msub><mrow><mtext> /H</mtext></mrow><mrow><mtext>total</mtext></mrow></msub><mtext>= 3%</mtext></mrow></math><img file="EP0798447A2_D0003.tif" /></maths> as a cylindrical section T_Z. is formed in order to then correspond to a transition function T_Ü to expire. In general, the thickening 11 is chosen such that b_M, b_R the profile cuts P different profile depths l_R or l_M but result in the dimension t in the axial direction of the thickening 11 unaffected. The dimension t is, however, for reasons of strength over the height of the airfoil h_total changeable.

0027An alternative course of the thickening 11 is shown in FIGS. 3a to 3f, the profile parameters determining the thickening 11 here starting from the profile section P near the edge_R steadily fall without turning point until the profile cuts P into the reference profile cut P_M pass over. The course of the parameters is approximated to a cosine function whereby the position and size of the secondary vertebrae is favorably influenced. The staggering and threading of the profile cuts P was left unchanged compared to the first variant according to FIG. 2.

0028A third embodiment of an airfoil design close to the edge is shown in FIGS. 4a to 4c. The further shaping of the airfoil 6 '' 'is defined by the threading of the profile cuts P in the circumferential direction. The airfoil 6 '' 'shown there has the same qualitative and quantitative thickening 11 and the same profile sections P as is expressed in the parameter profiles for the first exemplary embodiment according to the diagrams in FIGS. 2e to 2g. By threading the profile cuts P, after which the blade trailing edge 13 forms a straight line, the thickening 11 is shown in the top view according to FIG. 4b as a pressure-side thickening 11 of the airfoil 6 '' '. In this and in the other exemplary embodiments, the staggering angle b of the profile cuts P was chosen such that the influence of the thickening 11 on the suction-side pressure distribution to the profile cuts P is compensated, or at least the position of the pressure minimum relative to profile cuts P._M remains unchanged without thickening.

0029Further advantageous airfoil designs result from threading the edge cuts in the axial direction by using swept front edges, as shown in FIG. 1. The axial width of the profile cuts to the edges 8 or 9 can increase or decrease.

0030This exemplary embodiment shows that the thickening 11 can be carried out on the pressure and suction side, the threading of the profile cuts P in the axial direction and in the circumferential direction not making the location of the thickening 11 on the pressure or suction side directly recognizable in FIG. 4a.

0031While in the first embodiment, the height h_on of the area provided with the thickening 11 was selected as a function of the radial extent of the side wall boundary layer, the height h for the third example_on in dependence on the course of the outflow angle α at the trailing edge 13 of the airfoil. 5a to 5c show the relationships on an airfoil 6<i>''''</i> without thickening 11. In Fig. 5c is the course of the outflow angle <maths id="math0004" num=""><math display="inline"><mrow><msub><mrow><mtext>α = f</mtext></mrow><mrow><mtext>1</mtext></mrow></msub><mtext>(H)</mtext></mrow></math><img file="EP0798447A2_D0004.tif" /></maths> over the radial distance h from the side wall 8 for an airfoil without and f₂(h) shown for an airfoil 6 with thickening 11.

0032Characteristic of both courses is the area A of the over-deflection immediately following the side wall 8. This is followed by an area B of the less deflection, which at point C at the radial height h_C. changes into a constant course, which in turn is peculiar to the central region of an airfoil. The height h_C. defines the edge of the secondary flow areas in the wall areas of the blade and depends on the entry total pressure profile and the amount of deflection of the blade grille. The height h_on with the thickening 11 is now chosen such that about <maths id="math0005" num=""><math display="inline"><mrow><msub><mrow><mtext>H</mtext></mrow><mrow><mtext>on</mtext></mrow></msub><msub><mrow><mtext> = h</mtext></mrow><mrow><mtext>C.</mtext></mrow></msub></mrow></math><img file="EP0798447A2_D0005.tif" /></maths> applies. This leads to less reduction or over-redirection and to the desired reduction in secondary flow losses. The lower reduction and over-deflection also improves the inflow for the secondary grille, as a result of which the total loss of the step assembly is reduced.