JP2014066224A - Gas turbine blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a gas turbine blade capable of reducing pressure loss by reducing a separating region at a film cooling air inlet portion, reducing power of a cooling air supply source, and further improving film cooling efficiency.SOLUTION: A conical inlet portion is disposed at an inlet side of high pressure, of a film cooling hole. The conical shape is defined to have a bottom surface on a wall surface at a high pressure side, of a film cooling hole disposed, for example, on an aerofoil portion of a gas turbine blade, and an end wall portion at an outer peripheral side or an inner peripheral side of the aerofoil portion, to guide cooling air, and to guide the cooling air to a surface of the aerofoil portion and the like.

Description

  The present invention relates to a gas turbine blade, and more particularly to a gas turbine blade having a film cooling hole.

  The turbine blades of the gas turbine are exposed to hot gas. Therefore, in order to prevent high-temperature oxidation and thinning damage due to high-temperature gas, it is necessary to cool the airfoil portion, the inner peripheral side of the airfoil portion, and the end wall portion on the outer peripheral side. Therefore, the air extracted from the compressor is supplied to the cooling flow path formed inside the gas turbine blade so that the cooling flow path wall is convectively cooled, and a plurality of through holes (films) are formed on the surface of the gas turbine blade. The cooling of the gas turbine blade is suppressed by film cooling that sets a cooling hole) and jets air from the cooling flow path to the surface of the gas turbine blade and flows over the surface.

  The cooling air mixes with the mainstream gas and generates a mixing loss. This reduces the thermal efficiency of the turbine. Conventionally, in order to reduce cooling air, many ideas have been made to improve the cooling efficiency by devising the shape of the outlet side where the pressure of the film cooling hole is low.

  The shape on the inlet side where the pressure of the film cooling hole is high has not been studied so far. As what devised about the shape of the entrance side where the pressure of a film cooling hole is high, there exist some which were proposed by patent documents 1 and patent documents 2, for example. In Patent Document 1, the cooling air holes are configured to be perpendicular to the wall on the low temperature side, the cooling air hole density is made smaller on the low temperature side than on the high temperature side, and the convective cooling amount on the low temperature side is limited, It has been proposed to reduce the thermal stress in the wall material. Further, in Patent Document 2, a groove is provided in the blade inner wall in the blade height direction, a cooling hole opening from the bottom of the groove to the outer surface of the blade is provided, and the development of the boundary layer of the blade inner wall is reduced by the groove. It has been proposed to increase the heat transfer rate and increase the cooling efficiency. Moreover, as what devised the whole of the low side from the high pressure side of a film cooling hole, in patent document 3, air is made into a taper shape where a cross-sectional area reduces as it goes to a jet nozzle, and air is a film cooling hole. It has been proposed to reduce the mixing loss by reducing the speed difference from the mainstream by accelerating in the process of passing through.

JP-A-57-309 JP 59-113204 JP 2008-215233 A

  According to the study by the present inventors, it has been found that on the inlet side where the pressure of the film cooling hole is high, there is a separation region in the flow field due to non-uniformity of the flow from the cooling air plenum. This peeling region causes a pressure loss and causes an increase in the pressure difference between the inlet and outlet of the film cooling hole, which causes a reduction in film cooling efficiency. Moreover, it leads also to the power increase of the apparatus which supplies film cooling air.

  In the above-mentioned Patent Documents 1 to 3, no consideration is given to the separation region that occurs on the inlet side where the pressure of the film cooling hole is high, and even on the inlet side where the pressure of the film cooling hole is high even with the configuration of Patent Documents 1 to 3. The resulting peeled area cannot be reduced.

  An object of this invention is to provide the gas turbine blade which can make small the peeling area | region in a film cooling air inlet part.

  The present invention is characterized in that a conical inlet portion is provided on the inlet side where the pressure of the film cooling hole is high.

  According to this invention, the peeling area | region in a film cooling air inlet part can be made small. As a result, the pressure loss can be reduced, the power of the cooling air supply source can be reduced, and the film cooling efficiency can be improved.

  Problems, configurations, and effects other than those described above will be clarified by the following description of embodiments.

1 is a cross-sectional view showing a schematic configuration of a gas turbine to which the present invention is applied. The figure explaining an example of the cooling structure in a gas turbine moving blade. The figure which shows the cross-sectional shape of the airfoil part of a gas turn blade. Film cooling hole sectional drawing explaining the function of a film cooling hole. The figure explaining the peeling phenomenon which generate | occur | produces inside a film cooling hole. The film cooling hole sectional drawing in the Example of this invention. Explanatory drawing which looked at the film cooling hole in the Example of this invention from the cooling air pressurization side. It is sectional drawing of the film cooling hole in the other Example of this invention. Explanatory drawing which looked at the film cooling hole in the other Example of this invention from the cooling air pressurization side. Explanatory drawing which applied the Example of this invention to the film cooling hole of a wing | blade surface. The figure explaining the reduction effect of the compressor power in the Example of this invention.

  Hereinafter, embodiments of the present invention will be described in detail with reference to the drawings.

  First, a schematic configuration of a gas turbine to which the present invention is applied will be described with reference to FIG.

  A gas turbine is roughly divided into a compressor, a combustor, and a turbine. The compressor 5 compresses air sucked from the atmosphere as a working fluid, and the combustor 6 mixes fuel with the compressed air supplied from the compressor 5 and burns to generate high-temperature and high-pressure gas. The high-temperature and high-pressure gas is introduced into the turbine composed of the stationary blade 8 and the moving blade 4. In general, the moving blades 4 are alternately arranged together with the stationary blades 8 and are implanted in grooves provided on the outer peripheral side of the wheel portion of the rotor 1. The stationary blade 8 is supported by the casing 7. The high-temperature and high-pressure gas is blown to the moving blade 4 through the stationary blade 8 and drives the rotor 1 through the moving blade 4. The rotor 1 also includes moving blades of the compressor 5, and the compressor 5 is driven by the turbine via the rotor 1.

  Gas turbines tend to be heated to increase efficiency, and the surface temperature of gas turbine blades exposed to high-temperature combustion gas exceeds the limit temperature of the heat-resistant alloy used. Cooling of the stationary blade 8) is required. An example of the blade cooling structure will be described with reference to FIGS.

  As shown in FIG. 2, the moving blade 4 includes an inner peripheral end wall portion located on the turbine rotor side, and an airfoil portion 12 extending in a direction in which the radial position increases from the inner surface 10 of the inner peripheral end wall portion. Have. A gap 17 through which a fluid flows is formed between the blade tip where the radius of the airfoil portion 12 becomes the largest and the outer peripheral end wall portion outer surface 16 of the stationary blade forming the gas flow path surface. The airfoil portion 12 has a hollow portion, and is configured to cool the blade from the inside by flowing a cooling medium in the hollow portion. The compressor 5 is often used as a cooling medium (cooling air) supply source. Cooling air extracted from an intermediate stage or an outlet of the compressor 5 is introduced into the moving blade 4 using a cooling air introduction hole provided in the rotor 1. In FIG. 2, the cooling air is introduced into the airfoil 12 from the cooling air inlet 9, and the cooling air flows in the direction of the arrow to cool the airfoil. In FIG. 2, pin fins 9c and turbulent flow promoting ribs 9d are provided inside the blades to improve the convection cooling performance. The cooling air after cooling the inside of the blade is discharged from the discharge hole 15 provided in the inner peripheral wall, and is eventually discharged to the gas path. Further, the rotor 1 is also cooled, and the air that has cooled the rotor 1 flows into the gas path path as indicated by arrows 18 and 19 from the gap between the stationary vanes 13 and 14 and the moving blades 4 on the inner peripheral side.

  FIG. 3 shows a cross-sectional shape of the airfoil portion 12. The airfoil portion includes a pressure surface 10b (blade ventral side portion) that is concave in the chord direction, a negative pressure surface 10a (blade back side portion) that is convex in the chord direction, a blade leading edge 12a, The blade edge 12b is formed such that the blade thickness gradually increases from the front edge side toward the center side and gradually decreases from the middle toward the rear edge side. A hollow portion constituting an air cooling chamber is provided inside the airfoil portion, and the blade is cooled from the inside by flowing a cooling medium in the hollow portion. In FIG. 3, insert members 9a and 9b each having an impingement cooling hole are provided for impingement cooling the blade inner surface. Although not shown, the cooled cooling air is discharged from the discharge hole provided in the inner peripheral wall, and is eventually discharged to the gas path.

  Various cooling means such as convection cooling and other cooling means are used for the cooling structure of the turbine blade, but the present invention relates to the shape of the pressure side of the cooling hole for supplying the cooling air after cooling inside the blade to the gas path through which the mainstream gas flows. is there. That is, the present invention relates to the shape of the pressure side (inside of the blade) of the film cooling hole for covering the blade surface with the cooling air ejected from the inside of the blade to the outside of the blade and cooling the film.

  FIG. 4 is a conceptual diagram of the film cooling hole 20 installed in the blade metal part 30 of the turbine blade. The film cooling hole 20 is divided into a pressure side having a high pressure and a mainstream gas side having a low pressure. At this time, the cooling air passes through the cooling holes and mixes with the mainstream gas as indicated by arrows in FIG. The film cooling holes are set so that the cooling film is effectively formed, that is, the angle of the film cooling holes makes an acute angle with the wall so that the cooling air follows the metal part of the blade as much as possible. Yes. In addition, the film cooling holes are formed in an elliptical shape or the like with the intention of forming a cooling air layer widely on the surface of the gas turbine blade (see FIG. 7). Such a film cooling hole is provided in a portion to which a heat load is applied, and reduces the inflow of heat that moves from a high-temperature gas to an airfoil portion or the like by causing cooling air to flow over the surface of the portion to which the heat load is applied. Therefore, the present invention can be applied to the film cooling holes formed in any one of these blades and stationary blades, and the inner or outer end wall portions.

  FIG. 5 is a diagram showing a problem inside the film cooling hole as shown in FIG. As shown in FIG. 5, the flow field peels off on the pressure side of the film cooling hole guided by the cooling air (peeling region 40). This is because the cross-sectional area changes rapidly in the path through which the cooling air is guided. In particular, peeling occurs on the side of the cooling hole that forms an acute angle. Due to the presence of this peeling region, the pressure loss at the film cooling hole increases. Further, the cooling air passing through the separation region becomes a fluctuating flow field after passing through the separation region, and when guided to the mainstream gas side, mixing with the mainstream gas is promoted. Thereby, the temperature of film cooling air rises and it leads to a film efficiency fall.

  Further, the fluctuating film air generated inside the film cooling holes in this way causes a mixing loss when mixed with the mainstream gas, and reduces the energy of the mainstream gas, leading to a decrease in turbine efficiency.

  In the present invention, a conical inlet portion is provided in order to reduce the change in the cross-sectional area of the channel for guiding the film air on the inlet side where the pressure of the film cooling hole increases, and the cooling air is guided by this inlet portion shape. The flow passage cross-sectional area change is moderated, and the peeling region at the film cooling hole inlet is suppressed.

  The pressure-side film cooling hole shape in the embodiment of the present invention will be described with reference to FIG.

  In the present embodiment, the film cooling hole shape on the pressure side is made conical to reduce the peeling area. When the film air flows by using the conical shape, the rapid change of the flow field is suppressed by reducing the change in the cross-sectional area to the film cooling hole (main hole). In other words, in this embodiment, the film cooling hole is formed by a main hole on the mainstream gas side and a conical inlet portion on the pressurization side, and the flow field is sharpened by reducing the cross-sectional area change toward the mainstream gas side. Slow speed increase is suppressed. Thereby, a rapid pressure change at the inlet wall surface of the film cooling hole is suppressed. As a result, separation of the flow field at the film cooling hole inlet is suppressed, and rectified cooling air can be guided to the film cooling hole as compared with the case where the present embodiment is not applied.

  Since the rectified cooling air is guided to the film cooling holes, fluctuations in the flow field inside the film cooling holes are suppressed, and the temperature rise caused by mixing the film air with the mainstream gas is reduced, and the film efficiency is reduced. It leads to.

  Further, since the rectified cooling air is supplied to the mainstream gas, the mixing loss of the mainstream gas is reduced. Further, the reduction of the peeled area reduces the total pressure loss of the film cooling holes, reduces the differential pressure inside the film cooling holes, and reduces the Mach number. By this effect, the heat transfer coefficient at the wall surface is reduced, and the heat flux from the mainstream gas to the blade metal part is also suppressed.

  This conical inlet portion can be installed as a post-processing of a cylindrical film cooling hole (main hole).

  FIG. 7 shows the positional relationship between the film cooling hole (main hole) and the conical inlet in the embodiment of FIG. A conical inlet portion 50 of the film cooling hole processed into a conical shape is indicated by a broken line. In the embodiment of FIG. 6, the apex of the cone exists in a region where the angle formed by the metal portion on the pressing side and the film cooling hole (main hole) is an acute angle. Due to the position of the apex, the area where the cooling air is peeled off after the conical inlet is post-processed is reduced, and the film efficiency and the turbine efficiency are improved.

  In this embodiment, the film cooling hole (main hole) has an elliptical shape, the diameter of the bottom surface of the conical inlet portion is larger than the short axis of the elliptical film cooling hole, and the conical inlet portion. Is provided so as to cover the acute angle side of the film cooling hole. By configuring the conical inlet portion in this way, it is possible to effectively suppress peeling that is particularly likely to occur on the acute angle side of the film cooling hole.

  Since the peeling area generated at the entrance of the film cooling hole installed in the turbine blade can be reduced, the fluctuation of the cooling air generated by the peeling area can be suppressed. Moreover, since the reduction of the peeling area leads to the pressure loss reduction of the film cooling hole, the differential pressure for supplying the film air can be reduced. Since the pressure loss is reduced, it is possible to reduce the power of the apparatus that supplies the cooling air. In a normal gas turbine facility, cooling air is supplied by extracting air from a compressor. By applying the present invention, the compressed air can be extracted from the stage close to the inlet of the compressor and used as cooling air, leading to an improvement in compressor efficiency.

  Next, another embodiment of the present invention will be described with reference to FIGS.

  In the present invention, the pressure-side film cooling hole shape is conical to reduce the peeling region. In this embodiment, as shown in FIGS. 8 and 9, an elliptical film cooling hole shape is used. The bottom surface of the conical inlet 50 is positioned so as to cover the entire area 20 (main hole). In this embodiment, the apex of the cone is located on the film cooling hole.

  Also in this embodiment, as in the previous embodiment, it is possible to reduce the peeling area generated at the entrance of the film cooling hole. In particular, in this embodiment, the conical inlet region extends over the entire pressure-side region of the film cooling hole (main hole), and the peeling region can be effectively reduced.

  The apex angle of the cone is desirably set to 120 ° to 150 ° also in the embodiments of FIGS. 6 and 7 and the embodiments of FIGS. In many cases, the film cooling hole is set at around 30 ° with respect to the blade surface that allows cooling air to flow. In this case, it is possible to effectively reduce the separation region by forming the conical inlet portion so that the apex angle of the cone is 120 ° to 150 °.

  FIG. 10 shows a gas turbine blade to which the film cooling hole of the above-described embodiment is applied. In FIG. 10, the above-described embodiment is applied to the film cooling hole 20 provided on the blade surface. Although the film cooling hole 20 provided in the suction surface 10a (blade back side portion) is described in FIG. 10, it can be similarly applied to the film cooling hole provided in the pressure surface 10b (blade bell side portion).

  In this embodiment, a conical inlet is set in the film cooling hole inside the blade. As in this embodiment, when the conical inlet is set in the film cooling hole inside the blade, it can be manufactured by precision casting. By suppressing the pressure-side separation region inside the blade, flow field rectification inside the film cooling hole and total pressure loss reduction can be achieved. This effect contributes to the reduction of compressor power and contributes to the efficiency improvement of the entire gas turbine facility.

  Further, in the above-described embodiment, the case where it is applied to the airfoil portion of the gas turbine blade has been described, but it can be similarly applied to the film cooling hole formed in the outer peripheral side or the inner peripheral side end wall portion. An effect is obtained.

  The compressor power reduction effect when the present invention is applied will be described with reference to FIG. Let A be the pressure on the pressure side when the present invention is not applied. By applying the present invention, the total pressure loss inside the film cooling hole is reduced, so the pressure difference between the inlet and outlet of the film cooling hole (the difference between the pressure on the mainstream gas side and the pressure on the pressure side of the film cooling hole) Becomes smaller. Due to this effect, it is possible to provide the mainstream gas side with cooling air of the same degree as before application of the present invention with a small differential pressure. The pressure when the present invention is applied may be smaller than A as in B. Therefore, the cooling air can be extracted from the side having the lower number of stages of the compressor, thereby improving the compressor efficiency.

  In addition, this invention is not limited to an above-described Example, Various modifications are included. For example, the above-described embodiments have been described in detail for easy understanding of the present invention, and are not necessarily limited to those having all the configurations described. Further, a part of the configuration of one embodiment can be replaced with the configuration of another embodiment, and the configuration of another embodiment can be added to the configuration of one embodiment. Moreover, it is possible to add, delete, and replace other configurations for a part of the configuration of each embodiment.

  DESCRIPTION OF SYMBOLS 1 ... Rotor, 3 ... Rotating shaft, 4 ... Moving blade, 5 ... Compressor, 6 ... Combustor, 7 ... Casing, 8 ... Stator blade, 9 ... Cooling air inlet, 12 ... Airfoil part, 30 ... Blade metal part 20 ... film cooling holes, 40 ... peeling area, 50 ... conical inlet.

Claims (5)

A gas turbine blade used in a gas turbine,
An airfoil with a cooling channel formed therein;
The airfoil portion is formed with a cooling hole for ejecting cooling air from the inside to the outside,
The gas turbine blade according to claim 1, wherein a conical inlet portion having a surface inside the airfoil portion as a bottom surface is formed inside the airfoil portion of the cooling hole. In claim 1,
The gas turbine blade according to claim 1, wherein the cooling hole is formed in an elliptical shape, and a diameter of the bottom surface of the conical shape is larger than a short axis of the elliptical cooling hole. In claim 2,
The gas turbine blade according to claim 1, wherein the conical inlet portion has a bottom surface that covers the entire surface of the elliptical cooling hole. A gas turbine blade used in a gas turbine,
An airfoil having a cooling channel formed therein;
An inner peripheral end wall located on the inner peripheral side of the airfoil, and
Comprising an outer peripheral end wall located on the outer peripheral side of the airfoil,
Cooling air is guided to the airfoil part, the inner peripheral side end wall part or the outer peripheral side end wall part, and the airfoil part, the inner peripheral side end wall part or the outer peripheral side end wall part airfoil part A cooling air hole is formed on the surface of the cooling air so as to guide the cooling air;
The gas turbine blade according to claim 1, wherein a conical inlet portion is formed on a surface of the cooling hole where the cooling air inlet is formed. In any one of Claims 1-4,
A gas turbine blade characterized in that an apex angle of the conical cone is set to be 120 ° to 150 °.

Images