JPH1068304A - Method and apparatus for reducing turbine blade pressure loss - Google Patents

Abstract

(57) [Abstract] (with correction) [PROBLEMS] To reduce the pressure loss of a transonic turbine blade where the flow velocity of the mainstream gas at the cascade outlet becomes transonic or supersonic, without causing an increase in weight. The present invention relates to a turbine blade pressure loss reducing method effective in a wide turbine operating range and an apparatus therefor. SOLUTION: An ejection gas is supplied to the inside of a turbine blade 10, and the ejection gas and the main gas flowing from a trailing edge 20 pressure surface of the turbine blade along a turbine blade pressure surface 14 constituting a cascade are formed. Injection is performed by reducing the mass flow velocity ratio determined by the density and velocity with the mainstream gas to 1 or less, to reduce the generation of shock waves from the pressure surface at the trailing edge, and to the suction surface of the turbine blade adjacent to the turbine blade generating the shock wave. The shock wave was not injected into the developing boundary layer, and the enlargement of the boundary layer due to the disturbance of the incident shock wave was suppressed to reduce the pressure loss. As a result, the pressure loss of the turbine blade can be reduced and the thermal efficiency of the turbine can be improved.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention is applied to a turbine blade forming a turbine cascade in a gas turbine or the like, and is provided with a shock wave generated from a pressure surface at a trailing edge of the turbine blade and an adjacent shock blade. Turbine blade pressure loss reduction method that reduces pressure loss caused by interference with the boundary layer developed on the suction surface of installed turbine blades and contributes to improvement of thermal efficiency of turbine, and turbine blade pressure reduction for implementing the method Related to the device.

[0002]

2. Description of the Related Art FIG. 7 shows a turbine blade cascade formed by a conventional gas turbine or the like, and one of the turbine blades that generates power by interacting with a mainstream gas flowing into the turbine cascade. FIG. 3 shows a perspective view of a cooled turbine blade. However,
The non-cooling refers to a cooling blade having no cooling structure in which a cooling passage or the like for flowing a cooling medium such as compressed air is provided inside the turbine blade and the inside of the blade is cooled.

The turbine blade 010 shown in the figure is a nozzle (not shown) provided on the upstream side, and its flow direction is deflected and accelerated by expansion between stages, and the flow direction of the main flow gas such as combustion gas flowing into the turbine blade 010 is shown. The pressure surface 014 which is the ventral side facing the air, the negative pressure surface 016 which is the back side formed on the back of the pressure surface 014, the leading edge 018 which is the most upstream end with respect to the inflowing mainstream gas, and the most downstream end A trailing edge 020 is provided, and includes a streamlined airfoil portion 012 and a dovetail-shaped blade root portion 022 for attaching the airfoil portion 012 to a turbine disk (not shown).

[0004] The thus-formed turbine blade 010
As described above, are planted at equal pitches in the outer circumferential direction of the turbine disk to form an annular turbine cascade,
Power is generated by interaction with the mainstream gas that is ejected from the nozzle and flows in, and is output to the outside by rotation of the turbine rotor.

[0005] Among such turbine blades 010, the relative velocity of the mainstream gas with respect to the turbine blades 010 increases depending on the peripheral speed of the turbine blades 010 or the absolute speed and the ejection angle of the mainstream gas ejected from the nozzle. ,
In a so-called transonic turbine rotor blade in which the Mach number at the turbine blade cascade exit is very close to 1.0 or a supersonic flow exceeding 1.0, the mainstream gas flowing along the pressure surface 014 is a turbine. Negative pressure surface 01 on trailing edge 020 side of rotor blade 010
6, the expansion wave is generated, and due to the overexpansion due to the expansion wave, the pressure surface 014 close to the trailing edge 020 of the turbine blade 010, which is the main flow gas wraparound portion. , A shock wave is generated, and is incident on a boundary layer which is disposed adjacent to the turbine moving blade 010 having the pressure surface 014 for generating the shock wave and which develops the negative pressure surface 016 of the turbine moving blade 010.

In this manner, the disturbance of the shock wave generated at the trailing edge pressure surface 014 and incident on the boundary layer of the suction surface 016 of the adjacent turbine blade 010 flows through the subsonic flow in the boundary layer and reaches the upstream side. To the side, and at the same time, change the thickness of the boundary layer, the amount of separation, transition, etc.
A so-called interference with the boundary layer is caused, the boundary layer of the suction surface 016 behind the interference position is enlarged, the interval between the turbine cascades is substantially reduced, and the flow rate of the mainstream gas flowing through the turbine cascade is reduced. By increasing the pressure loss of the turbine blades 010, not only by reducing it, but also by changing the flow outside the boundary layer.

[0007] In order to reduce the pressure loss of the transonic turbine blade, (1) the Mach number at the turbine cascade outlet is suppressed to a subsonic level, and the main gas flows into the trailing edge negative pressure surface 016. Thereby, the shock wave generated from the pressure surface 014 at the trailing edge of the turbine blade 010 is weakened. (2) The shape of the suction surface 016 of the turbine blade 010 is devised so that the shock wave generated from the pressure surface 014 at the trailing edge of the turbine blade 010 and the boundary layer developed on the suction surface of the adjacent turbine blade 010. The shape of the negative pressure surface 016 is such that the enlargement of the boundary layer behind the interference position due to the interference can be suppressed. Such methods have been conventionally considered.

In the method (1), the number of turbine stages is increased, the expansion ratio per stage, and the load are reduced, so that the absolute velocity of the mainstream gas and the turbine blades are reduced.
Since the peripheral speed of the turbine blade 10 can be reduced and the Mach number at the turbine cascade outlet can be set to a subsonic level, it can be effective for a turbine having no weight limitation.

However, in the case of ordinary turbines, if they have the same performance, lower cost and downsizing are more strongly required, and the pressure loss of the turbine blade is reduced by the method (1). Is not a very effective method. In particular, in an aeronautical gas turbine, the light weight is the first proposition. Therefore, in order to reduce the weight, reduction of the number of turbine stages is always pursued, and the expansion ratio per stage and the load are inevitably high. The turbine cascade exit Mach number is usually transonic or supersonic,
The method (1) cannot be used as an effective method for reducing the pressure loss of the turbine blade.

The method (2) is an effective method,
Alternatively, a device for implementing the method has not been reported so far, and is a method for which further research and development are awaited. However, in gas turbines that operate by changing the load frequently, such as in aviation gas turbines, their practical application will be extremely difficult.

That is, like a gas turbine for aviation,
In a gas turbine in which the load fluctuates, the peripheral speed of the turbine blade 010 changes due to a change in the operating state corresponding to the load, so that the relative speed of the mainstream gas with respect to the turbine blade 010 changes, and a shock wave is generated. Since the position fluctuates, the interference position between the boundary layer of the suction surface 016 and the shock wave moves. For this reason, according to the change of the operating state,
This is because, unless the shape of the negative pressure surface 016 is changed, for example, even if the method is effective for reducing pressure loss at full load, it is not effective at partial load.

[0012]

SUMMARY OF THE INVENTION The present invention is directed to a so-called transonic turbine blade, in which the Mach number at the turbine blade cascade outlet becomes transonic or supersonic, in particular, the trailing edge of the so-called transonic turbine blade (pressure surface). In order to reduce the pressure loss due to the interference between the shock wave generated by the shock wave and the boundary layer developed on the back side (vacuum side) of the turbine blade adjacent to the turbine blade that generated the shock wave, the number of turbine stages was increased and the expansion per stage was increased. Without lowering the ratio and load, the flow of mainstream gas around the trailing edge of the turbine blade weakens the strength of the shock wave generated from the trailing edge pressure surface and changes the shape of the suction surface of the turbine blade. A turbine blade pressure loss reduction method that suppresses the enlargement of the boundary layer that develops on the suction surface without
An object of the present invention is to provide a turbine blade pressure loss reduction device for performing the method.

[0013]

For this reason, the method for reducing turbine blade pressure loss according to the present invention employs the following means. (1) Discharge gas, such as compressor bleed gas or combustion gas, supplied to the inside of a turbine blade, which is arranged in a circumferential direction to form a turbine cascade in which blades are arranged at an equal pitch. Out of the nozzle, or out of the preceding turbine rotor blade, flow into the turbine cascade, reduce the mass flow rate ratio to 1 or less, and eject it from the trailing edge pressure surface of the turbine blade into the mainstream gas. , Which is generated from the pressure surface of the trailing edge of the turbine blade and impinges on the suction surface of the adjacent turbine blade, reduces the intensity of the shock wave and reduces the interference with the boundary layer developed on the suction surface of the adjacent turbine blade. Thus, the enlargement of the boundary layer is suppressed, and the pressure loss of the turbine blade is reduced. The mass flow rate ratio is defined by Equation 1.

[0014]

(Equation 1)

Here, ρ indicates the density, V indicates the flow velocity, the subscript e indicates the gas for ejection, and the subscript a indicates the mainstream gas.

It is effective to use a gas having a large density ρ for the gas for ejection, and the speed of ejection into the mainstream gas also depends on the density ρ of the gas for ejection used. In the case where the gas is used as a gas for ejection, a gas having a mass flow rate ratio of about 0.2 is preferable. Further, a turbine blade refers to both a turbine blade and a turbine stationary blade forming a nozzle.

The method for reducing turbine blade pressure loss according to the present invention comprises:
By employing the above-described means, (a) at the trailing edge of the turbine blade, the jetting gas ejected from the trailing edge pressure surface of the turbine blade into the mainstream gas flows from the trailing edge pressure surface side of the turbine blade to the negative pressure surface arc portion. In order to prevent the flow of the mainstream gas from flowing to the side, the generation of an expansion wave generated by the mainstream gas flowing around is suppressed, and the overexpansion is weakened.
The intensity of the shock wave generated from the pressure surface of the trailing edge of the turbine blade in the wraparound portion is reduced or no longer generated.
It is possible to at least reduce the pressure loss generated due to the interference between the shock wave and the boundary layer on which the shock wave is incident and develops on the suction surface of the adjacent turbine blade. Thereby, the thermal efficiency can be improved, and an efficient turbine can be obtained.

Further, the turbine blade pressure loss reducing device of the present invention employs the following means in order to perform the turbine blade pressure loss reducing method described in (1) above. (2) compressor bleed gas supplied to the blade root end of the turbine blade via a passage provided in a turbine rotor or the like;
Alternatively, a fluid supply port for introducing jetting gas having a high density ρ, which is made of a combustion gas and preferably can increase the mass flow rate at the time of jetting, into the turbine blade is formed in the blade root. (3) A blade internal passage that evenly distributes the jetting gas introduced into the turbine blade through the liquid supply port in the height direction of the airfoil portion that interacts with the mainstream gas of the turbine blade is formed as a fluid. The inside of the airfoil was defined by communicating with the supply port. (4) An opening is arranged in the height direction of the airfoil portion on the pressure surface of the trailing edge portion of the airfoil portion, and the ejection gas supplied from the air passage inside the airfoil communicates with the mass defined by Formula 1 described above. The flow rate ratio was set to 1 or less, and a jet port for jetting the mainstream gas was provided. It is more preferable that the jet outlet is a slit-shaped opening communicating with the airfoil portion in the height direction.

The fluid supply port formed in the blade root portion and the blade internal passage defined in the airfoil portion are formed by a turbine blade having cooling passages therein, and a refrigerant gas passing through these cooling passages. With a cooling structure that cools from the inside,
In the case where the cooling turbine blades are used, the cooling passages through which these refrigerants pass can also be used. However, in such a case, refrigerant gas is used as the gas for ejection.

Further, the jet port for jetting the jetting gas into the mainstream gas from the pressure surface at the trailing edge of the airfoil portion is a cooling turbine blade having a cooling structure as described above. In the case of having a cooling device that performs impingement cooling on the trailing edge portion of the turbine blade, the cooling device may be used also as these cooling devices, or supplied from these cooling structures to the cooling device. A part of the refrigerant may be branched, introduced, and injected.

In the turbine blade pressure reducing device of the present invention, by means of (2) to (4) described above, (b) openings arranged at the height of the airfoil portion are formed on the turbine blade trailing edge pressure surface. The jet gas having a mass flow rate ratio smaller than 1 is injected into the mainstream gas flowing along the pressure surface from the ejection port having the gas at the trailing edge of the turbine blade from the pressure surface side of the turbine blade trailing edge at the trailing edge of the turbine blade. In order to prevent the mainstream gas from sneaking into the arc on the negative pressure side, the sneakage of the mainstream gas suppresses the generation of expansion waves generated on the pressure surface at the trailing edge of the turbine blade, and can reduce overexpansion. Thus, the shock wave generated from the pressure surface at the trailing edge of the turbine blade can be reduced, and the pressure loss due to the interference between the shock wave and the boundary layer developed on the suction surface of the adjacent turbine blade can be reduced.

As a result, the effect of improving the fuel efficiency of the turbine can be obtained, and a turbine having high thermal efficiency can be obtained. When the turbine blades have a cooling structure or an impingement cooling device, the cooling passages and cooling devices provided in these turbine blades are provided with a liquid supply port, a blade internal passage,
As described above, it can also be used for the jet port. Further, when the turbine blade has no cooling structure and does not need to be cooled, the trailing edge is not limited by the cooling design like the cooling turbine blade. It is also possible to freely set the state quantity of the ejection gas ejected from the partial pressure surface, and to select the flow rate of the ejection gas such that the pressure loss is minimized with respect to the operating state of the gas turbine. Thereby, even in the case of a partial load, it is possible to effectively reduce pressure loss remarkably.

[0023]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of a turbine blade pressure loss reducing device according to the present invention will be described below with reference to the drawings. FIG. 1 is a perspective view showing a first embodiment of a turbine blade pressure loss reducing device according to the present invention.

As shown in the figure, a gas turbine moving blade 10 as a turbine blade in which the present embodiment is provided is arranged in a flow direction of a combustion gas as a mainstream gas flowing from a nozzle provided on an upstream side (not shown). An airfoil portion 12 provided with a facing pressure surface 14, a suction surface 16 formed on the back surface of the pressure surface 14, a leading edge 18 at the most upstream end, and a trailing edge 20 at the most downstream end; Airfoil 1 on disk
2 for fixing the blade 2. In the gas turbine rotor blade 10 according to the present embodiment, the dovetail-shaped blade root portion 22 is required to attach the airfoil portion 12 to the turbine disk.
However, it is also acceptable to use a blisk (not shown) in which the airfoil portion 12 and the turbine disk are integrated.

A pressure surface 14 on the trailing edge 20 side of the gas turbine blade 10 (hereinafter referred to as a trailing edge pressure surface) has a jet opening opening in a rectangular slit shape continuous in the height direction of the airfoil portion 12. 24 are provided. From this spout 24
An ejection gas F such as a compressor bleed gas or a combustion gas is ejected in the direction of the arrow into the mainstream gas flowing along the pressure surface 14.

The jet port 24 shown in the figure has one rectangular slit-shaped opening continuous in the height direction of the airfoil portion, and jets the jetting gas F. 5 is a perspective view of a rectangular slit divided into a plurality of portions in the height direction of the portion 12 or a second embodiment of the turbine blade pressure loss reducing device according to the present invention. Alternatively, an opening such as an oblong hole may be used. When the injection port 24 is divided into a plurality in the height direction of the airfoil portion 12 by the girders as described above, the thickness of the gas turbine blade 10 must be reduced in view of the structural strength, particularly, the operation efficiency. And the structural strength of the trailing edge portion can be increased, and the airfoil portion 12 having the ejection speed of the ejection gas F can be increased.
Distribution in the height direction has an effect of being easily controlled, but in a region where the opening of the ejection port 24 is not provided,
There is also a problem that the effect obtained by injecting the ejection gas F into the mainstream gas described below is reduced.

Next, FIG. 2 is a cross-sectional view taken along the line AA shown in FIG. 1, FIG. 3 is a detailed view of a portion A in FIG. 2, and FIG. 4 is a vertical sectional view taken along the line BB shown in FIG. It is. As shown in these figures, compressor bleed air or combustion gas supplied to the airfoil portion 12 from a fluid supply port 32 formed in the blade root portion 22 through a gas passage (not shown) provided in the turbine disk. The jetting gas F is formed from the jet port 24 through the blade inner passage 30 defined in the blade height direction in the airfoil portion 12, and is uniformly distributed in the blade height direction. It is blown out to the mainstream gas flowing along.

It should be noted that the fluid supply port 32 for introducing the jetting gas F into the gas turbine blade 10 is shown to be formed only at one location of the blade root portion 22. A plurality of ports 32 may be provided in the blade root portion 22 or, in the case of using a blisk, in a turbine disk.

Further, the blade internal passage 30 is not simply defined in the blade height direction as shown in FIG. 4, but is formed in a complicated shape such as a return flow, a pin fin or the like which is found in a cooling turbine blade. But it doesn't matter. Further, the cooling structure such as a cooling passage provided in the cooling turbine blade is replaced with the fluid supply port 32 or the blade internal passage 3.
The cooling device can also be used as 0, and the cooling device that cools the trailing edge by impingement cooling with the cooling turbine blades can also be used as the injection port 24 in the cooling device. It is. However, in such a case, the refrigerant gas for cooling the cooling turbine blades is used as the jetting gas, and the injection speed of the jetting gas F into the mainstream gas is restricted. In some cases, the effect of reducing the turbine blade pressure may be reduced.

As shown in FIG. 3, the injection port 24 for injecting fluid into the mainstream gas has a shape along the curved surface of the suction surface 16 (hereinafter referred to as the blade trailing edge arc). The jetting gas F jetted from the nozzle is jetted into the mainstream gas along the arc of the blade trailing edge.

The turbine blade pressure reducing device of the present embodiment has the above-described configuration, and discharges gas F such as compressor bleed gas and combustion gas from the jet port 24 provided on the trailing edge pressure surface 14.
Is set to an appropriate flow velocity, that is, the mass flow velocity ratio defined by Equation 1 is set to 1 or less, preferably about 0.2, and is injected into the main flow gas, so that the main flow gas wraps around the wing trailing edge arc portion. Interaction between a shock wave generated from the pressure surface 14 of the trailing edge of the turbine blade 10 and a boundary layer developed on the suction surface 16 of the adjacent turbine blade 10 due to suppression of generation of the generated expansion wave and overexpansion of the mainstream gas. Pressure loss is reduced.

Thus, the pressure loss of the high-load, high-expansion-ratio turbine generated when the flow velocity of the mainstream gas at the outlet of the gas turbine blade 10 becomes transonic or supersonic can be reduced, and the fuel efficiency of the gas turbine and the like can be reduced. The effect of improving can be obtained.

When the gas turbine blade 10 has a cooling structure or an impingement cooling device, the cooling passages and cooling devices provided in these gas turbine blades are connected to the liquid supply port 32, the blade internal passage 30, and the jet port 24. Is also available,
The structure is simplified and the cost can be reduced.

Conversely, when the gas turbine blades need not be cooled without having a cooling structure, the jets blown out from the trailing edge pressure surface 14 are not limited by the cooling design as in the case of the cooling turbine blades. The amount of state of the gas F for use can be set freely, and it is possible to inject the gas F for ejection so as to minimize the pressure loss with respect to the operating state of the gas turbine. Even in the case of load,
It can be effective for remarkable pressure loss reduction of the turbine blade.

Next, FIG. 6 is an enlarged cross-sectional view of a trailing edge fluid ejection port, showing a third embodiment of the turbine blade pressure loss reducing device of the present invention. The ejection port 24 ′ of the present embodiment has a shape along the curved surface of the pressure surface 14, and the ejection port 24 ′
The jetting gas F jetted into the mainstream gas is jetted along the curved surface of the pressure surface 14 of the gas turbine blade 10.

According to the present embodiment, as can be seen from the comparison with FIG. 3 showing the first embodiment, the shape of the trailing edge 20 maintains the original airfoil shape of the trailing edge. For,
Compared with the first embodiment, there is an advantage that the outlet outflow angle can be easily controlled. On the other hand, when the optimal ejection flow velocity is too high, the ejection gas F does not follow the curved surface of the pressure surface 14 and directly flows into the mainstream. It also has the disadvantage that the gas is blown out and the effect of reducing turbine blade pressure loss is diminished.

Although the embodiment of the turbine blade pressure reducing device according to the present invention has been described above, the present invention is not limited to the above-described embodiment. That is, the same effects as those described above can be obtained by applying the present invention to a turbine moving blade other than a gas turbine or a turbine stationary blade instead of the above-described gas turbine moving blade. Furthermore, it is a matter of course that various modifications can be made to the form, structure, configuration, and the like of the present invention without impairing all the substantial advantages of the present invention within the technical scope of the present invention.

[0038]

As described above, according to the turbine blade pressure loss reducing method and apparatus of the present invention, according to the configuration shown in the claims, (1) the cascade outlet Mach number is transonic speed, Alternatively, it is possible to reduce the pressure loss of the high-load, high-expansion-ratio turbine that occurs at a supersonic speed. Thereby, the effect of improving the fuel efficiency of the gas turbine or the like can be obtained.

(2) Further, in the case of a cooling turbine blade in which the turbine blade has a cooling structure, the effect of reducing the pressure loss can be obtained by a simple structure modification, and non-cooling without a cooling structure. In the case of turbine blades, the amount of gas to be ejected from the trailing edge pressure surface can be set freely, without being limited by the design of the cooling structure, minimizing pressure loss with respect to the operating state of the gas turbine. It is possible to select the state quantity of the gas to be blown out. As a result, the pressure loss can be effectively reduced not only at the time of full load but also at the time of partial load, and the fuel efficiency of the gas turbine and the like can be improved in a wide range of operating states of the gas turbine and the like. be able to.

[Brief description of the drawings]

FIG. 1 is a perspective view showing a first embodiment of a turbine blade pressure loss reducing device according to the present invention;

2 is a cross-sectional view taken along the line AA shown in FIG.

FIG. 3 is a detailed cross-sectional view of a portion A shown in FIG. 2;

FIG. 4 is a longitudinal sectional view taken along a line BB in FIG. 1;

FIG. 5 is a perspective view showing a second embodiment of the turbine blade pressure loss reducing device according to the present invention;

FIG. 6 is a detailed cross-sectional view of a trailing edge portion of a turbine blade showing a third embodiment of the turbine blade pressure loss reducing device according to the present invention;

FIG. 7 is a perspective view of a conventional non-cooled turbine blade.

[Description of Signs] 10 Gas turbine rotor blade as turbine blade 12 Airfoil portion 14 Pressure surface 16 Negative pressure surface 18 Leading edge 20 Trailing edge 22 Blade root 24, 24 'Spout (rectangular slit) 26 Spout (circular hole) 30 Blade internal passage 32 Fluid supply port

 ──────────────────────────────────────────────────続 き Continued on the front page (72) Katsuyuki Shimauchi 1200, Higashi-Tanaka, Komaki City, Aichi Prefecture Mitsubishi Heavy Industries, Ltd. Nagoya Guidance and Propulsion System Works Kunihiro Shimizu 1200, Higashi-Tanaka, Kodai, Komaki-shi, Aichi Prefecture Mitsubishi Heavy Industries, Ltd.Nagoya Guidance Propulsion System Works Co., Ltd. Co., Ltd. Nagoya Guidance Propulsion System Works Co., Ltd.

Claims (2)

[Claims] 1. A jet gas supplied to the inside of a turbine blade, which is arranged in the circumferential direction and forms a turbine blade row of equal pitch, is supplied to the inside of the turbine blade at a mass flow rate ratio of 1 or less, and is provided after the turbine blade. By jetting into the mainstream gas from the edge pressure surface to reduce the generation of shock waves from the pressure surface of the turbine blade, disposed to face the pressure surface, the shock wave is incident on the turbine blade A method for reducing turbine blade pressure loss, characterized in that boundary layer enlargement of a suction surface of an adjacent turbine blade is suppressed to reduce pressure loss occurring in the turbine cascade. 2. A turbine blade, which is arranged in a circumferential direction and is formed in a blade root portion of a turbine blade forming a turbine blade row of equal pitch,
A fluid supply port for introducing a gas for ejection into the turbine blade, and a gas for ejection defined from the airfoil portion of the turbine blade and flowing from the fluid supply port, distributed in a height direction of the airfoil portion. And a jetting gas that is provided in the height direction of the trailing edge pressure surface of the airfoil portion and flows from the vane internal passage at a mass flow rate ratio of 1 or less into the mainstream gas. An apparatus for reducing pressure loss of a turbine blade, comprising an injection port.