JP2011094483A - Turbocompressor and tip clearance control method thereof - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a turbocompressor and a tip clearance control method thereof, capable of largely reducing a clearance between the tip of a moving blade cascade arranged on the outer periphery of a rotating disc and a case for surrounding the moving blade cascade, in the whole operation time up to stopping from starting. <P>SOLUTION: This turbocompressor has a plurality of stages of moving blade cascades, and compresses air in a multistage shape, and includes a cooling disc 10 having a plurality of radial directional outer end parts 10a respectively provided with a plurality of moving blade cascades C2-C6 including the final cascade, a single radial directional inner end part 10b having a hollow hole 12 for surrounding a rotary shaft 1 of the moving blade cascades at an interval and a radial directional intermediate part 10c of a plurality of thin plates respectively extending up to a plurality of respective outer end parts from the inner end part, and a cooling structure 20 for cooling an inner surface of the hollow hole 12 of the inner end part by cooling air 6 extracted from an upstream side main flow. The cooling disc 10 has a heat shielding coating 14 on a flow passage surface composed of the outer end part 10a and the moving blade cascades C2-C6. <P>COPYRIGHT: (C)2011,JPO&INPIT

Description

  The present invention relates to a turbo compressor and a tip clearance control method thereof.

  In a turbo compressor constituting a turbofan engine and a turbojet engine, the tip of a moving blade row (hereinafter referred to as “chip”) provided on the outer periphery of a rotating disk and the moving blade row throughout the entire operation time from start to stop. It is necessary to provide an appropriate gap (hereinafter referred to as “chip clearance”) between the case and the surrounding case.

  A compressor cooling device related to the present invention is disclosed in Patent Document 1, for example. Moreover, the control of the tip clearance in the turbo compressor is disclosed in Patent Document 2, for example.

US Pat. No. 5,658,158, “COMPRESSOR ROTOR COOLING SYSTEM FOR A GAS TURBINE” Japanese Patent Application Laid-Open No. 2006-22814, “Blade Clearance Control”

  FIG. 1 is a relationship diagram between the operation time (horizontal axis) of the turbo compressor and the amount of deformation (displacement) (vertical axis) in the radial direction of the case and the disk. In FIG. 1 (A), the case at startup (time 0) and the chip clearance of the disk are called initial clearance (ICL), and the minimum chip clearance when decelerating and stopping is called pinch point (PP).

  In order to ensure the normal operation performance of the turbo compressor, the tip clearance (CL) in the steady state is usually set to be small as long as a pinch point can be secured. However, the conventional turbo compressor has a problem that the displacement due to the thermal expansion of the case at the start-up is larger than the displacement of the disk, so that the tip clearance becomes excessive in a steady state and the performance of the turbo compressor is lowered.

Therefore, when a low thermal expansion material (for example, INCO 909) is used as the case material and the displacement of the case at the time of startup is reduced, the displacement of the case is as shown in FIG.
FIG. 1B shows a case where a conventional disk is used as it is. In this case, the chip clearance (CL) in the steady state can be set small, but the displacement condition with respect to the time of the disk does not change. Therefore, when the chip clearance (CL) in the steady state is set according to the disk, the initial clearance (ICL) ) Is larger than before.

  On the other hand, FIG. 1 (C) shows a case where the displacement of the disk is suppressed to be smaller than the conventional one. By using this disk and a case using a low thermal expansion material, the tip clearance (CL) in the steady state can be set smaller than the conventional one while maintaining the initial clearance (ICL), and a double-headed arrow in the figure. As shown in FIG. 2, the tip clearance at the time of starting and stopping can also be made smaller than before.

  However, a disk that has a moving blade row on the outer periphery and rotates at a high speed has severe mechanical requirements, so a low thermal expansion material such as a case cannot be used. I can't make it smaller. Therefore, there has been a demand for a turbo compressor and its tip clearance control method that can satisfy the above-described displacement condition, that is, can suppress the displacement of the disk to be smaller than the conventional one.

  The present invention has been devised to meet the above-described needs. That is, the object of the present invention is to greatly reduce the gap between the tip of the moving blade row provided on the outer periphery of the rotating disk and the case surrounding the moving blade row over the entire operation time from start to stop. An object of the present invention is to provide a turbo compressor that can be used and a tip clearance control method thereof.

According to the present invention, a turbo compressor having a plurality of stages of moving blade rows and compressing air in multiple stages,
A plurality of radially outer ends each provided with a plurality of blade rows including a final row, a single radially inner end portion having a hollow hole that surrounds the rotation axis of the blade rows at a distance, and A cooling disk having a plurality of thin plate radial intermediate portions extending respectively from a radially inner end portion to a plurality of radially outer end portions;
A cooling structure that cools the inner surface of the hollow hole at the radially inner end with cooling air extracted from the upstream main stream, and
The cooling disk has a thermal barrier coating on a flow path surface including a radially outer end portion and a moving blade row, and a multistage turbo compressor is provided.

According to an embodiment of the present invention, the thermal barrier coating has a thickness of 0.1 mm to 0.5 mm.
The thermal barrier coating is yttria stabilized zirconia.

According to the apparatus and method of the present invention described above, the inner surface of the hollow hole at the radially inner end is cooled by the cooling air extracted from the upstream main stream, so that the single radially inner end of the cooling disk is cooled by the cooling air. Can be kept at a low temperature close to.
Further, since a heat shielding coating (for example, yttria-stabilized zirconia) is provided on the flow path surface including the radially outer end portion and the moving blade row, heat input from high-temperature air (for example, 500 to 700 ° C.) flowing through the flow path surface is blocked. Heating can suppress the temperature rise at the radially outer end portion and the radially intermediate portion to be small, and the displacement due to the thermal expansion of the entire cooling disk can be reduced.
Therefore, by using the case using the cooling disk and the low thermal expansion material according to the present invention, the tip of the moving blade row provided on the outer periphery of the rotating disc and the moving blade row are arranged over the entire operation time from start to stop. A gap between the surrounding case and the surrounding case can be greatly reduced.

FIG. 6 is a relationship diagram between the operation time of the turbo compressor and the radial displacement of the case and the disk. It is a partial block diagram of the jet engine which has the turbo compressor by this invention. 2 is a partial model of a turbo compressor according to the present invention. It is a figure which shows the temperature difference in the partial model of FIG. It is a simulation result of the partial model of FIG. It is a related figure of the thickness of thermal barrier coating by simulation, and disk displacement.

  DESCRIPTION OF EXEMPLARY EMBODIMENTS Hereinafter, the best mode for carrying out the invention will be described with reference to the drawings. In addition, the same code | symbol is attached | subjected to the common part in each figure, and the overlapping description is abbreviate | omitted.

FIG. 2 is a partial configuration diagram of a jet engine having a turbo compressor according to the present invention.
In this figure, 1 is a rotating shaft of a moving blade row, 2 is a hollow long shaft connecting a turbine and a compressor, 3 is a case surrounding the moving blade row, and 4 is a turbine nozzle. 2A is a diagram showing only the upper half of the rotary shaft 1, and FIG. 2B is an enlarged view of the rotor blade row C5.

  In this figure, the turbo compressor according to the present invention has a plurality of stages (six stages in this example) of moving blade rows C1 to C6, and compresses the air 5 in multiple stages.

The turbo compressor of the present invention includes a cooling disk 10 and a cooling structure 20.
The cooling disk 10 has a plurality of radially outer end portions 10a, a single radially inner end portion 10b, and a plurality of radially intermediate portions 10c. Hereinafter, “radial direction” is omitted, and they are referred to as an outer end portion 10a, an inner end portion 10b, and an intermediate portion 10c, respectively.

A plurality (five in this example) of moving blade rows C2 to C6 including the last row of moving blade rows C6 are provided on the plurality (five in this example) of outer end portions 10a. Installation of the moving blade rows C2 to C6 on the outer end 10a is preferably integrally formed, but may be separately formed and assembled.
The mainstream temperature in the plurality of rotor blade rows including the final row is, for example, 500 to 700 ° C.

The single inner end portion 10b has a hollow hole 12 surrounding the rotating shaft 1 of the rotor blade row with a predetermined interval. The inner diameter of the hollow hole 12 is set so that the cooling air 6 flows at a predetermined flow rate through the gap between the hollow hole and the long shaft 2.
The inner end portion 10b may be assembled as long as it is “single”.

  The plurality (five in this example) of intermediate portions 10c are formed of a plurality (five in this example) of thin plates respectively extending from the inner end 10b to each of the plurality of outer ends 10a. Each thin plate is set to be thin (for example, 4 to 9 mm) so as to suppress heat conduction from the outer end portion 10a to the inner end portion 10b.

  The cooling structure 20 cools the inner end portion 10b by flowing the cooling air 6 extracted from the upstream main flow downstream along the inner surface of the hollow hole 12 of the inner end portion 10b. The temperature of the cooling air 6 is preferably 200 to 300 ° C.

The cooling disk 10 described above has a thermal barrier coating 14 on the flow path surface composed of the outer end portion 10a and the moving blade rows C2 to C6.
The thermal barrier coating 14 is preferably yttria stabilized zirconia (YSZ) having a thickness of 0.1 mm to 0.5 mm.

Using the turbo compressor described above, the tip clearance control method according to the present invention is:
(1) Provided with the cooling disk 10 described above,
(2) A thermal barrier coating 14 made of yttria-stabilized zirconia (YSZ) having a thickness of 0.1 mm to 0.5 mm is applied to the flow path surface formed of the outer end portion 10a of the cooling disk 10 and the moving blade row,
(3) The inner surface of the hollow hole 12 of the inner end portion 10b is cooled with 200 to 300 ° C. cooling air 6 extracted from the upstream main stream.

  The application position of the thermal barrier coating is not limited to this example, and is preferably after the middle stage, preferably after the bleed stage (extraction position stage). This is to obtain the cooling effect according to the present invention in the latter stage where cooling is difficult.

  According to the above-described apparatus and method of the present invention, the inner surface of the hollow hole 12 in the radially inner end portion 10b is cooled by the cooling air 6 extracted from the upstream main flow, so The end 10b can be held at a low temperature (for example, around 300 ° C.) close to the cooling air 6.

  Further, since the heat shielding coating 14 (for example, yttria-stabilized zirconia) is provided on the flow path surface including the radially outer end portion 10a and the moving blade row, heat input from high-temperature air (for example, 500 to 700 ° C.) flowing through the flow path surface. The temperature rise of the radially outer end portion 10a and the radially intermediate portion 10c can be suppressed small (for example, 5 to 20 ° C. lower than the conventional temperature), and the displacement due to the thermal expansion of the entire cooling disk can be reduced. it can.

  Therefore, by using the cooling disk 10 according to the present invention and the case using the low thermal expansion material (case B in FIG. 1), it is provided on the outer periphery of the rotating disk (cooling disk 10) in the entire operation time from start to stop. The gap between the tip of the moving blade row and the case (case B) surrounding the moving blade row can be greatly reduced.

FIG. 3 is a partial model of a turbo compressor according to the present invention.
In this figure, the width (axial length) of the outer end portion 10a of the cooling disk 10 is 25 mm, the radius is 170 mm, the thickness (axial length) of the intermediate portion 10c is 5 mm, and the radius of the hollow hole 12 is 55 mm. is there.
In this example, the main stream temperature in the moving blade row is 395 ° C., and the temperature of the cooling air 6 is 283 ° C.

FIG. 4 is a diagram showing a temperature difference in the partial model of FIG.
As shown in this figure, in the following description, the temperature difference between the mainstream center in the rotor blade row and the surface of the thermal barrier coating 14 is dT1, and the temperature difference between the surface of the thermal barrier coating 14 and the outer surface of the outer end portion 10a. dT2, the temperature difference between the outer surface of the outer end 10a and the inner surface of the hollow hole 12 is dT3, and the temperature difference between the inner surface of the hollow hole 12 and the flow path center of the cooling air 6 is dT4.

FIG. 5 shows a simulation result of the partial model of FIG. In this figure, the horizontal axis represents the thickness of the thermal barrier coating 14, and the vertical axis represents the temperature difference of each part (dT1 to dT4).
From this figure, the temperature difference dT2 due to the thermal barrier coating 14 is 5 to 20 ° C. when the thickness is 0.1 mm to 0.5 mm, compared to the case where the thermal barrier coating 14 is not present (thickness: 0). It can be seen that the temperature rise of the outer end portion 10a and the intermediate portion 10c can be suppressed to that extent.

FIG. 6 is a relationship diagram between the thickness of the thermal barrier coating and the disc displacement by simulation. In this figure, the horizontal axis represents the thickness of the thermal barrier coating 14, and the vertical axis represents the displacement of the disk (the tip of the rotor blade row).
From this figure, the displacement can be reduced to a maximum of 0.82 mm when the thickness is 0.1 mm to 0.5 mm, compared to the displacement 0.85 mm when the thermal barrier coating 14 is not provided (thickness: 0). Recognize.

Table 1 is a graph showing the relationship between the thickness of the thermal barrier coating and the performance of the turbo compressor by simulation. This table compares the case where there is no thermal barrier coating 14 (thickness: 0) and the case where the thickness is 0.5 mm.
The disc displacement is the sum of the thermal displacement of the disc and the mechanical displacement of the disc.

  From this result, it can be seen that the coating with a thickness of 0.5 mm reduces the displacement of the disk at high rotation by 0.03 mm (0.85 mm → 0.82 mm). In this case, the clearance during assembly (initial clearance in Table 1) is adjusted (0.45 mm → 0.42 mm) so that the tip clearance during high rotation is kept the same, so that the clearance during low rotation (Table 1) is adjusted. Clearance) is improved by 12% (0.26 mm → 0.23 mm).

  In addition, this invention is not limited to embodiment mentioned above, Of course, it can change variously in the range which does not deviate from the summary of this invention.

1 rotating shaft, 2 long shaft, 3 case,
4 turbine nozzle, 5 air, 6 cooling air,
10 cooling disk,
10a outer end portion, 10b inner end portion, 10c intermediate portion,
12 hollow holes, 14 thermal barrier coating,
20 Cooling structure

Claims (3)

A turbo compressor having a plurality of stages of moving blade rows and compressing air in multiple stages,
A plurality of radially outer ends each provided with a plurality of blade rows including a final row, a single radially inner end portion having a hollow hole that surrounds the rotation axis of the blade rows at a distance, and A cooling disk having a plurality of thin plate radial intermediate portions extending respectively from a radially inner end portion to a plurality of radially outer end portions;
A cooling structure that cools the inner surface of the hollow hole at the radially inner end with cooling air extracted from the upstream main stream, and
The turbo compressor according to claim 1, wherein the cooling disk has a thermal barrier coating on a flow path surface including a radially outer end portion and a moving blade row.   The turbo compressor according to claim 1, wherein the thermal barrier coating has a thickness of 0.1 mm to 0.5 mm.   The turbo compressor according to claim 1, wherein the thermal barrier coating is yttria stabilized zirconia.

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