JP2012072706A - Method for modifying gas turbine device - Google Patents

Abstract

PROBLEM TO BE SOLVED: To solve the problem of inevitable enormous loss in many aspects including the entire development process such as generation of temperature rise of a thrust bearing in the case thrust force exceeds planned value to drastically damage the reliability as a rotor in a test operation of the gas turbine at the final stage of development, need of design change for the entire gas turbine when drastically modifying the thrust force of the gas turbine after completion of assembly, and in particular, need of enormous cost and period even only for precision casting when dramatically changing the profile of a cooling blade of the turbine.SOLUTION: The efflux angle of the moving blade is deflected to the ventral side by increasing the maximum blade thickness by inclining the mounting angle of the moving blade to the ventral side with respect to the axial line of a gas turbine or bulging the back side profile of the moving blade for a moving blade of a turbine.

Description

  The present invention relates to a gas turbine apparatus, and more particularly, to a thrust force reducing method for improving the reliability of a gas turbine under development that has generated a thrust force exceeding a planned value.

  The air sucked in by the gas turbine passes through a cycle process of pressurization (compression) by the compressor, combustion by the combustor, expansion and exhaust at the turbine section and below. At this time, fluid force acts from the discharge side to the suction side in the compressor and from the inflow side to the exhaust side in the turbine. Conceptually speaking, a difference in fluid force acting on the compressor and the turbine is supported by a thrust bearing as a thrust force on the rotating body. In general, this thrust bearing employs an oil supply type, which not only supports thrust force but also lubricates the bearing and removes heat generated by bearing loss, contributing to maintaining the reliability of the bearing as a rotating body. is doing.

  Various measures for improving the performance of this thrust bearing have been studied. For example, Patent Document 1 discloses a high-accuracy thrust bearing that improves load capacity and has a small amount of movement in the axial direction.

Japanese Unexamined Patent Publication No. Sho 63-172012

  In order to efficiently realize a gas turbine that is one of turbomachines in accordance with the purpose, it is an effective means to configure a gas turbine with reference to the prior art. By the way, in recent years, in the gas turbine design, with the development of computers and the verification and progress of analysis software, work has been advanced quickly and accurately. However, it is still on the way and not completed. If the development of the gas turbine itself is considered as a process from planning, basic / detailed design, production drawing creation, parts production, assembly, and trial operation, a period of several years is required. One of the most fundamental factors in the efficient development of gas turbines is the elimination of design errors. However, as long as the design technology has not been completed, it is difficult to completely eliminate design errors including human errors.

  For example, in a trial operation of a gas turbine that is the final stage of development, the thrust force may exceed the planned value. In this case, the thrust bearing causes a temperature rise, and the reliability as a rotating body is greatly impaired. In addition, when a thorough review of the thrust force of a gas turbine that has been assembled is to be fundamentally reviewed, a design change for the entire gas turbine is imposed. In particular, when it is necessary to drastically change the blade shape of the turbine cooling blade, even if only the precision casting is seen, it will consume a great deal of cost and time. I cannot escape the loss.

  Accordingly, an object of the present invention is to provide a method for maintaining the reliability as a rotating body device by reducing the thrust force exceeding the planned value without causing a significant increase in cost or prolonging the production period. It is in.

  In order to achieve the above object, a gas turbine apparatus of the present invention mainly includes a compressor, a combustor, and a turbine having a stationary blade and a moving blade, and supports a fluid force acting on a rotating body including the moving blade. A gas turbine having a thrust bearing is characterized in that a thrust force is reduced by increasing an outflow angle of a moving blade of the turbine.

  According to the present invention, it is possible to maintain the reliability as a rotating body device by reducing the thrust force exceeding the planned value without causing a significant increase in cost or prolonging the production period.

It is the conceptual diagram which showed the structure of the gas turbine. FIG. 2 is a cross-sectional view showing the details of the thrust bearing, showing the AA portion of FIG. 1. It is sectional drawing of the turbine part which showed the structure of this invention. It is sectional drawing of the airfoil which shows the comparison with the conventional blade blade type | mold of this invention, and an improvement. It is explanatory drawing showing the pressure point when the gas turbine of this invention is drive | operated. It is sectional drawing of the turbine part which showed the thrust force which acts on the turbine 1st stage of this invention.

〔Example〕
Embodiments of the present invention will be described below with reference to FIGS. 1, 2, 3, 4, 5, and 6. In each figure, the same number represents the same apparatus or member.

  The overall configuration of the gas turbine will be described with reference to FIG. The gas turbine 1 includes a compressor 2 that mainly compresses intake air, a combustor 3 that burns the compressed air together with fuel to generate high-temperature and high-pressure gas, a turbine 4 that expands the high-temperature and high-pressure gas, and a generator 5, and a plurality of rotating members (not shown) in the compressor 2 and the turbine 4 are connected as a shaft 6 in one axis. Further, the generator 5 is connected to the shaft 6 via a coupling 11. The shaft 6 is supported by a first journal bearing 7 and a second journal bearing 8, and a first thrust bearing 9 and a second thrust bearing 10 are provided in the vicinity of the first journal bearing 7. . From the intermediate stage and the final stage of the compressor 2, a compressor intermediate stage bleed path 30 and a compressor final stage bleed path 31 are formed and introduced into cooling blades (to be described later) of the turbine 4.

  Next, the detailed cross section of the AA part of FIG. 1 is demonstrated using FIG. As described above, the shaft 6 is defined as an assembly of rotating members. In the compressor 2, the compressor shaft end shaft 19 and the plurality of compressor wheels 20 are connected by a compressor stacking bolt 21 and become one of the members constituting the shaft 6. The bearing housing 12 provided in the compressor casing 13 accommodates the first journal bearing 7, the first thrust bearing 9, and the second thrust bearing 10, and is provided on the compressor shaft end shaft 19. The annular projecting portion 24 is formed with an anti-load side pressure receiving surface 23a and a load side pressure receiving surface 23b, and the first thrust bearing 9 and the second thrust bearing 10 are installed facing the surfaces. These bearings are supplied with lubricating oil from a lubricating oil supply device (not shown), and the exhaust oil has a circulation line that returns to the lubricating oil supply device.

  The configuration on the turbine side will be described with reference to FIG. In order to clarify the intention of the present invention, only the two-stage portion on the front stage side of the turbine 4 is shown in FIG. The turbine 4 is mainly positioned between a turbine casing 14 that serves as a stationary member, a first stage stationary blade 15a and a second stage stationary blade 16a that are fixed to the turbine casing 14, and a first wheel 17a and a second wheel 17b that serve as rotating members. The spacer 18 includes a first stage moving blade 15b fixed to the outer peripheral end of the first wheel 17a, and a second stage moving blade 16b fixed to the outer peripheral end of the second wheel 17b. Similar to the compressor described above, the inner barrel 25, the first wheel 17a, the second wheel 17b, and the spacer 18 are connected by the turbine stacking bolt 22 and become one of the components of the shaft 6. Note that the bleed air passing through the compressor intermediate stage bleed passage 30 is supplied to the second stage stationary blade 16a, and the bleed air passing through the compressor final stage bleed passage 31 is supplied to the first stage stationary blade 15a. Similarly, there is a system for supplying extracted air to the moving blades, but this is omitted.

  With reference to FIG. 4, the airfoil of the present embodiment will be described in comparison with a conventional blade. The cross-sectional view of the airfoil of FIG. 4 shows the average diameter cross section of the one-stage moving blade 15b, with the conventional blade as a broken line and the improved blade as a solid line. First, the airfoil is defined using the airfoil 48 of the improved wing. The airfoil has a continuous leading edge 44 that is an arc in contact with the leading edge diameter 43, a back profile 41 a of the improved wing, a ventral profile 41 b of the improved wing, and a trailing edge 46 that is an arc in contact with the trailing edge diameter 45. Connect and form. At this time, the maximum blade thickness 42b of the improved blade is given as ΦD2. With respect to the axial line 50, an angle α1 at which the mainstream gas flows is defined as an inflow angle, and an angle α2 at which the mainstream gas flows out with respect to the auxiliary line 52 parallel to the axial line 50 is defined as an outflow angle. On the other hand, an angle formed between the auxiliary line 51 that is a straight line contacting the front edge 44 and the rear edge 46 and the auxiliary line 52 is set as a set angle β1. Referring to this, the maximum blade thickness 42a of the conventional airfoil 47 is given as ΦD1, the outflow angle of the improved airfoil 48 is α2_ improved, the set angle is β1_ improved, and the conventional airfoil airfoil The outflow angle of 47 is called α2_conventional and the set angle is called β1_conventional.

  The maximum blade thickness of the improved blade is in a relationship of ΦD1 <ΦD2 after correcting in the direction in which only the back profile 41a is expanded. The increase in the maximum blade thickness 42b of the improved wing due to the improved wing dorsal profile 41a causes the dorsal profile 41a of the improved wing to bulge to the right side in the drawing, resulting in a steep slope at the connection with the trailing edge 46. Therefore, the outflow angle is deflected about 1 degree in the clockwise direction on the drawing. Further, the improved wing airfoil 48 is rotated by about 1 degree in the direction of the ventral profile 41b (clockwise in the drawing) with respect to the conventional wing airfoil 47, so that β1_improvement> β1_conventional. This simply deflects the outflow angle once in the clockwise direction on the drawing. As a result, the outflow angle α2_improvement of the airfoil 48 of the improved wing has been deflected by about 2 degrees in total with respect to α2_conventional.

  In this embodiment configured as described above, the high-temperature and high-pressure working gas generated in the compressor 2 and the combustor 3 along with the operation of the gas turbine 1 has a total pressure of about 1.6 MPa and a temperature of about 1300 ° C. While the turbine works at each stage including the four first stage stationary blades 15a and four stage stationary blades 15b, the pressure and temperature are reduced and the final stage blade (not shown) flows out at about 600 ° C. At this time, the generator 5 connected to the shaft 6 rotates to obtain electric power.

  Since the turbine blades are exposed to high-temperature gas, a part of the high-pressure air obtained by the compressor 2 is extracted, and the turbine intermediate stage extraction path 30 and the compressor final-stage extraction path 31 are respectively extracted. The air is introduced into the two-stage stationary blade 16a and the first-stage stationary blade 15a, and is cooled below the working gas temperature. In addition, the description about a moving blade is omitted.

  The pressure at each point of the gas turbine 1 at this time will be described with reference to FIG. 5 on the basis of the total pressure. The compressor suction point P0 pressure of the compressor 2 is atmospheric pressure, and the compressor discharge point P1 pressure is The maximum is about 1.6 MPa. The turbine inlet pressure point P2 decreases to about 1.52 MPa due to pressure loss in the combustor 3, and after further decreasing due to subsequent turbine work, the turbine outlet pressure point P3 becomes about 0.11 MPa. Finally, since this exhaust gas is released into the atmosphere by the chimney, the exhaust pressure point P4 returns to the same atmospheric pressure as the compressor suction point P0 pressure. This pressure distribution is monotonously increased until the compressor discharge, and monotonically decreased until the exhaust, and roughly speaking, with reference to FIG. A fluid force is generated in the left direction in the drawing and a fluid force toward the turbine side (in the right direction in the drawing), and the difference acts as a thrust force, so that the shaft 6 moves in the axial direction. The axial movement of the shaft 6 is supported by the first thrust bearing 9 or the second thrust bearing 10. Lubricating oil from a lubricating oil system (not shown) is supplied to the bearing, and the frictional heat accompanying the bearing is recovered by this lubricating oil, so that the temperature rise of the bearing is usually suppressed.

  Regarding the present invention and the thrust force, the first stage of the turbine 4 will be described with reference to FIG. The pressure of the first stage stationary blade inlet 60 is about 1.52 MPa of the total pressure of the working gas that becomes the P2 point. Up to this point, for the sake of convenience, the explanation has been made with the total pressure, but the evaluation of the thrust force is based on the static pressure. In the conventional blade, the pressure is about 1.0 MPa at the first stage blade inlet 61 and about 0.82 MPa at the first stage blade outlet 62. These pressures act on the first stage rotor blade 15b and the first wheel 17a, which are rotating bodies, as indicated by arrows 63 and 64 in the figure. At this time, in the first stage of the turbine, the fluid force obtained by multiplying the pressure-receiving area (not shown) in the axial direction of the first stage blade 15b and the first wheel 17a to which the pressure of the first stage blade inlet 61 is applied is multiplied by the pressure. The fluid force obtained by multiplying the pressure-receiving area (not shown) in the axial direction of the first stage blade 15b and the first wheel 17a acting in the right direction of the drawing and conversely to which the pressure of the first stage blade outlet 62 is applied. Acts in the left direction of the drawing, and the difference is the thrust force acting on the first stage of the turbine. When this thrust force is added up for each stage, the above-described thrust force is directed toward the turbine side. Similarly, if applied to the compressor side, the thrust force is directed toward the compressor side suction side. Further, the sum is supported by a thrust bearing as a thrust force acting on the shaft 6. In the conventional blade, the total thrust force is greatly directed to the compressor side, so that the second thrust bearing 10 is supported and the bearing surface pressure is about 5 MPa, and the oil discharge temperature rises.

  As described above, the total thrust force is determined by the pressure between the stages and the axial pressure receiving area. However, if the shaft 6 has the same shape, the thrust force is a function of only the pressure between the stages. Is determined. The pressures at the first stage blade inlet 61 and the first stage blade outlet 62 vary depending on the work distribution (reaction degree) between the stages. By the way, in the improved wing, the outflow angle was deflected by about 2 degrees compared to α2_conventional, so the reaction degree increased and the work to be shared increased, and the outlet of the first stage blade Although there is no change in the pressure of about 0.82 MPa at 62, the pressure rises to about 1.05 MPa at the first stage blade inlet 61. As a result, regarding the first stage blade 15b and the first wheel 17a, the thrust force toward the exhaust side of the turbine is increased, and as a sum, the thrust force is directed toward the suction side of the compressor. The load borne by the second thrust bearing 10 is reduced. At this time, the bearing surface pressure is about 3 MPa, and an increase in the oil discharge temperature can be suppressed.

  In the gas turbine apparatus described above, the set angle of the moving blade with respect to the axial line of the gas turbine is inclined in the ventral profile direction, and the back profile of the moving blade is changed to the direction in which it is expanded to increase the maximum blade thickness. By increasing, the blade outflow angle is deflected to the back profile side of the blade, and the reaction force of the first stage of the turbine, that is, the static pressure between the stages is changed, thereby adjusting the thrust force generated in the stage. Therefore, the thrust force applied to the thrust bearing can be reduced, and a highly reliable gas turbine can be provided. Turbine blades are manufactured by precision casting using a core, but the change on the increase side of the back profile is a modification that cuts the existing core and can be diverted. Increase in cost and prolongation of the production period can be avoided.

  In the present embodiment, the first stage has been described as an example. However, the present invention is not limited, and even when applied to other stages, the same effect can be obtained and a highly reliable gas turbine apparatus can be expected.

  In addition, the change in the rotor blade outflow angle increases the incidence angle to the two-stage stationary blade on the rear stage, which causes an increase in loss. However, the stationary blade is dealt with by adopting a blunt blade with excellent incidence characteristics. Is possible.

  The present invention is applicable to gas turbines for industrial use and power generation.

DESCRIPTION OF SYMBOLS 1 Gas turbine 2 Compressor 4 Turbine 6 Shaft 9 1st thrust bearing 10 2nd thrust bearing 15b 1st stage moving blade 41a Back blade profile 41b of improved blade Front side profile 42b of improved blade Maximum blade thickness α2 of improved blade Angle β1 Set angle

Claims (7)

A compressor for compressing air;
A combustor for combusting compressed air and fuel compressed by the compressor;
A turbine driven by combustion gas generated in the combustor;
A shaft connecting the compressor and the turbine;
In a method for modifying a gas turbine apparatus having a journal bearing and a thrust bearing for supporting the shaft,
A method for remodeling a gas turbine device, wherein an outflow angle of a turbine rotor blade is set based on a thrust force acting on the shaft.   2. The gas turbine apparatus modification method according to claim 1, wherein the outflow angle of the moving blade is changed by changing a set angle of the moving blade with respect to an axial line of the gas turbine. Remodeling method.   2. The gas turbine apparatus remodeling method according to claim 1, wherein the outflow angle of the moving blade is changed by changing a maximum blade thickness of the moving blade.   4. The gas turbine apparatus remodeling method according to claim 3, wherein the change in the maximum blade thickness is performed by changing the back side profile larger than the ventral side profile of the moving blade. .   4. The gas turbine apparatus remodeling method according to claim 3, wherein the change in the maximum blade thickness is performed only by changing the back profile of the moving blade.   3. The gas turbine apparatus remodeling method according to claim 2, wherein the change of the set angle of the moving blade and the change of the blade profile of the moving blade are used together to change the outflow angle of the moving blade, A method for remodeling a gas turbine device, characterized in that the deflection amount to the outflow angle is made equal and each deflection amount is set to 1 deg or less.   7. The gas turbine apparatus remodeling method according to claim 2, wherein the deflection of the outflow angle of the moving blade is set to a ventral profile direction of the moving blade. Gas turbine device.

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