JP2011190714A - Axial flow compressor and gas turbine engine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To further extend the operation range of an axial flow compressor 1 to a low flow rate side while suppressing a stall of the axial flow compressor 1. <P>SOLUTION: The axial flow compressor is constructed so that an outlet cross-sectional area of one of compressor rotors 11 is larger than the inlet side cross-sectional area. In a tangential cross section, the outer peripheral face of a disk 13 at the one compressor rotor 11 shifts from a first curve R1, which is convex toward a downstream direction, to a second concave curve R2. An inflection point IP shifting from the first curve R1 to the second curve R2 lies between a remote point by 2.8% of code length from the maximum blade thickness position TP of a rotor blade 15 toward upstream, and the remote point by 5.0% of the code length from the maximum blade thickness position TP of the rotor blade 15 toward downstream. <P>COPYRIGHT: (C)2011,JPO&INPIT

Description

  The present invention relates to an axial flow compressor that is used as, for example, a component of a gas turbine engine and compresses and conveys gas in an axial direction.

  A general axial compressor includes a cylindrical compressor case as a base, and the compressor case extends in the axial direction. An annular gas flow path is formed inside the compressor case.

  In the compressor case, a plurality of stages of compressor stators are provided along the axial direction. The compressor stator at each stage includes a plurality of stationary blades arranged at equal intervals in the circumferential direction, and each stationary blade is located in the gas flow path.

  In the compressor case, a plurality of stages of compressor rotors are provided alternately with the plurality of stages of compressor stators along the axial direction. In addition, each stage of the compressor rotor includes a disk that is provided in the compressor case so as to be rotatable about its axis, and the outer peripheral surface of the disk is a channel surface radially inward of the gas channel. Part. Further, a plurality of blades are provided at equal intervals on the outer peripheral surface of the disk in the compressor rotor of each stage, and each blade is located in the gas flow path.

  Therefore, by rotating the multi-stage compressor rotor, the multi-stage compressor rotor and the multi-stage compressor stator cooperate with each other, and the gas taken into the gas flow path is compressed and conveyed in the axial direction. be able to.

  On the other hand, various developments have been made to expand the operating range of an axial compressor, and the applicant of the present application has applied for an axial compressor with an expanded operating range and has already obtained a patent ( Patent Document 1). In the axial compressor according to this prior art, in addition to the configuration of the general axial compressor described above, the outlet side cross-sectional area (outlet cross-sectional area) of any compressor rotor is the inlet side cross-sectional area (inlet cross-section It is configured to be larger than (area). As a result, the axial flow velocity (axial flow velocity) of the gas is reduced in any compressor rotor, and considering the velocity triangle of the moving blade, the work of the moving blade can be performed without increasing the warpage of the moving blade. It can be secured sufficiently. In other words, while ensuring sufficient work of the rotor blade, reducing warpage of the rotor blade and reducing the pressure difference between the pressure surface and the suction surface of the rotor blade, the negative pressure from the pressure surface side of the rotor blade is reduced. The clearance flow to the pressure surface side can be reduced to lower the stall point, and the stall can be prevented even at a low flow rate, and the operating range of the axial compressor can be expanded to the low flow rate side.

Japanese Patent No. 3743415

  By the way, in recent years, in the field of jet engines, the demand for expanding the operating range of the axial compressor has become stronger, and the development of an axial compressor that can further expand the operating range has become an urgent task. It is coming.

  Accordingly, an object of the present invention is to provide an axial flow compressor having a novel configuration capable of further expanding the operating range by improving the axial flow compressor according to the above-described prior art.

  As a result of repeating trial and error in order to solve the above-mentioned problems, the inventor of the present invention has determined that the outlet side sectional area (outlet sectional area) of any compressor rotor is larger than the inlet side sectional area (inlet sectional area). In the meridional section (side section), the outer peripheral surface of the disk in any of the compressor rotors is convex from the convex first curve to the concave second curve. 3 (in the case of the new compressor rotor shown in FIG. 3 (a)), when the outer peripheral surface of the disk is an inclined straight line (in the case of the comparative compressor rotor shown in FIG. 3 (b)) )), The new knowledge that the pressure difference between the pressure surface and the suction surface on the leading edge side of the moving blade can be made smaller (the portion surrounded by the one-dot chain line in FIG. 4), as shown in FIG. The present invention has been completed. In the case of the new compressor rotor shown in FIG. 3A, in the meridional section, the upstream end of the outer peripheral surface of the disk of any compressor rotor is the inner wall surface (radial direction) of the compressor stator on the upstream side. The downstream end of the outer peripheral surface of the disk of any compressor rotor has a diameter larger than the extension line of the inner wall surface of the downstream compressor stator. Located on the outside in the direction or on this extension.

  Here, FIG. 3 (a) is a schematic meridional section showing a new compressor rotor having characteristics relating to new knowledge, and FIG. 3 (b) is a compressor for comparison with the new compressor rotor. FIG. 4 is a schematic meridional section showing the rotor, for a new compressor rotor and a comparative compressor rotor, the distance from the leading edge of the blade and the coefficient of increase in static pressure at the 97% span of the blade (positive It is a figure which shows the relationship with the static pressure rise coefficient of a pressure side and a suction side, Comprising: This relationship is calculated | required by three-dimensional steady viscosity CFD (Computational Fluid Dynamics) analysis. In the new compressor rotor shown in FIG. 3A, the radial position (the radial position) of the tip of each rotor blade is the same from the leading edge to the trailing edge.

  A first feature of the present invention is an axial compressor that compresses and conveys gas in an axial direction, and extends in the axial direction, and has a cylindrical shape with an annular gas flow path formed inside. Compressor case, and a plurality of stages of compressor stators provided in the compressor case along the axial direction and provided with a plurality of stationary blades disposed at equal intervals in the circumferential direction and positioned in the gas flow path And a plurality of stages of compressor rotors provided alternately in the axial direction in the compressor case, the outer peripheral surface constituting a part of the radially inner flow path surface of the gas flow path, and the axial center A disk rotatable around, and a multi-stage compressor rotor provided with a plurality of moving blades provided at equal intervals on the outer peripheral surface of the disk and positioned in the gas flow path, The outlet side sectional area (outlet sectional area) of the compressor rotor is the inlet side sectional area (inlet sectional area). In the meridional section (side section), the outer peripheral surface of the disk in any one of the compressor rotors changes from a convex first curve to a concave second curve in the downstream direction. The upstream end of the outer peripheral surface of the disk in any one of the compressor rotors is radially inward from the extension line of the inner wall surface of the compressor stator on the upstream side or the extension line (the upstream compressor). An extension line of the inner wall surface of the stator), and the downstream end of the outer peripheral surface of the disk in any one of the compressor rotors is radially outer than the extension line of the inner wall surface of the compressor stator on the downstream side. The gist is that it is located on this extension line (extension line of the inner wall surface of the compressor stator on the downstream side).

  In addition, the said axial flow compressor is not restricted to the compressor used as a component apparatus of gas turbine engines, such as a jet engine.

  According to the first feature, the gas introduced into the gas flow path by rotating the plurality of stages of the compressor rotors so that the plurality of stages of the compressor rotors and the plurality of stages of the compressor stators cooperate with each other. Can be transported compressed in the axial direction.

  And the outlet side cross-sectional area of any one of the compressor rotors is configured to be larger than the inlet side cross-sectional area, and in the meridian cross section, the outer peripheral surface of the disk in any one of the compressor rotors is in the downstream direction. Since the transition from the convex first curve to the concave second curve is applied, the outer peripheral surface of the disk in any one of the compressor rotors is an inclined straight line when the above-described novel knowledge is applied. As compared with the case, the pressure difference between the pressure surface and the suction surface on the leading edge side of the moving blade, which greatly affects the stall of the axial compressor, can be further reduced.

  The gist of the second feature of the present invention is that the gas turbine engine includes the axial flow compressor having the first feature.

  According to the feature of the second feature, the same effect as that of the first feature is achieved.

  According to the present invention, since the pressure difference between the pressure surface and the suction surface on the leading edge side of the moving blade can be further reduced, the clearance flow from the pressure surface side to the suction surface side of the moving blade is reduced, and the shaft The operating range of the axial compressor can be further expanded to the low flow rate side while suppressing stall of the flow compressor.

It is a typical meridional sectional view showing the principal part of the axial flow compressor concerning the embodiment of the present invention. It is the typical figure which looked at some compressor rotors in the axial flow compressor concerning the embodiment of the present invention from the diameter direction outside. FIG. 3A is a schematic meridional sectional view (side sectional view) showing a new compressor rotor having characteristics relating to new knowledge, and FIG. 3B is a diagram for comparison of the new compressor rotor. It is a typical meridional sectional view showing a compressor rotor. For the new compressor rotor and the comparative compressor rotor, the distance from the leading edge of the rotor blade and the static pressure increase coefficient at the 97% span of the rotor blade (static pressure increase coefficient on the pressure side and suction side) It is a figure which shows a relationship. It is a figure which shows the relationship between a flow coefficient and a static pressure increase coefficient about the compressor rotor (new compressor rotor) which concerns on embodiment of this invention, and the compressor rotor for a comparison. It is a figure explaining the relationship between the distance from the largest blade thickness position of a moving blade to an inflection point, and the flow coefficient in a stall point about the compressor rotor which concerns on embodiment of this invention, and the compressor rotor for a comparison.

  An embodiment of the present invention will be described with reference to FIGS. In the figure, “F” indicates the forward direction (upstream direction), and “R” indicates the backward direction (downstream direction).

  As shown in FIGS. 1 and 2, an axial flow compressor 1 according to an embodiment of the present invention is used as a compressor that is a component device of a jet engine (an example of a gas turbine engine), and gas (implementation of the present invention). In the form, air is compressed in the axial direction (rearward direction) and conveyed. And the specific structure of the axial flow compressor 1 which concerns on embodiment of this invention is as follows.

  The axial flow compressor 1 includes a cylindrical compressor case 3 as a base, and the compressor case 3 extends in the axial direction (front-rear direction). An annular gas flow path 5 is formed inside the compressor case 3.

  In the compressor case 3, a plurality of stages (only two stages are shown in FIG. 1) of compressor stators 7 are provided along the axial direction. The compressor stator 7 at each stage includes a plurality of stationary blades 9 arranged (aligned) at equal intervals in the circumferential direction, and each stationary blade 9 is located in the gas flow path 5. Yes.

  In the compressor case 3, a plurality of stages of compressor rotors 11 (only one stage is shown in FIG. 1) are provided alternately with the plurality of stages of compressor stators 7 along the axial direction. The machine rotor 11 is integrally connected to a plurality of stages of turbine rotors (not shown) in a turbine (not shown) which is a component device of the jet engine. The compressor rotor 11 at each stage includes a disk 13 provided in the compressor case 3 so as to be rotatable around an axis, and the outer peripheral surface of the disk 13 is radially inward of the gas flow path 5. This constitutes a part of the flow path surface 5s. Further, a plurality of moving blades 15 are provided at equal intervals on the outer peripheral surface of the disk 13 in each stage of the compressor rotor 11, and each moving blade 15 is located in the gas flow path 5.

  Then, the principal part of embodiment of this invention is demonstrated.

  The outlet side cross-sectional area (outlet cross-sectional area) of any of the compressor rotors 11 (referred to as the compressor rotor 11 shown in the drawing and hereinafter simply referred to as the compressor rotor 11) is the inlet-side cross-sectional area (inlet cross-sectional area). ) Is configured to be larger than. The radial position (the radial position) of the tip 15t of each rotor blade 15 in the compressor rotor 11 is the same from the front edge 15a to the rear edge 15b. The radial position of the tip 15t of each rotor blade 15 in the compressor rotor 11 may change from the front edge 15a to the rear edge 15b.

  In the meridional section, the outer peripheral surface of the disk 13 in the compressor rotor 11 transitions from the convex first curve R1 to the concave second curve R2 in the downstream direction. Further, the upstream end (front end) of the outer peripheral surface of the disk 13 of the compressor rotor 11 is located radially inward from the extension line of the inner wall surface of the upstream compressor stator 7. The downstream end (rear end) of the outer peripheral surface of the disk 13 is located on the radially outer side of the extension line of the inner wall surface of the compressor stator 7 on the downstream side. The upstream end and the downstream end of the outer peripheral surface of the disk 13 of the compressor rotor 11 are on the extension line of the inner wall surface of the compressor stator 7 on the upstream side and on the extension line of the inner wall surface of the compressor stator 7 on the downstream side, respectively. It is desirable to be located.

  The inflection point IP that transitions from the first curved surface R1 to the second curved surface R2 is 2.8% of the cord length (−2.8% code) from the maximum blade thickness position TP of the moving blade 15 toward the upstream direction. And a point separated by 5.0% (5.0% code) of the cord length in the downstream direction from the maximum blade thickness position TP of the moving blade 15. The point at which the blade 15 is separated from the maximum blade thickness position TP in the upstream direction by 2.8% of the cord length is that the inflection point IP is the maximum blade thickness of the blade 15 as shown in FIG. This is because it has been found that the flow coefficient at the stall point becomes higher than the stall margin when the distance from the position TP exceeds 2.8% of the cord length in the upstream direction. On the other hand, the inflection point IP is a point that is separated from the maximum blade thickness position TP of the moving blade 15 by 5% of the cord length in the downstream direction. The inflection point IP is the maximum blade thickness position TP of the moving blade 15. When moving away from the downstream by more than 5% of the cord length, the axial flow velocity of the gas in the vicinity of the trailing edge 15b of the rotor blade 15 is rapidly decelerated, causing separation of the gas flow, and axial flow compression. This is because there is a concern that the pressure ratio of the machine 1 (pressure ratio of the compressor rotor 11) may decrease. The cord length refers to the length of a straight line connecting the leading edge 15a and the trailing edge 15b of the rotor blade 15.

  Then, the effect | action and effect of embodiment of this invention are demonstrated.

  A turbine is driven by the expansion of combustion gas from a combustor (not shown) in a jet engine to rotate a multi-stage turbine rotor, whereby a multi-stage compressor rotor 11 is rotated integrally with the multi-stage turbine rotor. Let Thereby, the multistage compressor rotor 11 and the multistage compressor stator 7 can cooperate with each other, and the gas taken into the gas flow path 5 can be compressed and conveyed in the axial direction.

  The outlet side cross-sectional area of the compressor rotor 11 is configured to be larger than the inlet side cross-sectional area. In the meridian cross section, the outer peripheral surface of the disk 13 in the compressor rotor 11 is convex in the downstream direction. Since the transition from the curve R1 to the concave second curve R2 is applied, when the above-described novel knowledge is applied, the outer peripheral surface of the disk 13 in the compressor rotor 11 is shown in the portion surrounded by the one-dot chain line in FIG. Compared with the case of an inclined straight line (see FIG. 3B), the pressures on the pressure surface 15p and the suction surface 15n on the leading edge 15a side of the moving blade 15 have a great influence on the stall of the axial compressor 1. The difference can be made smaller. At this time, the pressure difference between the pressure surface 15p and the suction surface 15n is widened at a portion away from the front edge 15a of the moving blade 15, but in the vicinity of the front edge 15a of the moving blade 15 surrounded by a one-dot chain line in FIG. The effect of reducing the pressure difference between the pressure surface 15p and the suction surface 15n is greater.

  Therefore, according to the embodiment of the present invention, the pressure difference between the pressure surface 15p and the suction surface 15n on the leading edge 15a side of the rotor blade 15 can be further reduced, so that the pressure surface 15p side to the suction surface 15n side of the rotor blade 15 can be reduced. As shown in FIG. 5, it was found that the operating range of the axial flow compressor 1 can be further expanded to the low flow rate side while reducing the stall flow of the axial flow compressor 1 as shown in FIG. In particular, the point where the inflection point IP is separated from the maximum blade thickness position TP of the moving blade 15 by 2.8% of the cord length in the upstream direction, and the cord in the downstream direction from the maximum blade thickness position of the moving blade 15. Since it exists between the points separated by 5.0% of the length, stall of the axial compressor 1 can be sufficiently suppressed while suppressing a decrease in the pressure ratio of the axial compressor 1.

  Here, as described above, FIG. 6 shows the compressor rotor (in other words, the novel compressor rotor shown in FIG. 3A) 11 and the comparative compressor rotor (FIG. 6) according to the embodiment of the present invention. 3 (b) is a diagram for explaining the relationship between the distance from the maximum blade thickness position TP to the inflection point IP of the moving blade 15 and the flow coefficient at the stall point for the comparative compressor rotor shown in FIG. This relationship is obtained by three-dimensional steady viscosity CFD analysis. Further, Example 1 in FIG. 5 is an analysis result in the case where the clearance between the moving blade 15 and the compressor case 3 is assumed as the design clearance in the compressor rotor 11 according to the embodiment of the present invention, Example 2 in FIG. 5 is an analysis result when the clearance between the moving blade 15 and the compressor case 3 is assumed to be twice the design clearance in the compressor rotor 11 according to the embodiment of the present invention. is there. Comparative Example 1 in FIG. 5 is an analysis result when the clearance between the moving blade and the compressor case is assumed as a design clearance in the comparative compressor rotor, and Comparative Example 2 in FIG. It is an analysis result at the time of assuming the clearance between a moving blade and a compressor case as a clearance twice as large as a design clearance in a compressor rotor for comparison. At this time, it can be seen that the stall point moves to the low flow rate side in any case, and the operating range is expanded to the low flow rate side.

  The present invention is not limited to the description of the above-described embodiment, but can be applied in various modes, such as applying the axial flow compressor 1 used as a component device of a gas turbine engine to other axial flow compressors. It can be implemented. The scope of rights encompassed by the present invention extends not only to the axial flow compressor 1 but also to a gas turbine engine equipped with the axial flow compressor 1.

R1 First curved surface R2 Second curved surface IP Inflection point TP Maximum blade thickness position 1 Axial flow compressor 3 Compressor case 5 Gas flow path 5s Wall surface 7 Compressor stator 9 Stator blade 11 Compressor rotor 13 Disc 15 Rotor blade 15p Positive Pressure surface 15n Negative pressure surface 15a Front edge 15b Rear edge 15t Tip

Claims (4)

An axial compressor that compresses and transports gas in the axial direction,
A cylindrical compressor case that extends in the axial direction and has an annular gas flow path formed inside;
A plurality of compressor stators provided in the compressor case along the axial direction and provided with a plurality of stationary blades disposed at equal intervals in the circumferential direction and located in the gas flow path;
A plurality of stages of compressor rotors are alternately provided in the compressor case along the axial direction, and the outer peripheral surface constitutes a part of the radially inner flow path surface of the gas flow path and around the axis. A rotatable disk, and a multi-stage compressor rotor provided with a plurality of moving blades provided at equal intervals on the outer peripheral surface of the disk and positioned in the gas flow path,
The outlet side cross-sectional area of any of the compressor rotors is configured to be larger than the inlet side cross-sectional area,
In the meridional section, the outer peripheral surface of the disk in any one of the compressor rotors transitions from a convex first curve to a concave second curve in the downstream direction, and the compressor rotor in any of the compressor rotors The upstream end of the outer peripheral surface of the disk is located radially inward of or extending from the extension line of the inner wall surface of the compressor stator on the upstream side, and downstream of the outer peripheral surface of the disk in any one of the compressor rotors The axial flow compressor is characterized in that the end is located on the radially outer side or on the extension line with respect to the extension line of the inner wall surface of the compressor stator on the downstream side.   The axial flow compressor according to claim 1, wherein a radial position of a tip of each rotor blade in any one of the compressor rotors is the same from a leading edge to a trailing edge.   The inflection point at which the first curve changes to the second curve includes a point away from the maximum blade thickness position of the moving blade by 2.8% of the cord length in the upstream direction, and the maximum blade of the moving blade. The axial flow compressor according to claim 1, wherein the axial flow compressor is present between the thickness position and a point separated by 5.0% of the cord length in the downstream direction.   A gas turbine engine comprising the axial flow compressor according to any one of claims 1 to 3.

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