JP2005194923A - Compressor moving blade - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a compressor moving blade of a gas turbine engine capable of suppressing vibration by increasing aerodynamic performance without reducing clearance at the tip of the compressor moving blade. <P>SOLUTION: In this compressor moving blade 51 of the gas turbine engine 1, a projection 53 projected in the blade thickness direction of the compressor moving blade 51 and extending in the chord length direction of the compressor moving blade 51 is formed on the abdominal side 51E of the compressor moving blade 51 near the tip part 51B of the compressor moving blade 51 in the span length direction. <P>COPYRIGHT: (C)2005,JPO&NCIPI

Description

  The present invention relates to a compressor rotor blade of a gas turbine engine, and more particularly to a blade provided with a protrusion on the blade surface.

  Panel mode vibration (vibration mode in which the phase of vibration differs in the code length direction of the blade); blade span length generated in the compressor blade (gas turbine axial compressor) of aerospace and industrial gas turbine engines As a technology to reduce the vibration mode with nodes extending in the direction), conventionally, by changing the blade length and blade thickness, the blade shape is changed to avoid resonance and vibration Techniques for reducing stress are known.

  On the other hand, in order to improve the aerodynamic performance (for example, compression efficiency) of the rotor blades of the compressor, particularly the rotor blades downstream of the compressor, the leakage flow of fluid such as air between the blade tips and the engine case is reduced. As a technology, there is known a flare blade that reduces the clearance at the tip of the moving blade (the distance between the tip of the moving blade and the engine case) and increases the cord length at the tip of the moving blade (for example, Patent Document 1). reference).

Note that the cord length may be increased not only at the tip of the rotor blade but also over the entire length in the span length direction of the rotor blade.
JP 09-032501 A

  By the way, if the blade length is shortened and the blade thickness is increased to avoid resonance in the panel mode and reduce the vibration stress, the aerodynamic performance is reduced by the shape change of the blade. There's a problem.

  In other words, when the blade thickness of the blade is increased, the cross-sectional area facing the flow increases and the resistance increases, the curvature of the blade surface increases, and the flow becomes easy to peel off. The length of the blade tip clearance in the front-rear direction of the compressor is shortened, the flow of air leaking through the clearance of the blade tip increases, and the pressure loss increases.

  On the other hand, it is conceivable to reduce the clearance at the tip of the rotor blade without changing the shape of the rotor blade, but if the clearance at the tip of the rotor blade is reduced, the environment changes during operation in the transient state of the gas turbine engine, Due to the change over time due to long-term operation, the clearance at the tip of the moving blade is lost, and the turbine case may be rubbed with the moving blade, so it is difficult to reduce the clearance at the tip of the moving blade.

  The present invention has been made in view of the above problems, and provides a compressor blade of a gas turbine engine that can improve aerodynamic performance and suppress vibration without reducing the clearance at the tip of the compressor blade. The purpose is to provide.

  According to the first aspect of the present invention, in the compressor rotor blade of the gas turbine engine, the blade thickness of the compressor rotor blade is set on the ventral side of the compressor rotor blade in the vicinity of the tip in the span length direction of the compressor rotor blade. A compressor blade having a protrusion protruding in the direction and extending long in the cord length direction of the compressor blade.

  According to a second aspect of the present invention, in the compressor blade according to the first aspect, the protrusion height of the protrusion is 50% to 80% of the maximum blade thickness of the compressor blade. The distance between the center position of the protrusion in the span length direction and the base end of the span of the compressor blade is 85% to 90% with respect to the entire span of the compressor blade, The compressor rotor blade has a width in the span length direction of the compressor rotor blade of 5% to 10% with respect to the span overall length of the compressor rotor blade.

  According to the present invention, it is possible to provide a compressor blade of a gas turbine engine that improves aerodynamic performance and suppresses vibration without reducing the clearance at the tip of the compressor blade.

  FIG. 1 is a cross-sectional view showing a schematic configuration of a gas turbine engine 1 provided with a compressor blade 51 according to an embodiment of the present invention.

  The gas turbine engine 1 is used for an aircraft as a jet engine, for example, and is an engine that obtains propulsive force or rotational force by ejecting high-temperature and high-pressure combustion gas.

  The gas turbine engine 1 includes an engine outer cylinder 3 and a hollow engine case 5 that is provided on the inner side of the engine outer cylinder 3 and is substantially concentric with the engine outer cylinder 3 as a base. An annular engine flow path 7 is formed inside the engine case 5, and an annular bypass flow path 9 is formed between the engine outer cylinder 3 and the engine case 5.

  A front support frame 11A is provided integrally with the engine case 5 at an inner front portion (upstream portion as viewed from the gas flow direction) of the engine case 5 with an engine flow path 7 therebetween. 5, a rear support frame 11B is provided integrally with the engine case 5 with an engine flow path 7 therebetween. The front support frame 11A and the rear support frame 11B are connected to a low-pressure turbine shaft via bearings. The front support frame 11A and the rear support frame 11B support the hollow high-pressure turbine shaft 15 so as to be rotatable and concentric with the low-pressure turbine shaft 13 via bearings. ing.

  A fan 17 that feeds air into the engine flow path 7 and the bypass flow path 9 is provided on the front end side of the low-pressure turbine shaft 13.

  A low-pressure compressor 19 is provided on the upstream side of the engine flow path 7. The low-pressure compressor 19 sends air to the downstream side (downstream side in the gas flow direction, right side in FIG. 1) while compressing the air at low pressure.

  The low-pressure compressor 19 is provided on the downstream side of the fan 17 with an annular blade support member 21 provided integrally with the low-pressure turbine shaft 13 and on the outer periphery of the blade support member 21 along the engine flow path 7. A plurality of low-pressure compression blade arrays 23 and a plurality of low-pressure compression blade arrays 25 provided alternately with the plurality of low-pressure compression blade arrays 23 along the engine flow path 7 inside the engine case 5. It has.

  A high-pressure compressor 27 is provided on the downstream side of the low-pressure compressor 19 in the engine flow path 7 so that the high-pressure compressor 27 sends the air compressed at a low pressure by the low-pressure compressor 19 to the downstream side while compressing the air at high pressure. It has become.

  The high-pressure compressor 27 includes a plurality of stages of high-pressure compressor blades 29 provided on the high-pressure turbine shaft 15 along the engine flow path 7 and a plurality of stages along the engine flow path 7 inside the engine case 5. A plurality of high-pressure compression stator blade rows 31 provided alternately with the high-pressure compression blade row 29;

  An annular combustion chamber 33 is provided on the downstream side of the high-pressure compressor 27 in the engine flow path 7, and the combustion chamber 33 burns fuel in compressed air to generate high-temperature and high-pressure combustion gas. .

  A high-pressure turbine 35 is provided on the downstream side of the combustion chamber 33 in the engine flow path 7, and the high-pressure turbine 35 obtains rotational force by the expansion of high-temperature and high-pressure combustion gas from the combustion chamber 33, The shaft 15 is rotationally driven.

  The high-pressure turbine 35 is provided on the high-pressure turbine shaft 15 along the engine flow path 7 and is rotated by high-temperature and high-pressure combustion gas. A plurality of high-pressure turbine rotor blade rows 37 and high-pressure turbine stationary blade rows 39 provided alternately are provided.

  A low-pressure turbine 41 is provided downstream of the high-pressure turbine 35 in the engine flow path 7, and the low-pressure turbine 41 obtains a rotational force by the expansion of high-temperature and high-pressure combustion gas from the combustion chamber 33, and the low-pressure turbine shaft 13 is driven to rotate.

  The low-pressure turbine 41 is provided on the low-pressure turbine shaft 13 along the engine flow path 7 and is rotated by high-temperature and high-pressure combustion gas. 7, the plurality of low-pressure turbine rotor blade rows 43 and the plurality of low-pressure turbine stationary blade rows 45 provided alternately are provided.

  Next, the compressor blade 51 constituting the high pressure compressor blade row 29 of the high pressure compressor (high pressure compressor) 27 will be described.

  FIG. 2 is an arrow view showing a IIA-IIB cross section in FIG.

  3 is a diagram showing an arrow IIIA-IIIB in FIG. 2, FIG. 4 is an enlarged view of a IV portion in FIG. 3, and FIG. 5 is an enlarged view of a V portion in FIG.

  FIG. 6 is a view showing a cross section taken along the line VIA-VIB in FIG.

  The compressor blades 51 are held by the high-pressure turbine shaft 15 on the base end portion 51 </ b> A side, and a plurality of compressor blades 51 holding the base end portion 51 </ b> A side are arranged in the circumferential direction of the high-pressure turbine shaft 15. Has been.

  A slight clearance T1 is formed between the tip 51B of the compressor rotor blade 51 and the engine case 5 in order to prevent the compressor rotor blade 51 and the engine case 5 from interfering with each other. (See FIG. 2).

  In the blade portion 51C of the compressor blade 51, the radial length L1 (see FIG. 2) of the high-pressure turbine shaft 15 is the span length, the width L3 of the blade portion 51C is the cord length, and the blade portion 51C The thickness L5 (see FIG. 3) is the blade thickness. In FIG. 3, the compressor rotor blade 51 rotates in the direction indicated by the arrow AR1. In the blade portion 51C of the compressor moving blade 51, the front side is a back 51D and the rear side is a belly 51E.

  Protrusions (ribs) 53 that protrude in the direction of the blade thickness of the compressor blade 51 are formed on the ventral side 51E of the compressor blade 51 in the vicinity of the tip 51B in the span length direction of the compressor blade 51. Is provided.

  The protrusion 53 is, for example, a streamline of the tip 51B of the compressor blade 51 (when the compressor blade 51 is viewed from a direction perpendicular to the span length direction and the cord length direction, the compressor blade 51 (Line formed at the tip end surface) 51F and extending substantially in the cord length direction of the compressor rotor blade 51 (blade portion 51C) substantially in parallel with the 51F.

  The protrusion height L7 of the protrusion 53 is 50% to 80% (preferably 60% to 70%, more preferably 65%) of the maximum blade thickness of the compressor rotor blade 51 (see FIG. 3). ).

  Further, in the projection 53, a distance L9 between the position of the center 53 of the projection 53 in the span length direction of the compressor blade 51 and the base end 51G of the compressor blade 51 in the span length direction is the compressor 53. It is 85% to 90% (preferably 88%) with respect to the span total length L1 of the moving blade 51 (see FIG. 2).

  Further, in the protrusion 53, the width L11 of the compressor rotor blade 51 in the span length direction is 5% to 10% (preferably 8%) with respect to the span total length L1 of the compressor rotor blade 51 ( (See FIG. 2).

  In the protrusion 53, the length L13 of the compressor blade 51 in the cord length direction is about 90% of the cord length L3 of the compressor blade, and the fluid is efficiently fed to the ventral side 51E. Therefore, no protrusion is formed on a portion 51H corresponding to 5% of the cord length L3 from the front end in the cord length direction of the compressor rotor blade 51 toward the rear side (see FIG. 3).

  Further, in order to efficiently flow out the fluid introduced to the ventral side 51E to the rear side, it corresponds to a length of 5% of the cord length from the rear end in the cord length direction of the compressor blade 51 toward the front side. The projection 51I is not formed on the portion 51I to be operated (see FIG. 3).

  Furthermore, the skirt part (base end part) 51J of the protrusion 53 is interpolated by, for example, a curved surface formed by a cubic curve, and is formed smoothly. The protrusion 53 rises smoothly from the surface on which the protrusion on the ventral side 51E is not provided. By being formed in this way, the occurrence of stress concentration can be suppressed and the fluid can flow smoothly.

  Further, the edge on the tip 51K side of the protrusion 53 is also smoothly formed by a curved surface formed by, for example, a cubic curve so that the fluid flows smoothly (see FIGS. 3 to 6).

  Next, the vibration of the compressor blade 51 when the gas turbine engine 1 employing the compressor blade 51 is operated will be described.

  FIG. 7 is a diagram for explaining the vibration of the compressor blades 51 when the gas turbine engine 1 employing the compressor blades 51 is operated.

  The horizontal axis in FIG. 7 indicates the rotational speed of the rotor (high pressure turbine shaft) 15 in which the compressor blades 51 are used, and the rotational speed increases toward the right of the horizontal axis. The vertical axis indicates the natural frequency of the compressor rotor blade 51, and the frequency increases as it goes upward.

  The graph G1 is a graph showing the vibration applied to the compressor rotor blade 51 depending on the number of stator blades in the stator blade row 31 and the rotational speed of the rotor 15.

  Graph G3A is a graph showing the natural frequency of the compressor blade 51 in the third-order panel mode, and graph G3B is a conventional compressor blade in the third-order panel mode (compressor blade not provided with the protrusion 53). ) Is a graph showing the natural frequency.

  Graph G4A is a graph showing the natural frequency of the compressor blade 51 in the fourth-order panel mode, and graph G4B is a conventional compressor blade in the fourth-order panel mode (compressor blade not provided with the protrusion 53). ) Is a graph showing the natural frequency.

  Although there are other vibration modes generated in the compressor rotor blade 51, the gas turbine engine 1 using the compressor rotor blade 51 having a longer cord length in order to reduce the leakage of compressed air is used as the idling speed. When actually operating between the maximum number of rotations, the problems of the third-order panel mode and the fourth-order panel mode are particularly problematic. Therefore, in this specification, the vibration of the third-order panel mode and the vibration of the fourth-order panel mode will be described.

  Here, at the point P3A where the graph G1 and the graph G3A intersect each other, the resonance vibration in the third-order panel mode occurs in the compressor blade 51, and at the point P3B where the graph G1 and the graph G3B intersect each other, Resonant vibration in the third-order panel mode is generated in the conventional compressor blade.

  In the gas turbine engine 1, by providing the protrusion 53, the vibration level (for example, amplitude) of the compressor blade 51 can be reduced to about a half of that of the conventional compressor blade.

  Further, in the fourth-order panel mode, it is possible to avoid the occurrence of resonance in the operation region of the gas turbine engine 1 by raising the resonance frequency of the compressor blade 51.

  That is, by raising the graph G4A to the upper side of the graph G4B, the graph G1 and the graph G4A can be prevented from crossing each other within the range of the idle speed and the maximum speed.

  Here, the vibration in the panel mode will be described.

  FIG. 8 is a diagram for explaining the vibration of the third-order panel mode in the compressor rotor blade 51, and FIG. 9 is a diagram for explaining the vibration of the fourth-order panel mode in the compressor rotor blade 51.

  8 and 9 are views of the compressor blades as viewed from the direction of the arrow AR3 shown in FIG.

  In the vibration of the third-order panel mode, the three nodes CL5 extending in the span length direction appear at intervals in the cord length direction, and the compressor blades 51 vibrate. That is, vibrations having different phases are generated at the “+” and “−” portions shown in FIG.

  In the fourth-order panel mode vibration, four nodes CL7 extending in the span length direction appear at intervals in the cord length direction, and the compressor blades 51 vibrate. That is, vibrations having different phases are generated at the “+” and “−” portions shown in FIG.

  Next, the aerodynamic performance of the compressor 27 using the compressor rotor blade 51 will be described.

  FIG. 10 is an enlarged view of a portion X in FIG. 1, showing a flow of air in a compressor employing a conventional compressor blade, and FIG. 11 is an enlarged view of a portion X in FIG. It is a figure which shows the flow of the air of the compressor in which the compressor rotor blade 51 is employ | adopted.

  In addition, the arrow shown in FIG.10 and FIG.11 shows the direction and flow velocity which air flows, and the flow velocity becomes so high that the arrow is long.

  By providing the protrusions 53 on the compressor blades 51, the air flow in the XI section (the vicinity of the protrusions 53) shown in FIG. 11 is different from that shown in FIG.

  That is, in the XI part shown in FIG. 11, the flow F1 toward the engine case 5 side and the flow F3 toward the rotor 15 side are generated. In addition, on the rear side of the protrusion 53, the air flow is fast (see F5).

  FIG. 12 is a diagram showing the air velocity in the section IIA-IIB shown in FIG. 1, and is a diagram showing the air velocity of a compressor employing a conventional compressor blade, and FIG. 13 is shown in FIG. It is a figure which shows the speed of the air in the IIA-IIB cross section shown, and is a figure which shows the speed of the air of the compressor in which the compressor moving blade 51 is employ | adopted.

  Here, the part PT1 shown in FIGS. 12 and 13 has a low air flow rate, and the air flow rate increases toward the part PT3 and part PT5. Part PT1 corresponds to a region of air leaking through the clearance from the ventral side to the back side of compressor blade 51.

  As can be seen from a comparison between FIG. 12 and FIG. 13, the region where the flow velocity generated on the engine case 5 side is lower in the case where the compressor blade 51 is used than in the case where the conventional compressor blade is used. Although it is small, and not shown in FIGS. 12 and 13, the minimum value of the flow velocity also appears in the case of using the conventional compressor blade. This indicates that the amount of leaked air has decreased.

  Here, the reason why the aerodynamic performance, that is, the leakage is reduced in the compressor using the compressor blade 51 as compared with the conventional compressor blade is described.

  By providing the protrusion 53, the flow F1 from the protrusion 53 toward the engine case 5 side on the ventral side 51E of the compressor rotor blade 51 and the base end portion 51A side of the compressor rotor blade 51 from the protrusion 53 are provided. A flowing flow F3 is generated (see FIG. 11).

  Of these flows F1 and F3, the flow F1 is directed from the protrusion 53 toward the engine case 5 and is downstream of the clearance between the compressor blades 51 and the engine case 5 (the ventral side). The air flow rate of the air is increased compared with the case where no protrusion is present, and the pressure on the downstream side (ventral side) is lower than the case where the protrusion is not present. The amount of air leaking through the clearance T1 from the ventral side 51E to the back side 51D of the wing 51 is reduced as compared with the case where the protrusion 53 is not present.

  Further, the flow F3 from the protrusion 53 toward the proximal end portion 51A side of the compressor blade 51 causes the centrifugal blade to move from the proximal end portion 51A side to the distal end portion 51B side by the centrifugal force, that is, the rotor 15 side. Therefore, the flow of air toward the clearance between the compressor blades 51 and the engine case 5 is offset and lessened, and the amount of air leaking from the clearance is smaller than when no protrusion is present.

  Further, since the protrusion 53 is provided, the air flow path becomes narrower in the portion where the protrusion 53 is provided than in the case where the protrusion is not provided. Since the flow velocity is increased and the flow velocity is increased, the pressure on the downstream side (ventral side) is lower than that in the case where no protrusion is present, and therefore the clearance from the ventral side 51E of the compressor blade 51 to the back side 51D is reduced. The amount of air leaking through T1 is less than when no protrusion is present.

  According to the compressor rotor blade 51, the compressor rotor blade 51 protrudes in the direction of the blade thickness of the compressor rotor blade 51 on the ventral side 51E of the compressor rotor blade 51 in the vicinity of the tip in the span length direction of the compressor rotor blade 51. Since the protrusion 53 extending in the cord length direction of the moving blade 51 is provided, the rigidity against vibration of the panel mode can be increased without shortening the cord length of the compressor moving blade 51, as described above. The vibration of the compressor blades 51 during the operation of the compressor provided with the compressor blades 51 can be suppressed.

  Further, according to the compressor blade 51, since the cord length is not shortened, the compressor blade 51 passes through the clearance without reducing the clearance between the compressor blade 51 and the engine case 5. The amount of air leaking from the ventral side 51E to the back side 51D can be reduced, and the protrusion 53 is provided in the vicinity of the tip of the compressor rotor blade 51 in the span length direction. The amount of air that passes through the clearance and leaks from the ventral side 51E of the compressor blade 51 to the back side 51D can be further reduced, and the aerodynamic performance of the compressor provided with the compressor blade 51 can be improved. it can.

  That is, according to the compressor rotor blade 51, vibration can be suppressed and aerodynamic performance can be improved.

  Further, according to the compressor rotor blade 51, the protrusion 53 is provided on the ventral side 51E of the compressor rotor blade 51. Therefore, when the compressor provided with the compressor rotor blade 51 is in operation (of the gas turbine engine 1). During operation, air is difficult to separate from the compressor rotor blade 51 on the back side 51D. Since the ventral side 51E is a side where air is compressed, even if the protrusion 53 exists, the protrusion 53 hardly causes air to peel off from the blade surface of the compressor rotor blade 51.

  Even when the protrusion 53 is provided on the back side 51D, it is possible to suppress vibration during operation of the compressor provided with the compressor blade 51.

  In the above-described embodiment, the description has been given by taking the example of the moving blade used in the compressor of the two-shaft gas turbine engine having the high-pressure shaft and the low-pressure shaft. The compressor blade according to the above embodiment can also be applied to the blade.

1 is a cross-sectional view showing a schematic configuration of a gas turbine engine provided with compressor rotor blades according to an embodiment of the present invention. It is an arrow line view which shows the IIA-IIB cross section in FIG. 1, and is the figure which expanded a part of high-pressure compression blade cascade. It is a figure which shows the IIIA-IIIB arrow in FIG. It is an enlarged view of the IV section in FIG. It is an enlarged view of the V section in FIG. It is a figure which shows the VIA-VIB cross section in FIG. 3, and is a figure which shows the cross section by the side of the front-end | tip part of a compressor moving blade. It is a figure explaining the vibration of a compressor moving blade at the time of driving | operating the gas turbine engine by which the compressor moving blade is employ | adopted. It is a figure explaining the vibration of the 3rd panel mode in a compressor blade. It is a figure explaining the vibration of the 4th panel mode in a compressor blade. It is an enlarged view of the X section in Drawing 1, and is a figure showing the flow of the air of the compressor in which the conventional compressor rotor blade is adopted. It is an enlarged view of the X section in Drawing 1, and is a figure showing the flow of the air of the compressor in which the compressor blade is adopted. It is a figure which shows the speed of the air in the IIA-IIB cross section shown in FIG. 1, and is a figure which shows the speed of the air of the compressor in which the conventional compressor rotor blade is employ | adopted. It is a figure which shows the speed of the air in the IIA-IIB cross section shown in FIG. 1, and is a figure which shows the speed of the air of the compressor by which the compressor blade is employ | adopted.

Explanation of symbols

1 Gas Turbine Engine 51 Compressor Blade 51B Tip 53 Projection

Claims (2)

In compressor blades of gas turbine engines,
Protruding in the direction of the blade thickness of the compressor blade and extending in the cord length direction of the compressor blade on the ventral side of the compressor blade near the tip of the compressor blade in the span length direction Compressor blades characterized by that. The compressor blade according to claim 1, wherein
The protrusion height of the protrusion is 50% to 80% of the maximum blade thickness of the compressor blade,
In the projection, the distance between the center position of the projection in the span length direction and the base end of the span of the compressor blade is 85% to 90% with respect to the span length of the compressor blade. ,
The compressor blade according to claim 1, wherein a width of the compressor blade in the span length direction is 5% to 10% with respect to a span length of the compressor blade.

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