JPH116445A - Gas turbine engine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a gas turbine engine to improve compressor performance and engine performance. SOLUTION: A heat insulation structure member 42 is inserted between a turbine nozzle plate 40 and a compressor diffuser plate 41 and fastened against each other by means of bolts, not shown. In the heat insulation structure member 42, recesses are respectively formed on the turbine nozzle plate 40 side and the compressor diffuser plate 41 side and a space is formed. Further, an annular fin 45 is attached to the outer peripheral part of the heat insulation structure member 42.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

[0001] The present invention relates to a gas turbine engine.

[0002]

2. Description of the Related Art Conventional gas turbine engines include:
For example, as disclosed in JP-A-2-238132, there has been a structure as shown in FIGS.

The configuration of this prior art is as follows. The gas turbine 1 comprises the basic components of a compressor 2, a combustor 3, and a turbine 4, and a turbine rotor 11 driven by hot gas and a centrifugal compressor impeller 12 for pressurizing cold gas are coaxial via a shaft 13. Connected, the shaft 13 drives a load via a pair of rolling bearings.

The high-temperature exhaust gas discharged from the turbine 4 is supplied from a turbine housing 26 via a diffuser to a recuperator (heat exchanger) 5 as indicated by an arrow in the drawing.
And absorbs the heat of the exhaust gas to reduce the fuel consumption rate. As shown in FIG. 11, the diffuser 25 is connected to the turbine housing 26 via the plenum chamber 27, and the joining flange 25A to the turbine housing 26 side is circular, and the plenum chamber 27
Are formed in a quadrangular shape.

In the recuperator 5, a flow path (not shown) for passing high-temperature exhaust gas sent from the turbine 4 is formed in the rotation direction of the shaft 13 so that the flow path is along the flow direction of the exhaust gas discharged from the turbine 4. Be placed. A flow path through which the low-temperature compressed air sent from the compressor 2 passes is formed in parallel with and opposed to the exhaust gas flow path via a heat transfer wall.

The recuperator 5 and the combustor 3 are arranged in parallel with each other, and a chamber 21 is provided to cover both the recuperator 5 and the combustor 3. Low-temperature compressed air discharged from the compressor 2 passes through the chamber 21 to the recuperator 5. Be guided.

[0007] Low-temperature compressed air pumped from the compressor 2 flows into the chamber 21 through an annular flow path 23 formed in the flange 22, as indicated by an arrow in the figure. The chamber 21 has a plurality of bolts 2
4 are fastened.

The recuperator 5 is formed with an inlet 5A for low-temperature compressed air that opens into the chamber 21, and the compressed air that flows in from the inlet 5A and is heated through the recuperator 5 is bent by the header 5C. After that, it flows into the combustor via the joint flange 5B.

Although the opposed recuperator 5 is used in the prior art, a cross-flow recuperator in which the exhaust gas and the low-temperature compressed air intersect at right angles to each other may be used. Is relatively large.

In the combustor 3, the fuel injected from the fuel injection valve 6 is burned to produce a high-temperature gas, and the combustion gas flowing out of the combustor 3 is supplied to a flow path defined between the plenum chamber 27 and the heat insulating material 33. After being bent at 28, it is guided to the turbine rotor 11 through a flow path 29 defined spirally between the turbine housing outer peripheral walls 34, expands, and applies a rotational force to the turbine rotor 11. I have.

The plenum chamber 27 includes a diffuser 2.
5 and a joining flange 27B for the combustor outer chamber 34, respectively. The heat insulating material 33 is provided to be joined to the flange 22 side, and insulates the combustion gas guided to the turbine rotor 11.

As shown in FIG. 12, a fuel injection valve 6 and a spark plug 7 are attached to a cap 35 attached to the end of the combustor 3, and the cap 35 has a flow path 36 formed in the hollow cap 35. In addition to cooling the fuel injection valve 6 and the spark plug 7, the low-temperature compressed gas taken out of the chamber 21 flows from the inlet 37 through a pipe and flows into the combustor outer chamber 34 from a plurality of outlets 38. , The combustor 3 is insulated.

[0013]

However, in such a conventional gas turbine engine, a part of the heat of the combustion gas in the turbine stage flows into the compressor stage, which is back-to-back, and the compressor stage is There is a problem that the performance of the step is reduced.

The present invention has been made in view of such conventional problems, and has as its object to provide a gas turbine engine capable of improving compressor performance and improving engine performance.

[0015]

SUMMARY OF THE INVENTION In order to solve the above-mentioned problems, the present invention provides a radial turbine rotor that is driven to rotate by exhaust gas and air sent to a combustor by a radial impeller disposed coaxially with the rotor. In the gas turbine engine obtained by compressing the gas turbine engine, radiating fins are provided in the outer radial direction on the outer peripheral portion of the heat insulating structural member between the compressor stage and the turbine stage, on the compressor side from the center position of the axial thickness, and the compressor diffuser is provided. It is structured to be exposed outside the outer diameter to the flow path downstream of the compressor stage.

Further, in the above-described gas turbine engine, a structure is provided in which a rectifying vane is provided at an outer peripheral end of the radiation fin.

Further, the gas turbine engine has a structure in which the radiation fins are bent in the direction of the turbine outlet at the maximum outer diameter position, and overlap with the outer wall of the turbine plenum in the axial direction.

[0018]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Hereinafter, embodiments of a gas turbine engine according to the present invention will be described in detail with reference to the accompanying drawings.

(First Embodiment) FIGS. 1 to 4 are views showing a first embodiment of a gas turbine engine according to the present invention. First, the structure will be described. As shown in FIG. 1, a heat insulating structural member 42 is inserted between a turbine nozzle plate 40 and a compressor diffuser plate 41 and fastened with bolts (not shown). As shown in FIG. 2 which is an enlarged view of a main part of FIG. 1, the heat insulating structural member 42 is provided with recesses on the turbine nozzle plate 40 side and the compressor diffuser plate 41 side, respectively, and forms spaces 43 and 44. ing. An annular fin 45 is attached to an outer peripheral portion 42a of the heat insulating structural member 42. The plan view is shown in FIG. Impeller 12
A diffuser with a vane 46 as shown in FIG. 4 which is a cross section AA in FIG.
The diameter of the outer circumference 45a at the tip of the previous annular fin 45 is larger than the diameter of the outer circumference 49 of the vane 46.
It projects into the compressor outlet channel 48. The axial position of the annular fin 45 is closer to the compressor diffuser plate 41 than the center position 50 of the axial thickness of the heat insulating structural member 42.

Next, the operation of the first embodiment will be described. The turbine nozzle plate 40 is provided for the turbine rotor 1.
The fuel gas is heated by the combustion gas flowing into the fuel cell 1 and has a high temperature.
When this heat flows into the compressor, it is heated in the middle of the compression stroke, which lowers the performance of the compressor.

In the first embodiment, a turbine nozzle plate 40 and a compressor diffuser plate 41 are fastened via a heat insulating structural member 42. Although the heat of the high-temperature turbine nozzle plate 40 flows into the heat-insulating structural member 42, the surface of the heat-insulating structural member 42 on the turbine side has a depression, and the solid contact area with the turbine nozzle plate 40 is equal to the surfaces 42 b and 42 c. It is made small and mostly faces the turbine nozzle plate 40 via a low thermal conductivity air layer in the space 43. As a result, the amount of heat flowing from the turbine nozzle plate 40 to the heat insulating structural member 42 is reduced.

The compressor-side surface of the heat insulating structural member 42 also has a recess, and the solid contact area with the compressor diffuser plate 41 is reduced to the surfaces 42d and 42e. It is opposed to the compressor diffuser plate 41 via a low-efficiency air layer. Thereby, the amount of heat flowing from the heat insulating structural member 42 to the compressor diffuser plate 41 is reduced.

Further, the outer peripheral portion 42 of the heat insulating structural member 42
The heat is radiated from the heat insulating structural member 42 to the low-temperature compressed air in the compressor outlet channel 48 by the annular fin 45 attached to the a, and the amount of heat flowing into the compressor diffuser plate 41 is reduced. Here, the air flowing out of the impeller 12 is decelerated and compressed while passing through the flow path 47 between the vanes 46 and flows into the compressor outlet flow path 48. The diameter of the outer circumference 45a at the tip of the annular fin 45 is
Since the diameter of the outer periphery 49 of the vane 46 is larger than the diameter of the outer periphery 49 and protrudes into the compressor outlet flow path 48, the annular fin 45 comes into contact with air having a high flow velocity in the flow path, so that the heat conductivity is increased and the heat insulation structure is increased. The amount of heat radiated from the member 42 to the low-temperature compressed air in the compressor outlet channel 48 increases,
The amount of heat flowing into the compressor diffuser plate 41 is further reduced.

The axial position of the annular fin 45 is closer to the compressor diffuser plate 41 than the central position 50 of the axial thickness of the heat insulating structural member 42.
The air layer between the turbine nozzle plate 40 and the plenum outer wall 51 and the annular fin 45 becomes thicker, and the annular fin 45
The amount of heat flowing into the air can be reduced. Therefore, compared with the case where the annular fin 45 is located on the turbine nozzle plate 40, it is possible to reduce a decrease in the turbine inlet temperature due to the heat escaping from the turbine.

(Second Embodiment) FIGS. 5 and 6 show a second embodiment of the gas turbine engine according to the present invention.

First, the structure will be described. In addition to the structure of the first embodiment, a rectifying vane 52 is attached to the outer periphery 45a of the annular fin 45 in the second embodiment. The axial position of the leading edge 52a of the straightening vane 52 is set at the center of the axial thickness of the vane 46 of the compressor diffuser,
The radial position is determined by experiment. In addition, the cross section of the rectifying vane 52 has an R shape. The annular fin 45 has a through hole 53 between the inner diameter of the rectifying vane 52 and the outer diameter 49 of the vane 46 of the compressor diffuser.
Is provided. The rectifying vane 52 is not limited to the case where the rectifying vane 52 is attached over the entire circumference, for example, as shown in FIG.
Only the lower 180 degrees may be attached. In this case, the circumferential range of the through hole 53 is also substantially the same as the range of the rectifying vane 52.

Next, the operation of the second embodiment will be described. Since the axial position of the leading edge 52a of the rectifying vane 52 attached to the outer periphery of the annular fin 45 is located at the center of the axial thickness of the compressor diffuser vane 46, the flow path 47 between the compressor diffuser vanes 46 The flow rate of the outflowing air is divided approximately by half with the flow straightening vane 52 interposed therebetween. The air 54 that has flowed into the turbine side of the leading edge of the straightening vane 52 flows along the vane, and flows downstream through the through hole 53 formed in the annular fin 45. As described above, it is generally known that when the flow direction changes by 90 degrees, when the rectifying vanes 52 are attached, pressure loss is reduced by suppressing separation. Further, since the heat from the annular fins 45 is radiated to the air from both surfaces of the rectifying vanes 52, the heat from the heat insulating structural member 42 to the low-temperature compressed air in the compressor outlet flow path 48 is higher than in the first embodiment. The amount of heat dissipated is increased, and the amount of heat flowing into the compressor diffuser plate 41 is further reduced.

(Third Embodiment) FIGS. 7 and 8 show a third embodiment of the gas turbine engine according to the present invention.

First, the structure will be described. In addition to the structure of the first embodiment, as shown in FIG.
Has a portion 58 that bends toward the turbine outlet at the maximum outer diameter position and overlaps with the turbine plenum outer wall 55 in the axial direction with a space 56 interposed therebetween. The cross section of the bent portion 57 has an R shape. The length of polymerization is determined by experiments.
The portion 58 to be superimposed may not only extend over the entire circumference, but also, for example, only at the lower 180 degrees as shown in FIG.

Next, the operation of the third embodiment will be described. In addition to the operation of the first embodiment, the air flowing out of the flow path 47 between the vanes 46 of the compressor diffuser flows downstream along the bend of the annular fin 45.
Since the cross section of the bent portion 57 has an R shape, the cross section is less likely to be peeled off than in the case of a right angle, and the pressure loss can be reduced. Since the air in the space 56 hardly flows, the heat transfer coefficient between the turbine plenum outer wall 55 and the overlapping portion 58 is low, and the amount of heat flowing from the plenum outer wall to the overlapping portion 58 is small. Therefore, compared with the case where there is no portion 58 to be polymerized,
A decrease in turbine inlet temperature due to heat escaping from the turbine to the low-temperature compressed air in the compressor outlet passage 48 can be reduced.

[0031]

As described in detail above, according to the present invention, the structure is constituted by (1) a radial turbine rotor which is driven to rotate by exhaust gas and a radial impeller which is arranged coaxially with the rotor. In a gas turbine engine formed by compressing air sent to a combustor, a radiation fin is provided on an outer peripheral portion of a heat insulating structural member between a compressor stage and a turbine stage, at a position closer to the compressor than a central position in an axial thickness, and And exposed to the flow path downstream of the compressor stage outside the outer diameter of the compressor diffuser. Or (2) in addition to the above (1), a rectifying vane is provided at the outer peripheral end of the radiation fin. Or (3) In addition to the above (1), the radiation fin is bent in the direction of the turbine outlet at the maximum outer diameter position and overlaps with the outer wall of the turbine plenum in the axial direction. effective.

(First Embodiment) The compressor performance is improved by dissipating heat from the heat insulating structural member 42 to the low-temperature compressed air in the compressor outlet flow path 48 and reducing the amount of heat flowing into the compressor diffuser plate 41. Improve engine performance.

(Second Embodiment) In addition to the effects of the first embodiment, the amount of heat radiation from the heat insulating structural member 42 to the low-temperature compressed air in the compressor outlet flow path 48 is increased to further improve the compressor performance. . Further, the engine performance is improved by reducing the pressure loss in the curved flow passage at the compressor outlet.

(Third Embodiment) In addition to the effects of the first embodiment, the decrease in turbine inlet temperature caused by heat escaping from the turbine to the low-temperature compressed air in the compressor outlet passage 48 is reduced. Improve engine performance.

[Brief description of the drawings]

FIG. 1 is a diagram showing a configuration of a first embodiment of a gas turbine engine according to the present invention.

FIG. 2 is an enlarged view of a main part of FIG.

FIG. 3 is a plan view of the annular fin and the heat insulating structural member according to the first embodiment.

FIG. 4 is a sectional view taken along line AA of FIG. 1;

FIG. 5 is a diagram illustrating a configuration of a second embodiment.

FIG. 6 is a plan view of an annular fin and a heat insulating structural member according to a second embodiment.

FIG. 7 is a diagram illustrating a configuration of a third embodiment.

FIG. 8 is a plan view of an annular fin and a heat insulating structural member according to a third embodiment.

FIG. 9 is a side sectional view of a prior art gas turbine.

FIG. 10 is a front view of a conventional gas turbine.

FIG. 11 is an exploded perspective view of a conventional gas turbine.

FIG. 12 is a sectional view of a main part of a conventional gas turbine.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 11 Turbine rotor 12 Impeller 40 Turbine nozzle plate 41 Compressor diffuser plate 42 Heat insulation structural member 42a Outer peripheral portion 42b, 42c, 42d, 42e of heat insulation structural member 42 Surface 43,44 Space 45 Annular fin 45a Outer periphery of annular fin 45 tip 46 Vane 47 flow path 48 compressor outlet flow path 49 outer periphery of vane 46 50 central position of axial thickness of heat insulation structural member 42 51 plenum outer wall 52 rectifying vane 52a leading edge 53 of rectifying vane 52 53 through hole 54 air 55 turbine plenum outer wall 56 space 57 Bent part 58 Overlapping part

Claims (3)

[Claims] 1. A gas turbine engine comprising: a radial turbine rotor rotatably driven by exhaust gas; and a radial impeller arranged coaxially with the rotor to compress air sent to a combustor. Radiation fins are provided on the outer peripheral portion of the heat insulating structural member between the central position of the thickness in the axial direction and on the compressor side in the outer diameter direction, and are exposed outside the outer diameter of the compressor diffuser and downstream of the compressor stage. A gas turbine engine characterized by the above. 2. The gas turbine engine according to claim 1, wherein a rectifying vane is provided at an outer peripheral end of said radiating fin. 3. The gas turbine engine according to claim 1, wherein the radiating fins are bent toward a turbine outlet at a maximum outer diameter position, and overlap with a turbine plenum outer wall in an axial direction. Characteristic gas turbine engine.