CN212774523U - Interstage sealing structure of radial flow type gas turbine - Google Patents
Interstage sealing structure of radial flow type gas turbine Download PDFInfo
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- CN212774523U CN212774523U CN202021966378.4U CN202021966378U CN212774523U CN 212774523 U CN212774523 U CN 212774523U CN 202021966378 U CN202021966378 U CN 202021966378U CN 212774523 U CN212774523 U CN 212774523U
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Abstract
The utility model discloses an interstage sealing structure of a radial flow type gas turbine.A far axial core part of a sealing disk of a compressor is fixed on the rear side surface of a front end disk of a compression diffuser, which is close to the axial core part; the nozzle ring blades are uniformly arranged along the circumferential direction; cold air through holes penetrating through the nozzle ring front disc, the nozzle ring blades and the nozzle ring rear disc are formed corresponding to the nozzle ring blades respectively; the rear side of the nozzle ring rear disc is coaxial with the cold air through hole and is axially fixed with cooling rings at intervals; a circular sealing push ring is formed at the front side of the front disc of the nozzle ring near the axis center; the back side surface of the front end wall of the sealing disk of the compressor is provided with a radial sealing ring corresponding to the sealing push ring; the seal push ring is press-fit to the radial seal ring. The utility model discloses a radial flow formula gas turbine interstage seal structure makes gas turbine interstage sealing performance prediction simple, and manufacturing process requires lowly to radial flow formula gas turbine interstage sealed reliability and economic nature have been improved.
Description
Technical Field
The utility model relates to a gas turbine, in particular to radial flow type gas turbine interstage seal structure.
Background
With the continuous improvement of the performance requirements of modern gas turbines, sealing is an important research direction of the gas turbines, and the excellent sealing structure plays an important role in improving the efficiency of the gas turbines and ensuring the normal operation of the gas turbines. The sealing structure is an important component of the gas turbine, the sealing performance of the sealing structure plays a key role in the performance of the gas turbine, and particularly, the gas path sealing directly influences the improvement of the supercharging ratio and the turbine efficiency of the gas turbine.
The present common structural forms of the interstage sealing device of the radial flow type gas turbine are piston ring sealing and elastic ring sealing, and the interstage sealing device has different advantages and disadvantages due to the adoption of different sealing principles.
In a thermal state, the piston ring sealing structure utilizes the fact that a piston ring is heated to expand and is tightly attached to the surfaces of other parts, so that interstage sealing with high pressure difference is achieved, but two difficulties exist: firstly, how to ensure that the expanded piston ring can be tightly attached to the surface of a part without generating a large gap in a thermal state; complex and accurate calculation is required to meet the condition, and if the boundary condition has errors, gaps may occur in the actual structure; secondly, when the piston ring works, gas leakage occurs, especially when the pressure difference between two sides is large, the leakage amount is increased, so that the efficiency of the gas turbine is greatly influenced, and how to accurately control the leakage amount of the gas is a difficult point for ensuring the normal work of the gas turbine.
The elastic ring seal is another common structural form of the prior interstage seal of the radial-flow gas turbine, when in use, the elastic ring seal is arranged between a sealing disc and a nozzle ring of a compressor, and two ends of the elastic ring structure are tightly pressed on the sealing disc side and the nozzle ring side of the compressor by thermal expansion under a thermal state, but the elastic ring seal has two defects: firstly, the elastic ring has larger stress in a hot state, the material generally enters yielding, and the fatigue life and the creep life are shorter due to higher working temperature; and secondly, the elastic ring realizes air path sealing by means of contact between the side plane and the planes of other parts, and because the contact surface is large, the requirement on the manufacturing process precision is high, and the sealing performance can be seriously influenced by poor surface roughness and manufacturing tolerance.
SUMMERY OF THE UTILITY MODEL
The to-be-solved technical problem of the utility model is to provide a radial flow formula gas turbine interstage seal structure, make gas turbine interstage sealing performance prediction simple, manufacturing process requires lowly to radial flow formula gas turbine interstage sealed reliability and economic nature have been improved.
In order to solve the technical problem, the nozzle ring 1, the compressor sealing disc 2, the compressed air diffuser 7 and the cooling ring 15 of the interstage seal structure of the radial flow type gas turbine provided by the utility model are all annular structural members;
the nozzle ring 1, the compressor sealing disk 2, the compression diffuser 7, the cooling ring 15 and the annular combustion chamber 4 are coaxially assembled together;
the far axial center part of the sealing disk 2 of the compressor is fixed on the rear side surface of the front end disk 70 close to the axial center part of the air diffuser 7;
the nozzle ring 1 comprises a nozzle ring front disc 11, nozzle ring blades 12 and a nozzle ring rear disc 13;
the plurality of nozzle ring vanes 12 are uniformly arranged in the circumferential direction;
the front end of each nozzle ring blade 12 is attached and fixed to a nozzle ring front disc 11, and the rear end is attached and fixed to a nozzle ring rear disc 13;
cold air through holes 14 penetrating through the nozzle ring front disc 11, the nozzle ring blades 12 and the nozzle ring rear disc 13 are formed corresponding to the nozzle ring blades 12;
a cooling ring 15 is fixed at the rear side of the nozzle ring rear disc 13 corresponding to the cold air through hole 14; the cooling ring 15 is coaxial with and axially spaced from the nozzle ring rear disk 13;
the far axial center end of the front disc 11 of the nozzle ring is fixed on the back side surface of the far axial center end of the front disc 70 of the air pressure diffuser 7;
a circular sealing push ring 112 is formed on the front side of the near axle center end of the nozzle ring front disc 11;
a circular groove is formed on the back side surface of the front end wall 21 of the compressor sealing disc 2 corresponding to the sealing push ring 112;
the front end of the radial sealing ring 3 is fixedly embedded into the annular groove;
the sealing push ring 112 is pressed on the rear end of the radial sealing ring 3.
Preferably, the radial sealing ring 3 is an E-ring.
Preferably, the side of the nozzle ring front plate 11 away from the axial center is connected with the annular combustion chamber outer ring 41 of the annular combustion chamber 4 in a sealing way through a piston ring 5.
Preferably, the front end wall 21 of the compressor sealing disk 2 is formed with a limit ring 22 far from the rear side of the axial center part;
the seal pusher ring 112 is located on the proximal side of the stop collar 22.
Preferably, the far axial center of the compressor sealing disk 2 is fixed to the near axial center rear side of the front end disk 70 of the compressor diffuser 7 by bolts.
Preferably, N diffuser positioning bosses 72 are uniformly formed on the rear side surface of the far-shaft center part of the front end disc 70 of the air compression diffuser 7 along the circumferential direction, wherein N is an integer greater than 2;
n nozzle ring front disc positioning bosses 111 are uniformly formed on the front side of the far shaft end of the nozzle ring front disc 11 along the circumferential direction;
each diffuser positioning boss 72 is fixedly connected with the corresponding nozzle ring front disc positioning boss 111 through a positioning pin 114.
Preferably, N is 4, 6, 7 or 9.
Preferably, the compressor diffuser 7 is fixed inside the diffuser casing 81;
a radial cold air flow channel is formed between the front end disc 70 of the air compressor diffuser 7 and the diffuser casing 81;
a far shaft annular wall 23 and a near shaft annular wall 24 which are coaxial are formed on the rear side of the front end wall 21 of the compressor sealing disk 2;
a circle of heat insulation tiles 61 are arranged between the rear ends of the far collar wall 23 and the near collar wall 24;
the front end wall 21 of the compressor sealing disk 2, the far collar wall 23, the near collar wall 24 and the heat insulation tile 61 form an inter-stage sealing inner cavity 20 together.
Preferably, the heat insulation tile 61 is in a fan ring shape;
two wings of the heat insulation tile 61 are respectively provided with a groove and a convex rib;
the heat insulation tiles 61 are matched with the convex ribs through the grooves and are connected end to form a complete heat insulation tile sealing ring.
Preferably, the far axial end of the heat insulation tile 61 is fixedly connected to the far collar wall 23 of the compressor sealing disc 2 through a heat insulation tile positioning pin 62;
a heat insulation tile installation groove is formed in the rear end of the far shaft side of the near shaft annular wall 24 of the compressor sealing disk 2 along the circumferential direction;
the end of the heat insulation tile 61 close to the shaft is inserted into the heat insulation tile installation groove and is in sealing fit with the annular wall 24 close to the shaft of the compressor sealing disc 2.
Preferably, the cold gas through holes 14 in the nozzle ring vanes 12 pass through the geometric center of the nozzle ring vanes 12 in the axial direction.
Preferably, the nozzle ring vanes 12 are hollow vanes.
The utility model discloses a radial flow formula gas turbine interstage seal structure, radial flow formula gas turbine is out of work the time, sealed ejector ring 112 pressfitting is to radial seal ring 3 rear ends, realizes compressing air diffuser 7 and separates with the radial seal of 1 steam jet passage of nozzle ring. When the radial-flow gas turbine works, the cold air through holes 14 of the nozzle ring 1, which penetrate through the nozzle ring front disc 11, the nozzle ring blades 12 and the nozzle ring rear disc 13, guide high-pressure cold air to the rear side of the nozzle ring rear disc 13 and cool the rear side of the nozzle ring through the cooling ring 15, so that the hot position of the nozzle ring front disc sealing push ring 112 can seal and separate high-pressure compressor outlet gas and low-pressure turbine inlet gas, the gas turbine can efficiently and reliably run, the defects of difficult piston ring sealing performance prediction and high requirement on an elastic ring manufacturing process are overcome, and the reliability and the economical efficiency of the interstage seal of the radial-flow gas turbine are improved.
Drawings
In order to more clearly illustrate the technical solution of the present invention, the drawings required for the present invention are briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
Fig. 1 is an axial symmetry sectional view of an embodiment of the interstage seal structure of the radial flow gas turbine of the invention.
Fig. 2 is a partially enlarged view of fig. 1.
Description of reference numerals:
1 a nozzle ring; 2, sealing a disc of the compressor; 21 a front end wall; 22 a stop collar; 23 a distal collar wall; 24 a proximal annular wall; 7 a compressed gas diffuser; 70 a front end disk; 72 diffuser positioning boss; 4 an annular combustion chamber; 11 nozzle ring front disc; 12 a nozzle ring vane; 13 nozzle ring rear disc; 14 cold air through holes; 15 cooling the ring; 112 sealing the push ring; 3, radial sealing ring; 41 an annular combustor outer ring; 5 a piston ring; 111 nozzle ring front disc positioning boss; 114 a dowel pin; 81 diffuser casing; 61 heat insulation tiles; 20 interstage seal inner chambers; 62 insulating tile locating pins.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the accompanying drawings, and it should be understood that the described embodiments are only some embodiments, but not all embodiments, of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by a person of ordinary skill in the art without creative efforts belong to the protection scope of the present invention.
Example one
As shown in fig. 1 and 2, in the interstage seal structure of the radial flow gas turbine, a nozzle ring 1, a compressor seal disk 2, a compression diffuser 7 and a cooling ring 15 are all annular structural members;
the nozzle ring 1, the compressor sealing disk 2, the compression diffuser 7, the cooling ring 15 and the annular combustion chamber 4 are coaxially assembled together;
the far axial center part of the sealing disk 2 of the compressor is fixed on the rear side surface of the front end disk 70 close to the axial center part of the air diffuser 7;
the nozzle ring 1 comprises a nozzle ring front disc 11, nozzle ring blades 12 and a nozzle ring rear disc 13;
the plurality of nozzle ring vanes 12 are uniformly arranged in the circumferential direction;
the front end of each nozzle ring blade 12 is attached and fixed to a nozzle ring front disc 11, and the rear end is attached and fixed to a nozzle ring rear disc 13;
cold air through holes 14 penetrating through the nozzle ring front disc 11, the nozzle ring blades 12 and the nozzle ring rear disc 13 are formed corresponding to the nozzle ring blades 12;
a cooling ring 15 is fixed at the rear side of the nozzle ring rear disc 13 corresponding to the cold air through hole 14; the cooling ring 15 is coaxial with and axially spaced from the nozzle ring rear disk 13;
the far axial center end of the front disc 11 of the nozzle ring is fixed on the back side surface of the far axial center end of the front disc 70 of the air pressure diffuser 7;
a circular sealing push ring 112 is formed on the front side of the near axle center end of the nozzle ring front disc 11;
a circular groove is formed on the back side surface of the front end wall 21 of the compressor sealing disc 2 corresponding to the sealing push ring 112;
the front end of the radial sealing ring 3 is fixedly embedded into the annular groove;
when the radial-flow gas turbine does not work, the sealing push ring 112 is pressed at the rear end of the radial sealing ring 3, so that the radial sealing isolation of the air compression diffuser 7 and the hot gas injection passage of the nozzle ring 1 is realized.
Preferably, the radial sealing ring 3 is an E-ring.
The axial symmetry section of the radial flow gas turbine is shown in fig. 1, and the working principle is to convert chemical energy into output mechanical energy; the compressor absorbs external air to compress, the annular combustion chamber 4 mixes and burns air with a certain pressure ratio and fuel gas, the burnt gas drives the turbine to rotate through the nozzle ring 1, and the turbine outputs mechanical energy (in a high-speed rotating rotor mode) to a user end through a rotor system. The direction of flow of the gas in the gas turbine is shown by the hollow arrows in fig. 1. The interstage sealing structure of the radial flow type gas turbine mainly functions to seal and separate outlet gas of a high-pressure compressor and inlet gas of a low-pressure turbine, so that the gas turbine can operate efficiently and reliably.
In the interstage sealing structure of the radial flow type gas turbine according to the first embodiment, when the radial flow type gas turbine does not work, the sealing push ring 112 is pressed on the rear end of the radial sealing ring 3, so that the radial sealing isolation of the gas compression diffuser 7 and the hot gas injection passage of the nozzle ring 1 is realized. When the radial-flow gas turbine works, the cold air through holes 14 of the nozzle ring 1, which penetrate through the nozzle ring front disc 11, the nozzle ring blades 12 and the nozzle ring rear disc 13, guide high-pressure cold air to the rear side of the nozzle ring rear disc 13 and cool the rear side of the nozzle ring through the cooling ring 15, so that the hot position of the nozzle ring front disc sealing push ring 112 can seal and separate high-pressure compressor outlet gas and low-pressure turbine inlet gas, the gas turbine can efficiently and reliably run, the defects of difficult piston ring sealing performance prediction and high requirement on an elastic ring manufacturing process are overcome, and the reliability and the economical efficiency of the interstage seal of the radial-flow gas turbine are improved.
Example two
According to the interstage seal structure of the radial flow type gas turbine of the first embodiment, the side, away from the axial center, of the nozzle ring front disc 11 and the annular combustion chamber outer ring 41 of the annular combustion chamber 4 are connected together in a sealing mode through the piston rings 5.
In the interstage sealing structure of the radial flow type gas turbine in the first embodiment, the annular combustion chamber 4 and the nozzle ring front disc 11 are sealed through the piston ring 5, so that the gas in the combustion chamber is prevented from leaking through a gap between the annular combustion chamber and the nozzle ring front disc 11, and the axial sealing isolation of the front end disc 70 of the compressor diffuser 7 and the combustion chamber is realized.
EXAMPLE III
Based on the interstage sealing structure of the radial flow type gas turbine in the first embodiment, a limiting ring 22 is formed on the front end wall 21 of the compressor sealing disk 2 far away from the rear side face of the shaft center part;
the seal pusher ring 112 is located on the proximal side of the stop collar 22.
Preferably, the far axial center of the compressor sealing disk 2 is fixed to the near axial center rear side of the front end disk 70 of the compressor diffuser 7 by bolts.
Example four
Based on the interstage sealing structure of the radial flow type gas turbine in the first embodiment, N diffuser positioning bosses 72 are uniformly formed on the rear side surface of the far shaft center part of the front end disc 70 of the compressor diffuser 7 along the circumferential direction, wherein N is an integer greater than 2;
n nozzle ring front disc positioning bosses 111 are uniformly formed on the front side of the far shaft end of the nozzle ring front disc 11 along the circumferential direction;
each diffuser positioning boss 72 is fixedly connected with the corresponding nozzle ring front disc positioning boss 111 through a positioning pin 114.
Preferably, N is 4, 6, 7, or 9, etc.
In the interstage seal structure of the radial flow gas turbine according to the fourth embodiment, the nozzle ring 1 is axially positioned by the nozzle ring front disc positioning boss 111, the nozzle ring front disc positioning pin 114 and the diffuser positioning boss 72.
EXAMPLE five
Based on the interstage sealing structure of the radial flow type gas turbine in the first embodiment, the compressor diffuser 7 is fixed in the diffuser casing 81;
a radial cold air flow channel is formed between the front end disc 70 of the air compressor diffuser 7 and the diffuser casing 81;
a far shaft annular wall 23 and a near shaft annular wall 24 which are coaxial are formed on the rear side of the front end wall 21 of the compressor sealing disk 2;
a circle of heat insulation tiles 61 are arranged between the rear ends of the far collar wall 23 and the near collar wall 24;
the front end wall 21 of the compressor sealing disk 2, the far collar wall 23, the near collar wall 24 and the heat insulation tile 61 form an inter-stage sealing inner cavity 20 together.
Preferably, the heat insulation tile 61 is in a fan ring shape; two wings of the heat insulation tile 61 are respectively provided with a groove and a convex rib; the heat insulation tiles 61 are matched with the convex ribs through the grooves and are connected end to form a complete heat insulation tile sealing ring.
Preferably, the far axial end of the heat insulation tile 61 is fixedly connected to the far collar wall 23 of the compressor sealing disc 2 through a heat insulation tile positioning pin 62; a heat insulation tile installation groove is formed in the rear end of the far shaft side of the near shaft annular wall 24 of the compressor sealing disk 2 along the circumferential direction; the end of the heat insulation tile 61 close to the shaft is inserted into the heat insulation tile installation groove and is in sealing fit with the annular wall 24 close to the shaft of the compressor sealing disc 2. The heat insulating tile 61 is axially positioned by the heat insulating tile positioning pin 62 and the heat insulating tile mounting groove.
EXAMPLE six
According to the interstage seal structure of the radial flow type gas turbine in the first embodiment, the cold air through holes 14 in the nozzle ring blades 12 penetrate through the geometric centers of the nozzle ring blades 12 in the axial direction.
Preferably, the nozzle ring vanes 12 are hollow vanes.
EXAMPLE seven
The simulation design method of the interstage seal structure of the radial flow type gas turbine comprises the following steps:
firstly, the position of the sealing push ring 112 under the working state (thermal state) of the gas turbine is estimated through the numerical simulation of the structure of the gas turbine.
And secondly, if the clearance between the sealing push ring 112 and the front end wall 21 of the compressor sealing disk 2 is larger than a set value, the cooling effect at the position of the cooling ring 15 is enhanced.
Because the temperatures of different areas of the gas turbine are obviously different when the gas turbine works, the temperature deformation differences of parts of different areas of the gas turbine are obvious, the sealing push ring 112 is possibly far away from the front end wall 21 of the sealing disk 2 of the compressor and cannot compress the front end of the radial sealing ring 3, and the radial sealing isolation of the gas compressor diffuser 7 and the hot gas injection channel of the nozzle ring 1 is difficult to ensure.
The cooling effect at the cooling ring 15 is increased, so that the sealing push ring 112 and the front end wall 21 of the sealing disc 2 of the compressor move oppositely, the sealing push ring 112 is pressed to the front end of the radial sealing ring 3, and the radial sealing isolation of the air compressor diffuser 7 and the hot air injection passage of the nozzle ring 1 is ensured.
And thirdly, based on the target axial force of the front side surface and the rear side surface of the E ring, combining the structure numerical value of the gas turbine, and determining the size of the heat exchange coefficient applied to the cooling ring 15 in a simulation mode.
The target axial force is X times the minimum value of the effective seal of the E ring calibrated by a manufacturer, and X is between 1 and 1.3 (for example, X is 1.2).
The sealing effect of the E ring is in direct proportion to the magnitude of the axial force acting on the two side faces of the E ring, and in order to achieve the optimal sealing effect, the numerical values of the target axial forces acting on the front side face and the rear side face of the E ring are required to be larger than the minimum value of the effective sealing of the E ring calibrated by a manufacturer. The target axial force may be a factory calibrated minimum value of the E-ring effective seal multiplied by an amplification factor X, taking into account errors in the numerical calculation.
And fourthly, combining the size of the heat exchange coefficient applied to the cooling ring 15 and the aerodynamic value, simulating and determining the size of the cold air through holes 14 of the nozzle ring blades.
The hole size of the nozzle blade cold air channel 14 determined by the simulation design method of the interstage seal structure of the radial flow type gas turbine can ensure that the seal push ring 112 moves towards the front end wall 21 of the compressor seal disc 2 in the opposite direction under the working state (thermal state) of the gas turbine, so that the interstage seal of the radial flow type gas turbine is realized.
The above are merely preferred embodiments of the present application and are not intended to limit the present application. Various modifications and changes may occur to those skilled in the art. Any modification, equivalent replacement, improvement and the like made within the spirit and principle of the present application shall be included in the protection scope of the present application.
Claims (12)
1. The interstage seal structure of the radial-flow gas turbine is characterized in that a nozzle ring (1), a compressor seal disk (2), a compression diffuser (7) and a cooling ring (15) are all annular structural members;
the nozzle ring (1), the compressor sealing disc (2), the compression diffuser (7), the cooling ring (15) and the annular combustion chamber (4) are coaxially assembled together;
the far shaft center of the sealing disk (2) of the compressor is fixed on the rear side surface of the front end disk (70) of the air compressor diffuser (7) close to the shaft center;
the nozzle ring (1) comprises a nozzle ring front disc (11), nozzle ring blades (12) and a nozzle ring rear disc (13);
the nozzle ring blades (12) are uniformly arranged along the circumferential direction;
the front end of each nozzle ring blade (12) is attached and fixed to a nozzle ring front disc (11), and the rear end is attached and fixed to a nozzle ring rear disc (13);
cold air through holes (14) which penetrate through the nozzle ring front disc (11), the nozzle ring blades (12) and the nozzle ring rear disc (13) are formed corresponding to the nozzle ring blades (12);
a cooling ring (15) is fixed at the rear side of the nozzle ring rear disc (13) corresponding to the cold air through hole (14); the cooling ring (15) is coaxial with and axially spaced from the nozzle ring rear disc (13);
the far axle center end of the front disc (11) of the nozzle ring is fixed on the rear side surface of the far axle center end of the front disc (70) of the air pressure diffuser (7);
a circular sealing push ring (112) is formed on the front side of the near-axis end of the nozzle ring front disc (11);
the back side surface of the front end wall (21) of the compressor sealing disk (2) is provided with a circular groove corresponding to the sealing push ring (112);
the front end of the radial sealing ring (3) is fixedly embedded into the annular groove;
the sealing push ring (112) is pressed at the rear end of the radial sealing ring (3).
2. The interstage seal structure of a radial flow gas turbine according to claim 1,
the radial sealing ring (3) is an E ring.
3. The interstage seal structure of a radial flow gas turbine according to claim 1,
the side of the front disc (11) of the nozzle ring, which is far away from the axle center, is connected with an outer ring (41) of the annular combustion chamber (4) in a sealing way through a piston ring (5).
4. The interstage seal structure of a radial flow gas turbine according to claim 1,
a limiting ring (22) is formed on the rear side surface of the far shaft center part of the front end wall (21) of the compressor sealing disk (2);
the sealing push ring (112) is positioned on the side close to the axial center of the limiting ring (22).
5. The interstage seal structure of a radial flow gas turbine according to claim 1,
the far shaft center part of the sealing disk (2) of the compressor is fixed to the rear side surface of the front end disk (70) of the air compressor diffuser (7) close to the shaft center part through bolts.
6. The interstage seal structure of a radial flow gas turbine according to claim 1,
n diffuser positioning bosses (72) are uniformly formed on the rear side surface of the far shaft center part of a front end disc (70) of the air compression diffuser (7) along the circumferential direction, and N is an integer greater than 2;
n nozzle ring front disc positioning bosses (111) are uniformly formed on the front side of the far shaft end of the nozzle ring front disc (11) along the circumferential direction;
each diffuser positioning boss (72) is fixedly connected with the corresponding nozzle ring front disc positioning boss (111) through a positioning pin (114).
7. The interstage seal structure of a radial flow gas turbine according to claim 6,
n is 4, 6, 7 or 9.
8. The interstage seal structure of a radial flow gas turbine according to claim 1,
the gas compressor diffuser (7) is fixed in the diffuser casing (81);
a radial cold air flow channel is formed between a front end disc (70) of the air compressor diffuser (7) and a diffuser casing (81);
a far shaft annular wall (23) and a near shaft annular wall (24) which are coaxial are formed on the rear side of the front end wall (21) of the compressor sealing disk (2);
a circle of heat insulation tiles (61) are arranged between the rear ends of the far shaft ring wall (23) and the near shaft ring wall (24);
the front end wall (21) of the compressor sealing disc (2), the far shaft annular wall (23), the near shaft annular wall (24) and the heat insulation tile (61) jointly form an interstage sealing inner cavity (20).
9. The interstage seal structure of a radial flow gas turbine according to claim 8,
the heat insulation tile (61) is in a fan ring shape;
two wings of the heat insulation tile (61) are respectively provided with a groove and a convex rib;
the heat insulation tiles (61) are matched with the convex ribs through the grooves and are connected end to form a complete heat insulation tile sealing ring.
10. The interstage seal structure of a radial flow gas turbine according to claim 9,
the far shaft end of the heat insulation tile (61) is fixedly connected to the far shaft annular wall (23) of the compressor sealing disc (2) through a heat insulation tile positioning pin (62);
a heat insulation tile installation groove is formed in the circumferential direction at the rear end of the far shaft side of the near shaft annular wall (24) of the compressor sealing disk (2);
the end close to the shaft of the heat insulation tile (61) is inserted into the heat insulation tile mounting groove and is in sealing fit with the annular wall (24) close to the shaft of the compressor sealing disc (2).
11. The interstage seal structure of a radial flow gas turbine according to claim 1,
the cold air through holes (14) on the nozzle ring blades (12) penetrate through the geometric centers of the nozzle ring blades (12) along the axial direction.
12. The interstage seal structure of a radial flow gas turbine according to claim 1,
the nozzle ring vanes (12) are hollow vanes.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
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| CN202021966378.4U CN212774523U (en) | 2020-09-10 | 2020-09-10 | Interstage sealing structure of radial flow type gas turbine |
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| CN202021966378.4U CN212774523U (en) | 2020-09-10 | 2020-09-10 | Interstage sealing structure of radial flow type gas turbine |
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| CN212774523U true CN212774523U (en) | 2021-03-23 |
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