RU2369749C1 - Two-stage turbine of has turbine engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: two-stage turbine of has turbine engine comprises inner shelves of cooled second-stage nozzle vanes representing four-wall boxes with contact sites. Turbine interdisk space is sealed by front and rear intermediate disks that make together with aforesaid boxes a labyrinth seal. The said seal is divided into front seal arranged between front intermediate disk knife edges and mating surface of the box side wall with enlarged radial gap H, and rear seal. Contact sites are made on the front wall of the said box over its entire height. Front space of two-space nozzle vane communicates with slot-like space of front labyrinth seal with H/h=2…5, where H is radial gap of the front labyrinth seal, and h is that of the rear labyrinth seal.

EFFECT: higher reliability due to reduced temperature of vane contact sites and intermediate disk rim.

3 dwg

Description

The invention relates to two-stage turbines of gas turbine engines of aviation and ground applications.

A two-stage turbine of a gas turbine engine is known, the nozzle blades of the second stage in which, along the inner shelves, are fixed between themselves by radial trunnions telescopically mounted in the retaining ring. (S. A. Vyunov. “Design and Design of Aircraft Gas Turbine Engines.” Moscow, “Mechanical Engineering”, p. 183, Fig. 4.41a).

A disadvantage of the known design is its low reliability due to wear of the radial trunnions and retaining ring during vibration of the nozzle blades of the second stage.

Closest to the claimed one is the design of a two-stage turbine of a gas turbine engine, in which the inner shelves of the nozzle blades of the second stage are made in the form of four-wall boxes fixed to each other by contact pads located on the rear continuous wall of the box along the gas flow, and the inter-disk cavity is sealed with the front and rear downstream gas intermediate discs (patent RU No. 2151884).

A disadvantage of the known design adopted for the prototype is its low reliability due to the hardening and wear of the contact pads of the nozzle blades and the increased temperature of the intermediate disks washed around the periphery by high-temperature gas.

The technical problem solved by the invention is to increase the reliability of a two-stage turbine of a gas turbine engine by lowering the temperature of the contact pads of the second stage nozzle blades and the rim of the intermediate disks.

The essence of the invention lies in the fact that in a two-stage turbine of a gas turbine engine, the internal shelves of the cooled nozzle blades of the second stage in which are made in the form of four-wall boxes with contact pads, and the interdisk cavity is sealed with front and rear intermediate air disks forming a labyrinth seal with the boxes, according to the invention, the seal is divided into the front between the combs of the front intermediate disk and the mating surface of the front side wall of the box with increased radial clearance H and the rear seal, while the contact pads are made on the upstream side wall of the box and along the entire height of this wall, and the front cavity of the two-cavity nozzle vanes at the outlet is connected through the channel between the contact areas of adjacent vanes with the slot cavity of the front labyrinth seal , while H / h = 2 ... 5, where:

N - radial clearance along the front labyrinth seal;

h is the radial clearance along the rear labyrinth seal.

Separation of the labyrinth seal on the front between the combs of the front intermediate disk and the mating surface of the front side wall of the box with an increased radial clearance and on the rear seal reduces the dependence of the cooling air flow through the front cavity of the second nozzle blade on the radial clearance in the labyrinth seal, since in the case of a significant reduce the radial clearance in the rear labyrinth seal (for example, when operating in transient conditions) the cooling air can can flow through the front labyrinth seal with increased radial clearances, which increases the reliability of a two-stage turbine of a gas turbine engine.

The implementation of the contact pads on the upstream side wall of the box and along the entire height of this wall allows to reduce parasitic gas leakage through the axial gaps between the inner shelves of adjacent second stage nozzle blades, since the contact pads in this case work as sealing elements, which increases efficiency and reliability turbines.

The connection at the outlet of the front cavity of the blade through the channel between the contact pads of adjacent blades with the slot cavity of the front labyrinth seal allows you to organize a labyrinth seal between the inner shelf of the second stage nozzle blade and the intermediate disks with high pressure cooling air, which significantly reduces the flow of high-temperature gas through the labyrinth seal with the corresponding lowering the temperature of the rim of the intermediate disks and increasing their reliability. At the same time, the temperature of the pads pressurized by cooling air is reduced. Such repeated use of cooling air makes it possible to use its coolant to the fullest extent and minimize the deterioration of the parameters of the gas turbine engine during the selection of air for cooling the turbine.

With H / h <2, it is possible to “lock” the flow rate of cooling air flowing through the front cavity of the second stage nozzle blade, especially during transient operation of the gas turbine engine, which can cause overheating and breakage of the second stage nozzle blade.

At H / h> 5, the reliability of a two-stage turbine of a gas turbine engine decreases due to the increased mixing of high-temperature gas with cooling air, with heating and a decrease in the reliability of the intermediate disks.

Figure 1 shows a longitudinal section of a two-stage turbine of a gas turbine engine.

Figure 2 is a section aa in figure 1.

Figure 3 - element 1 in figure 1 in an enlarged view.

A two-stage turbine 1 of a gas turbine engine consists of a stator 2 with nozzle blades of the first stage 3 and nozzle blades of the second stage 4, as well as a rotor 5 with impellers of the first stage 6 and second stage 7, the interdisk cavity 8 between which is sealed by an intermediate gas stream 9 intermediate the disk 10 and the rear intermediate disk 11. The feather 12 of the nozzle blade of the second stage 4 is made cooled and divided by the radial partition 13 into the front 14 and rear 15 along the gas stream 9 internal cavities. The rear internal cavity 15 is connected at the inlet to an intermediate stage of the compressor (not shown), and the front internal cavity 14 is connected at the inlet to the high pressure air cavity 16. Such separate cooling of the nozzle blade of the second stage allows you to save energy-expensive high-pressure air, which helps to improve the parameters of the gas turbine engine and increase the reliability of the turbine. The inner shelf 17 of the nozzle blade of the second stage 4 is made in the form of a four-wall box 18, on the continuous front wall 9 of the wall 19 which are located contact pads 20 and 21, made along the entire height of the front wall 19. The inner surface 22 of the front wall 19 forms together with the sealing combs 23 of the front intermediate disk 10, the front labyrinth seal 24, and the inner wall 25 of the box 18 together with the honeycomb seal 26 form together with the sealing combs 27 of the front intermediate disk 10 and the rear intermediate disk 11, the rear labyrinth seal 28. The front inner cavity 14 of the second stage nozzle blade 4 at the outlet through the transition channel 29 and through the channel 30 between the contact pads 20, 21 is connected to the slotted cavity 31 of the front labyrinth seal 24. To compensate for thermal expansion, the inner shelves 17 nozzle blades of the second stage 4 are made with axial clearances 32.

The device operates as follows.

When the two-stage turbine 1 of the gas turbine engine is operating, the nozzle blades of the second stage 4 are in contact with each other by the contact pads 20 and 21, which prevents vibration and breakage of the blades 4. During long-term operation on the working surface of the contact pads 20 and 21, a rivet could form, which would lead to wear contact pads and the appearance of vibration stresses on the blades 4. However, this does not happen, since the cooled pads 20 and 21 throughout the resource retain high hardness, which eliminates their hardening and wear, which, in turn, increases the reliability of the turbine.

Claims (1)

A two-stage turbine of a gas turbine engine, the internal shelves of the cooled nozzle blades of the second stage in which are made in the form of four-wall boxes with contact pads, and the inter-disk cavity is sealed with front and rear intermediate air disks forming the labyrinth seal with the boxes, characterized in that the seal is divided into the front between the combs of the front intermediate disk and the mating surface of the front side wall of the box with an increased radial clearance H and the rear densification, while the contact pads are made on the upstream side wall of the box and along the entire height of this wall, and the front cavity of the two-cavity nozzle vane at the outlet is connected through the channel between the contact pads of adjacent vanes with the slot cavity of the front labyrinth seal, with H / h = 2 ... 5,
where N is the radial clearance along the front labyrinth seal,
h is the radial clearance along the rear labyrinth seal.

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