RU2519127C1 - Turbine of gas turbine engine and method for adjustment of radial clearance in turbine - Google Patents

Abstract

FIELD: machine building.

SUBSTANCE: turbine of a bypass gas turbine engine comprises a cooled stage with a nozzle diaphragm with cavities above and below it and a turbine rotor with a cooled runner. The turbine stator comprises at least two turbine casings with cavities in-between and a radial clearance adjustment system including a ring insertion above the turbine runner. The cavity above the nozzle diaphragm is connected to a compressor output by an air bleed pipeline including a flow regulator. The radial clearance adjustment system comprises an on-board computer, sensors for radial clearance measurement and microwave radiation sources set above the insertion. The radial clearance measuring sensors and the microwave radiation sources are connected by electric links with the on-board computer. The microwave radiation sources are made so that to be able of heating the ring insertion. The radial clearance is measured and basing on its value the microwave radiation sources are switched on.

EFFECT: effective control of radial clearances in a turbine in all modes, increasing engine thrust in takeoff and afterburner modes, increasing turbine efficiency factor and reliability.

4 cl, 7 dwg

Description

The group of inventions relates to engine building, including aircraft and stationary gas turbine engines of gas turbine engines, having two circuits, and can find application in aircraft building, shipbuilding, gas pumping stations and for peak power plants as a drive for an electric generator designed to generate electricity.

A known turbine of a gas turbine engine according to the invention patent No. 2435039 IPC F01D 11/24, published April 27, 2008. The turbine housing includes a radial wall and contains a support for fastening the ring surrounding the turbine blades from its inner surface. The support comprises a peripheral wall surrounding the ring coaxially with it. The housing includes many perforations that provide air for uniform ventilation of the outer surface of the peripheral wall. Perforations are formed through the radial wall of the housing, passing radially inward. The wall essentially encompasses the ventilation chamber, which is also formed by the inner surface of the housing and the outer surface of the peripheral wall of the support. The ventilation chamber includes a small hole between the radial rib of the support and the inner surface of the radial wall to release air from the chamber.

Disadvantages - structural complexity and the inability to regulate the radial clearance in all engine operating modes.

Known gas turbine engine according to the patent of the Russian Federation for the invention No. 2304221, IPC F01D 11/14, publ. 08.10.07, This gas turbine engine contains a compressor having several axial stages containing a housing, guiding apparatuses and impellers, and a turbine comprising a housing and at least one stage with a nozzle apparatus and an impeller, as well as at least radial clearance control means one stage of the compressor and / or turbine.

Disadvantages - low efficiency of regulation of the radial clearance, especially in transition modes, when forcing or throttling the engine, the structural complexity of the device for regulating the radial clearance.

A gas turbine, such as a high-pressure turbine for a turbomachine, such as the French patent publication No. 2688539 disclosed in France, typically contains a plurality of fixed vanes arranged so that they alternate with a plurality of movable vanes located in the path of the hot gas coming from the combustion chamber of the turbomachine. Moving turbine blades are surrounded around their periphery by a stationary annular assembly. The stationary annular assembly forms a passage along which hot gas flows through the turbine blades.

In order to increase the efficiency of such a turbine, as is known, the gap that exists between the vertices of the moving turbine blades and the parts of the stationary annular assembly facing them is reduced to the value that is as small as possible.

For this, tools have been developed that provide the ability to change the diameter of the stationary annular assembly.

Nevertheless, this solution is considered insufficient if the support to which the ring is attached is also exposed at its periphery to uneven thermal deformation, when such deformation leads to deformation of the turbine ring.

Also known is a turbine engine with adjustable radial clearances according to the patent of Russian Federation No. 2435039, IPC F01D 11/04, a prototype of the method and device.

This method of controlling the radial clearance in a turbine includes cooling and / or heating the rotor and / or stator.

This turbine contains an outer, inner and intermediate case, a stage with a nozzle apparatus and an impeller with an annular insert above the impeller, and also means for regulating the radial clearances of at least one stage of the turbine, while the annular insert above the impellers is fixed on the intermediate and external cases .

The disadvantages of the method and device is a sharp increase in the radial clearance during engine boosting due to the rapid heating of the housing.

The technical result achieved during the creation of the invention is to maintain radial clearances constant in all modes of operation of the turbine.

The group of inventions relates to gas turbine engines.

Objectives of the invention: effective regulation of radial clearances in the turbine in all modes, increasing engine thrust in take-off and afterburning modes, increasing the efficiency and reliability of the turbine.

The solution of these problems was achieved in a turbine of a gas turbine engine containing at least one cooled stage with a nozzle apparatus with cavities above and below it and a turbine rotor with a cooled impeller, as well as a turbine stator containing at least two turbine bodies with cavities between them and a radial clearance control system comprising an annular insert above the turbine impeller, in that, according to the invention, the cavity above the nozzle apparatus is connected by an air sampling pipe containing a regulator p the exit from the compressor, the radial clearance control system includes an on-board computer, radial clearance measurement sensors and microwave radiation sources mounted above the insert, while radial clearance measurement sensors and microwave radiation sources are connected by electrical connections to the on-board computer. The annular insert may be hollow. The inner cavity of the annular insert may be filled with a heat storage substance.

The solution to these problems has been achieved in a method for regulating the radial clearance in a turbine of a double-circuit gas turbine engine, including cooling the rotor and heating the stator, by the fact that according to the invention, the radial clearance is measured and, depending on its size, the microwave sources are turned on to heat the annular insert.

The invention is presented in the drawings (figures 1-7), where:

- figure 1 shows a diagram of a gas turbine engine,

- figure 2 presents a diagram of a turbine and a radial clearance control system in a turbine using the example of one stage of a two-stage turbine,

- figure 3 shows a view of the device,

- figure 4 shows a view of a device with an annular insert having a coating,

- figure 5 shows a view of the device and an annular insert with panels of "honeycomb seals",

- figure 6 shows a diagram of the change in air flow for cooling the turbine rotor depending on the temperature in front of the turbine,

- Fig.7 shows a diagram of the change in air flow for cooling the turbine stator depending on the time of operation of the gas turbine engine.

The design of the gas turbine engine is shown in FIG. 1-7. A gas turbine engine (GTE) comprises an input device 1, with an intake fairing 2, a fan 3, a main body 4, a nozzle 5, a compressor 6, a combustion chamber 7 with a body 8, a flame tube 9 and nozzles 10, a turbine 11, shafts 12 and 13, supports 14 ... 17 (figure 1). The shafts in the turbine 11 can be not only two, but also one or three.

The compressor 6 includes a housing 18, at least one stage 19, which in turn contains a guide apparatus 20 and rotor blades 21 and disks 22.

The turbine 11 contains at least one rotor 23 and the stator 24. The turbine 11 has at least one stage 25. Figure 1 shows a turbine 11 with two stages 25, each of which, in turn, contains a nozzle apparatus 26 and an impeller 27 with the working blades 28 and the disk 29. The steps 25 of the turbine 11 may be one or more two. The nozzle apparatus 26 and the working blades 28 are made cooled, for example perforated. The disk 29 has front and rear deflectors 30 and 31 on both sides. The steps 25 of the turbine 11, as mentioned earlier, can be one, three or as many as you want, and the radial clearance control tool is used on one, or several, or all stages 25 of the turbine 11. The most effective is the use of radial clearance control in the first stages of the turbine due to the high pressure drop across them.

A dual-circuit gas turbine engine can be single-circuit or have two circuits: the first 32 and second 33 (Fig. 1).

The turbine 11 comprises means for adjusting the radial clearance. The radial clearance control means comprises an annular insert 34 mounted inside the stator 24 above the rotor blades 28 of the turbine 11 to form a radial clearance 5. The annular insert 34 can be solid or hollow, i.e. contain a cavity 35. The annular insert 34 may be segmented and consist of segments 36.

The invention is further described by the example of one first stage 25 of a high pressure turbine (first), but can be applied to other (all) stages 25 of a turbine 11.

The working blades 28 can be made with retaining shelves (this option is not shown). The working blades 28 contain a locking part 37 (figure 2). In the disk 29, holes 38 are made for supplying cooling air to the working blades 28. The front deflector 30 is sealed relative to the shaft 13 and the stator parts with seals 39 relative to the ring 40 and the seal 41 relative to the stator part 42. Openings 43 for supplying cooling air are made in the front deflector 30.

The radial clearance control means, in addition to the previously listed means, includes a swirling device 44, an internal cooling air supply pipe 45, an internal cavity 46, openings 47, an internal cavity 48 of the nozzle apparatus 26, openings 49, an upper cavity 50 inside the annular manifold 51, openings 52, high pressure pipe 53, air flow regulator 54 with drive 55. The other end of the high pressure pipe 53 is connected to a manifold 56 at the outlet of the compressor 6.

The shaft 13, the disk 26 with the deflectors 30 and 31 and the impeller 27 form the rotor 23 of the turbine 11 (figure 2) The turbine 11, as mentioned earlier, has a stator 24. The stator 24 contains several buildings (from two or more).

The following is a description of a turbine 11 with three casings: an outer casing 57, an inner casing 58, and an intermediate casing 59 installed between them. The intermediate casing 59 has a conical shape (truncated cone shape) with a radial flange 60 connected to the flange 61 of the outer casing 57. In addition, the intermediate casing 59 has a front radial baffle 62 mounted with an annular gap 63 inside the outer casing 57. The outer casing 57 has a first radial baffle 64 with rectangular windows 65 and a second radial baffle rodku 66 with holes 67 for air vent (1 and 2), a cooling turbine 24, a stator 11 (Figures 2 and 5). As a result, four cavities 68, 69, 70 and 71 are formed in the stator 24 of the turbine 11.

The first radial baffle 64 comprises an "omega-shaped" part 72, which is connected by a weld seam 73 to the first radial baffle 64. On the other side of the (internal) "omega-shaped" part 72 of the first radial baffle 64, an annular part 75 is welded with an annular a groove 76 for accommodating therein an annular protrusion 77 provided on the annular insert 34 for centering it.

The main features of the turbine 11 is the presence of radial clearance measurement sensors 78, microwave radiation sources 79 and an on-board computer 80 connected by electrical connections 81. Only one radial clearance measurement sensor 78 can be used, but this is extremely undesirable since failure of the sensor 78 may lead to an emergency.

On-board computer 80 is based on modern digital processors.

On the inner surface of the annular inserts 34, a soft, easy-to-abrade coating 85, for example graphite (FIG. 4), or honeycomb inserts 86 are attached (FIG. 5).

Figure 6 shows a diagram of the change in air flow for cooling the rotor of the turbine 11 pos.87 depending on the temperature in front of the turbine - Tg, from which it follows that the air flow g1 cooling the rotor 23 of the turbine 11 should increase with increasing temperature of the combustion products in front of the turbine Tg. This dependence can be linear, for example, as shown in Fig.6. Figure 7 shows a diagram of the change in air flow for cooling the turbine stator depending on the operation time of the gas turbine engine. For clarity, the calculated cooling air flow rates g2 are given for cooling the turbine stator 11 in three sections of the gas turbine engine operation (in boost mode 88, 89, and 90). Positions 91, 92 and 93 show the actual change in air flow rate g2.

In the intermediate casing 59, holes 94 are made connecting the cavities 69 and 56 for supplying cooling air to the cavities 68, 69 and 70 (Fig. 4).

The operation of the turbine GTE is as follows (figure 1 ... 7).

With a sharp change in the operating mode of the turbine of a gas turbine engine, for example, when it is forced, the temperature of the combustion products in front of the turbine increases. In the nominal mode, the radial clearance 60, has a calculated value, and in the afterburner (maximum) mode, the radial clearance 5 at the initial moment in the absence of regulation would increase sharply. When forcing a gas turbine engine, the temperature of the combustion products rises sharply. In this case, the turbine bodies 57 ... 59 and the disk 29 with rotor blades 28 are warmed up. But the mass of the disk 29 of the turbine 11 is much larger than the mass of all the bodies 57 ... 59, so the gap would increase without the use of radial clearance control. The presence of a hollow annular insert 34 filled with heat-accumulating material 36 will slow down the heating of the hollow annular insert 34 and the housings 57 ... 59, which will prevent an increase in the radial clearance.

Passing through the high pressure pipe 53 through the flow regulator 54 cooling air cools the disk 29 of the turbine 11 and the blades 28.

Moreover, the change in the flow rate of cooling air through the flow regulator 54 is carried out only depending on the operating mode of the engine Tg and the change in the flow rate of this air is not controlled by the radial clearance, since an increase in the flow rate of this air reduces the efficiency of the turbine 11. In this case, the high pressure pipe 53 can only be connected to the exit of the compressor 6 (i.e., beyond its last stage, otherwise the cooling air pressure will not be enough to cool the perforated nozzle apparatus 26 and perforated side blades 28 of the turbine 11).

With an increase in the radial clearance, the radial clearance sensors 78 record this fact, and the on-board computer 80 via the communication channel 84 to the drive 55 of the flow controller 54 to increase the flow rate of cooling air and simultaneously to the switch 82 to turn on microwave sources 79. Sources of microwave radiation 79 practically instantly and evenly heat up the annular insert 34. This prevents the blades 28 from touching the annular insert. If the specified gap is exceeded by microwave sources 79, the on-board computer 80 turns them off.

Sources of microwave radiation 79 are short-term and warm up a small mass, so the energy costs are low.

When the radial clearance decreases below the permissible limit, on the contrary, the cooling air flow is reduced. As a result, the proposed system can very accurately maintain radial clearances constant in almost all modes.

The use of a group of inventions allowed:

1. To provide effective smooth regulation of radial clearances in the turbine of a gas turbine engine in all modes.

2. To provide an increase in engine power in afterburner (maximum) modes by reducing the radial clearance in these modes.

3. Ensure reliable take-off of the aircraft with engines equipped with such radial clearance control systems without preliminary heating of the turbine engine, or significantly reduce the warm-up time of the gas engine. This is necessary for military aircraft.

4. Ensure reliable take-off at high ambient temperatures, ie in conditions when the takeoff thrust of the gas turbine engine decreases.

5. Almost instantly transfer the operation mode of the gas turbine engine from cruising to afterburning mode. This is especially important for military aircraft.

6. Simplify the design of the elements of the radial clearance control system, reduce its weight and place the gas turbine engine outside the tract in the low temperature zone, which will increase the reliability of the turbine.

Claims (4)

1. A turbine of a gas turbine engine, comprising at least one cooled stage with a nozzle apparatus with cavities above and below it, and a turbine rotor with a cooled impeller, and a turbine stator comprising at least two turbine bodies with cavities between them and a control system a radial clearance containing an annular insert above the impeller of the turbine, characterized in that the cavity above the nozzle apparatus is connected by an air intake pipe containing a flow regulator with an outlet from the compressor, a reg the radial clearance measuring device comprises an on-board computer, radial clearance measurement sensors and microwave radiation sources mounted above the insert, wherein the radial clearance measurement sensors and microwave radiation sources are electrically connected to the on-board computer, the microwave radiation sources being configured to heat the annular insert . 2. A dual-circuit gas turbine engine containing a turbine according to claim 1, characterized in that the annular insert is made hollow. 3. The dual-circuit gas turbine engine according to claim 2, characterized in that the inner cavity of the insert is filled with a heat-accumulating substance. 4. The method for controlling the radial clearance in the turbine of a double-circuit gas turbine engine according to claim 1, including cooling the rotor and heating the stator, characterized in that the radial clearance is measured and, depending on its size, the microwave sources are turned on to heat the annular insert.

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