RU2496991C1 - Bypass gas turbine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed turbine comprises cooled stage with distributor provided with cavities there above and below, turbine rotor with cooled wheel and swirler arranged there ahead and turbine stator. Stator comprises turbine housing, circular insert above turbine wheel and radial clearance adjustment system. Turbine housing consists of two parts, i.e. front and rear parts with circular inserts arranged there between. Stator is cooled by second-stage air. Stator cooling air feed system is composed of second-stage air intake, driven flow rate controller and air intake casing. Perforated jacket is arranged in air intake casing. Said system incorporates also onboard computer, radial clearance transducers and flow rate controller drive. Radial clearance transducers are electrically connected with onboard computer.

EFFECT: higher engine thrust at takeoff and afterburning, engine efficiency and reliability.

5 cl, 15 dwg

Description

The invention relates to engine building, including aircraft and stationary gas turbine engines of a gas turbine engine 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.

The known turbine of a gas turbine engine according to the patent for invention No. 2435039 IPC F01D 11/24 UBL 04/27/08, the turbine housing includes a radial wall and contains from its inner surface a support for fastening the ring surrounding the movable blades of the turbine. 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 devices and working stakes, and a turbine containing a housing and at least one stage with a nozzle apparatus and an impeller, as well as a means for regulating radial clearances, according to at least 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, for example, a high pressure turbine for a turbomachine, such as the French patent disclosed in Patent No. 2688539, typically contains a plurality of fixed blades arranged so that they alternate with a plurality of movable blades in the path of the hot gas 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 the Russian Federation No. 2435039, IPC F01D 111/04, 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 enclosures

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 the take-off and afterburning mode, increasing the efficiency and reliability of the turbine.

The solution of these problems was achieved in a turbine of a double-circuit gas turbine engine, containing at least one cooled stage with a nozzle apparatus with cavities above and below it, and at least one turbine rotor with a cooled impeller and a spin device in front of it, as well as a stator a turbine comprising, a turbine housing, an annular insert above the turbine impeller and a radial clearance control system, in accordance with the invention, the turbine housing is made of two parts: front and rear, an annular insert renewed between them, the stator is made of cooled air of the second circuit, while the supply system of the cooling stator air is made in the form of an air intake installed in the second circuit, a flow regulator with a casing drive in the cavity of which a perforated casing is installed, and also contains an on-board computer and radial clearance sensors, the drive of the flow regulator, and radial clearance sensors 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. Ribs may be provided on the annular insert and / or turbine housing. Radial openings can be made in the turbine housing in front of the annular insert, and through axial bushings can be made in the annular insert

The invention is presented in the drawings (figures 1-15), 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 presents the second diagram of the turbine and the control system of the radial clearance in the turbine on the example of one stage of a two-stage turbine,

- figure 4 shows the view of the insert,

- figure 5 shows a view A,

- figure 6 shows the insert with heat-storage filler,

- Fig.7 shows a view of an annular insert with holes in it,

- Fig.8 shows an annular insert with ribs,

- figure 9 shows the annular insert with turbulators,

- figure 10 shows the annular insert coated with a soft abradable material,

- figure 11 shows the annular insert with panels of "honeycomb seals",

- Fig.12 shows the appearance of the air intake,

- Fig.13 shows a view In

- Fig.14 is a diagram of the change in air flow for cooling the turbine rotor depending on the temperature in front of the turbine,

- Fig.15 shows a diagram of the change in air flow for cooling the turbine stator depending on the operation time of the gas turbine engine.

The design of a dual-circuit gas turbine engine shown in the drawings of figures 1-15. A double-circuit gas turbine engine (GTE) contains an input device 1, with an input 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 I, 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, the impeller 27 with the 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 any number, and the radial clearance control tool is applied on one or more or all of the steps 25 of the turbine 11. The most efficient the use of radial clearance control in the first stages of the turbine 11 due to the high pressure drop across them.

The dual-circuit gas turbine engine has two circuits: the first 32 and the second 33 (figure 1). The air of the second circuit 33 has a lower temperature than the air in the compressor 6 due to the fact that when the air is compressed, its temperature rises. As a result, it is preferable to use the air of the second circuit 33 to control the radial clearances in the turbine 11.

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 δ. The annular insert 34 can be solid (Fig. 5) or hollow (Fig. 6), i.e. contain a cavity 35. The cavity 35 may be filled with a heat storage substance 36. A heat storage substance 36 is a material having high heat capacity and phase transition heat, for example, based on sodium acetate.

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 in figure 1 ... 16). The working blades 28 contain a locking part 37. 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 8 and the stator parts with seals 39 and 40. In the front deflector 30, holes 41 are made for supplying cooling air.

The cooling system of the rotor 23 of the turbine 11 includes a swirl apparatus 42, an internal cooling air supply pipe 43, an internal cavity 44, holes 45, an internal cavity 46 of the nozzle device 26, holes 47, an upper cavity 48 of the sleeve 49, a high pressure pipe 50, a flow regulator 51. The other end of the high pressure pipe 50 is connected to the outlet of the compressor 6.

In addition, the radial clearance control means has air inlets 52 installed in the second circuit 33 above the stator 24 of the turbine 11. Each air inlet 52 has an air intake pipe 53 and a flow regulator 54. The air inlets 52 are installed in the second circuit 33 and are designed for dosed intake of cooling air from the second circuit 33. In total, from 2 to 12 air inlets 52 can be applied. In more detail, the design of the air inlets 52 is shown in FIGS. 4, 14 and 15.

The air inlet 52, in addition to the air intake pipe 53 and the flow regulator 54 with the drive 55, contains a housing 56 made concentric with the stator 24 of the turbine 11 and forming a cavity 57 inside which a perforated casing 58 is installed to supply air in the form of a “shower” to the stator 24 of the turbine 11 and the ring insert 34 .. To intensify the cooling on the stator 24 can be performed longitudinal ribs 59.

The stator 24 includes a housing 60 with a flange 61, an annular insert 34 and an inner shell 62. The housing 60 consists of two parts of the front 63 and the back 64 between which the annular insert 34 is installed. The housing 8 of the combustion chamber 7 has a flange 65. connected to the flange 61. Between flanges 61 and 65 are fixed conical spacer 666 with flange 67.

The second embodiment of the cooling circuit of the stator 24 of the turbine 11 is shown in Figs. 8 and 9. To implement this method, holes 68 are made in the front part 63 of the housing 60 connecting the cavity 57 with the cavity 69 between the front part 63 of the housing 60 and the annular insert 34.

The rear part 64 of the housing 60 has a radial partition 70 with holes 71, which contains an annular part 72, which is welded 73 to the radial partition 70. An annular groove 74 is made in the annular part 72 to accommodate the annular protrusion 75 present on the annular insert 34 to center it. On the other hand, the annular insert 34 is made of an annular protrusion 76 mounted in the cavity 77 ..

The heat storage material 36, as noted above, is a material that has a high heat capacity and a high specific heat of phase transition. An example of such a material is sodium acetate trihydrate.

Thermophysical properties of this material:

- heat of fusion 220 kJ / kg,

- heat capacity of the solid phase 2 kJ / kg,

- heat capacity of the liquid phase 2, 8 kJ / kg

Heat storage is usually carried out due to the heat of the phase transition. By selecting the volume of the heat-accumulating material 36, the heating time of the disk 29 and the housing 60 of the turbine 11 and the annular insert 34 can be made the same and, as a result, the increase in the radial clearance in the boosting modes can be prevented.

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

Fig. 7 shows an annular insert 34 with ribs 81 on an annular insert 34 and / or 59 on a housing 60. The use of ribs 81 and 59 intensifies cooling of the annular inserts 34. Fig. 8 shows an annular insert 34 with turbulators 82, also made on the outer surface of the annular insert 34. The turbulators 82 can be made in the form of cylinders of small size or any other shape.

Through the sleeve 34 can be installed through the sleeve 83, combined with holes 74 for the passage of air from the cavity 84 into the cavity 85. In the inner wall 62 holes 86 are made for venting air into the cavity 87 behind the turbine 11. The front part 53 and the rear part 64 of the housing 60 are interconnected via an annular insert 34 with bolts 88 mounted between through bushings 83.

On the inner surface of the annular inserts 34, a soft, easy-to-abrade coating 88, for example graphite (Fig. 11), or honeycomb insert inserts 89 (Fig. 12) are attached.

Figures 14 and 15 show the design of the air intake 52, which includes an air intake pipe 53 and a flow regulator 54, a housing 55 with a cavity 57, which is connected to the cavity 57 by openings 68. The housing 55 has two brackets 80, which it is attached with bolts 91 to the flange 92 of the housing 8 of the combustion chamber 7. The flow regulator 54 may be of any design. For example, a flow controller 54 is shown in the form of a cylinder 93 with rectangular openings 90. A shaft 95 with a drive 55 is connected to the cylinder 93. The drive 55 is connected by an electrical connection 79 to the on-board computer 80 (Fig. 14) and is fixed by an arm 96 to the housing 60.

At the same time, it is expedient to increase the air velocity in the cavity 57 in comparison with the air velocity in the second circuit 33. This will increase the heat transfer rate. This result is achieved by reducing the cross-sectional area of the gap 57 in comparison with the inlet area of the intake pipe 53.

On Fig shows a diagram of the change in air flow for cooling the rotor of the turbine 11 pos.97 depending on the temperature in front of the turbine - Tg, from which it follows that the flow rate of air 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 may be linear, for example, as shown in FIG. On Fig 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 98 ... 100. Positions 101 ... 103 show the real change in air flow rate g2.

TURBINE WORK

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 δ0 has a calculated value, and in the afterburner (maximal) mode, the radial clearance δ 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 parts 62 and 63 of the housing 60 and the annular insert 34, which will prevent an increase in the radial clearance.

Passing through the high pressure pipe 50 through the flow regulator 51, the 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 51 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. Moreover, the high pressure pipe 50 can only be connected to the outlet of the compressor 6 (i.e., beyond its last stage, otherwise pressure cooling air will not be enough to cool the perforated nozzle apparatus 26 and the perforated working blades 28 of the turbine 11.

The cooling air from the second circuit 33 passing through the air intake 52 and the flow control 54 enters the annular collector 47, then through the sleeves 46 into the cavity 45 and then through the holes 66 in the cavities 47 and 49 and cools the housing 38 ... 40 and the annular insert 34. When this, in order to maximize the efficiency of the system, it is necessary to use relatively "cold" air, which should be taken due to the intermediate stage of the compressor 12 (figure 1). The flow regulators 51 and the actuators 55 of the regulators 54 are connected by electrical connections 80 to the on-board computer 79 to control the flow of cooling air g1 and g2 (11 and 12)

The use of heat-accumulating material 36 equalizes the thermal inertia of the rotor 23 and stator 24. With an increase in the radial clearance, the radial clearance sensors 78 record this fact, and the on-board computer 79 sends a command to the drive 55 of the flow controller 54 to increase the flow of cooling air through the communication channel 80. 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 application of the invention allowed:

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

2. To ensure 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 systems for regulating the radial clearance 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 (5)

1. A turbine of a dual-circuit gas turbine engine, comprising at least one cooled stage with a nozzle apparatus with cavities above and below it, and at least one turbine rotor with a cooled impeller and a spin apparatus in front of it, and also a turbine stator containing turbine housing, an annular insert above the turbine impeller and a radial clearance control system, characterized in that the turbine housing is made of two parts: front and rear, an annular insert is installed between them, the stator is cooled air of the second circuit, while the supply system of the cooling air stator is made in the form of an air intake installed in the second circuit, a flow regulator with a drive, an air intake case, and a perforated casing is installed in the air intake case, and also contains an on-board computer and radial clearance sensors, a regulator drive flow sensors and radial clearance sensors are electrically connected to the on-board computer. 2. The turbine of a dual-circuit gas turbine engine according to claim 1, characterized in that the annular insert is made hollow. 3. The turbine of a 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 turbine of a dual-circuit gas turbine engine according to claim 1, characterized in that the ribs are made on the annular insert and / or turbine housing. 5. The turbine of a dual-circuit gas turbine engine according to claim 1, characterized in that radial holes are made in the turbine housing in front of the annular insert, and through axial bushings are made in the annular insert.

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