RU2499894C1 - Bypass gas turbine engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises compressor with at least one stage, combustion chamber with fire tube, turbine including at least one cooled stage with distributor with cavities above and under it. Turbine rotor incorporated cooled impeller and swirler arranged there ahead. Turbine stator comprises at least two turbine casings with cavities there between and radial clearance adjustment system including ring insert arranged above the impeller. Cavity above distributor is communicated via air bleed pipeline including flow rate controller with compressor outlet. One of said cavities between turbine casings is communicated via pipeline including second flow rate controller with compressor mid stage. Radial clearance adjustment means comprises radial clearance transducers and onboard computer electrically connected with flow rate controllers and radial clearance transducers.

EFFECT: efficient control over radial clearance, higher takeoff and augmenter thrust, efficiency and reliability.

3 cl, 16 dwg

Description

The group of inventions relates to engine building, including aircraft and stationary gas turbine engines (GTE), having the bottom of the circuit, and can be used in aircraft building, shipbuilding, gas pumping stations and for peak power plants as a drive for an electric generator designed to generate electricity .

Known turbine for a gas turbine engine according to the patent for invention No. 2435039, IPC F01D 11/24, publ. 04/27/08, the turbine housing includes a radial wall and contains on the side of its inner surface a support for fixing 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 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 the Russian Federation No. 2435039, IPC F01D 11/04, a prototype 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 the take-off and afterburning mode, increasing the efficiency and reliability of the turbine.

The solution of these problems was achieved in a gas turbine engine containing a compressor having at least one stage, a combustion chamber containing a heat pipe installed with a gap relative to the body of the combustion chamber, a turbine containing at least one cooled stage with a nozzle apparatus with strips above it and beneath it, and a turbine rotor with a cooled impeller and a spinning apparatus in front of it, as well as a turbine stator containing at least two turbine bodies with cavities between them and a radial regulation system a gap, containing an annular insert above the impeller of the turbine, in that according to the invention, the cavity above the nozzle apparatus is connected by an air intake pipe containing a flow regulator to the compressor outlet, and one of the cavities between the turbine bodies is connected by a pipe containing a second flow regulator, with intermediate compressor stage, the radial clearance control system contains an on-board computer and radial clearance measurement sensors, a flow regulator, valve actuators and measurement sensors Dial clearance are connected by electrical connections.

The invention is presented in the drawings (figures 1-16), 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 A,

- figure 4 shows a section bB,

- figure 5 shows the installation diagram of the annular insert,

- Fig.6 shows a hollow insert,

- Fig.7 shows an insert with heat-storage filler,

- Fig.8 shows a view of an annular insert with holes in it,

- figure 9 shows the annular insert with ribs,

- figure 10 shows the annular insert with turbulators,

- figure 11 shows the annular insert coated with a soft abradable material, type A,

- Fig.12 shows an annular insert with panels of "honeycomb seals",

- Fig.13 shows the appearance of the air intake,

- Fig.14 shows a view In

- Fig.15 shows a diagram of the change in air flow for cooling the turbine rotor depending on the temperature in front of the turbine,

- Fig.16 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 the gas turbine engine is shown in the drawings of figures 1-16. 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 is a guide apparatus 20, 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 s working blades 28 and the disk 29. The steps 25 of the turbine 11 may be one or more than 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.

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

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 a high heat capacity and a 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, but can be applied to the other (all) stages 25 of the 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 radial clearance control means, in addition to the previously listed means, includes a swirl device 42, an internal cooling air supply pipe 43, an internal cavity 44, holes 45, an internal cavity 46 of the nozzle device 26, openings 47, an upper cavity 48 inside the annular manifold 49, a high pressure pipe 50, air 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. Each air inlet 52 has an air intake pipe 53 and a flow regulator 54. 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, 2- x up to 12 air inlets 52. The construction of air inlets 52 is shown in detail in FIGS. 13 and 14. The air inlets 52 are connected by bushings 55 to the main cavity 56. From the main cavity 56, the zduha for cooling the stator 24 of the turbine 11 (Figure 2).

The shaft 8, the disk 26 with the deflectors 30 and 31 and the impeller 29 form the rotor 23 of the turbine 11 (figure 2). The turbine 11, as previously indicated, 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 rod 66 with holes 67 for venting air (FIGS. 1 and 2) cooling the stator 24 of the turbine 11 (FIGS. 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 heat storage material 36 is, as noted above, 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, it is possible to make the heating time of the disk 29 and turbine bodies 57 ... 59 the same and, as a result, to prevent an increase in the radial clearance in the boost modes.

The main features of the turbine 11 is the presence of radial clearance measurement sensors 78 and an on-board computer 79 connected by electrical connections 80. Only one radial clearance measurement sensor 77 can be used, but this is extremely undesirable, because Sensor failure can lead to an emergency.

Figure 4 presents the Assembly of the stator 23 of the turbine 11 of the turbine engine. In the design of the stator 21 of the turbine 11 can be applied holes 81 made in the intermediate casing 59 and 82, made in the annular insert 34.

7 shows an annular insert 34 with ribs 83, the use of which intensifies the cooling of the annular inserts 34. Fig. 8 shows an annular insert 34 with turbulators 84, also made on the outer surface of the annular insert 34. The turbulators 84 can be made in the form of small cylinders size.

On the inner surface of the annular inserts 34, a soft, easy-to-abrade coating 85, for example graphite (Fig. 11), or honeycomb inserts 86 are attached (Fig. 12).

13 and 14 show the design of the air intake 52, which includes an air intake pipe 53 and a flow regulator 54, a manifold 87 with a cavity 88, which is connected by a sleeve 55 to the cavity 56. The manifold 87 has two brackets 89, which it is attached with bolts 90 to the flange 91 of the housing 8 of the combustion chamber 4. The flow regulator 54 may be of any design. For example, a flow controller 54 is shown in the form of a cylinder 92 with rectangular openings 93. A shaft 94 with a drive 95 is connected to the cylinder 92. The drive 95 is connected by an electrical connection 80 to the on-board computer 79 (Fig. 14).

On Fig shows a diagram of the change in air flow for cooling the rotor of the turbine 11 pos.96 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 consumption g2 for cooling the turbine stator 11 in three sections of the gas turbine engine operation (in boost mode 97, 98, and 99) is shown. Positions 100,101 and 102 show the actual change in air flow rate g2.

In the intermediate housing 39, holes 103 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 ... 16).

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 (maximum) mode, the radial clearance (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 warm 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 control radial clearance. 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 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. In this case, the high pressure pipe 50 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.

The cooling air from the secondary circuit 33 passing through the air intake 52 and the flow regulators 54 enters the annular manifold 47, then through the bushings 46 into the cavity 45 and then through the openings 66 in the cavities 47 and 49 and cools the casings 38 ... 40 and the annular insert 29 In this case, 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 (Fig. 1). Flow controllers 61 and 63 are connected by electrical connections 37 to the on-board computer 36 to control the flow of cooling air g1 and g2 (11 and 12)

The use of heat-accumulating material 51 equalizes the thermal inertia of the rotor 19 and stator 21. With an increase in the radial clearance, the radial clearance measurement sensors 35 record this fact, and the on-board computer 36 instructs the second flow regulator 63 to increase the flow rate of cooling air via the communication channel 37. If 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 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 radial clearance control systems without preliminary heating of the turbine engine or significantly reduce the warm-up time of the gas turbine 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 (3)

1. A double-circuit gas turbine engine containing a compressor having at least one stage, a combustion chamber containing a heat pipe installed with a gap relative to the housing of the combustion chamber, a turbine containing 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 spinning apparatus in front of it, and a turbine stator comprising at least two turbine bodies with cavities between them and a radial clearance control system, comprising an annular insert above the turbine impeller, 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, and one of the cavities between the turbine bodies is connected by a pipeline containing a second flow regulator with an intermediate compressor stage, a control system radial clearance contains an on-board computer and radial clearance measurement sensors, a flow regulator, valve actuators and radial clearance measurement sensors are connected electrical connections. 2. The dual-circuit gas turbine engine 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.

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