The compressor of gas turbine engine

 

The invention relates to compressors of gas turbine engines for aircraft and ground application. The technical problem on which the invention is directed, is to improve reliability due to the gradual draining of oil in the flow-through portion of the impeller of the first stage. The essence of the technical solutions is that in the compressor of the gas turbine engine with an annular cavity formed by the blade and the disk rim of the first impeller, and an annular radial rib mounted on the inner surface of the rim of the disk from the side of the entrance to the first impeller, according to the invention the annular cavity is connected with a flowing part of the impeller radial holes formed between the slots under the blades along the middle lines of the projections of the disk, while the number of holes is 2-10, and the ratio D/d lies in the range 50-500, where D is the diameter of the sleeve of the first impeller inlet to the compressor; d - diameter radial holes. 3 Il.

The invention relates to compressors of gas turbine engines for aircraft and ground application. Known compressor of a gas turbine engine, in which Barakzai is the reduced reliability due to the additional load of the rim of the disc by centrifugal forces from the balancing weights.

Closest to the claimed is the compressor of the gas turbine engine, on the inner side of the rim of the disk which is made annular radial force edge for setting the balancing weights [2].

A disadvantage of the known designs adopted for the prototype is the low reliability of the compressor due to an imbalance of its rotor when the oil in the annular cavity formed by the rim and the blade disk and the annular radial rib on the inner surface of the rim of the disk. Some transitional modes of operation of the compressor, for example when you reset the gas, due to the reverse pressure drop some oil through the labyrinth seal oil cavity falls to the canvas disk first stage, and then under the action of centrifugal force is accumulated in the annular cavity. After stopping the engine oil accumulated in the lower part of the rim of the disk and consueta, causing an imbalance of the rotor of the compressor and higher vibration during engine operation, which can cause damage to bearings and parts bearings.

The technical problem on which the invention is directed, is to improve reliability due to the gradual draining of the oil compressor gas turbine engine with an annular cavity, formed by the blade and the disk rim of the first impeller, and an annular radial rib mounted on the inner surface of the rim of the disk from the side of the entrance to the first impeller, according to the invention the annular cavity is connected with a flowing part of the impeller radial holes formed between the slots under the blades along the middle lines of the projections of the disk, while the number of holes is 2...10, and the ratio D/d lies in the range 50...500, where D is the diameter of the sleeve of the first impeller inlet to the compressor; d - diameter radial holes.

In case of contact with oil in the annular cavity drain oil in the setting of the compressor of the gas turbine engine must be slow, i.e., stretched in time, as instantaneous release of oil in an air path at the entrance of the compressor will cause a high concentration of smoke in the air taken from the compressor to the pressurization of the aircraft cabin. This is due to the fact that the temperature of the air output from the compressor in modern gas turbine engines is much higher than the combustion temperature and coking oil.

After stopping of the engine of the separated oil will accumulate in the lower part of the ring on which ATOC eliminated due to the fact, what in the inventive design of the annular cavity between the sheet and the rim of the disk of the first impeller is connected with the flowing part of the impeller radial holes, which are made between the slots under the blades along the middle lines of the projections of the disk. In such a constructive implementation of the oil through the radial holes will gradually drain into the flowing part of the impeller of the first stage, evaporating and burning in the compressor without the formation of high concentrations of smoke. Instantaneous release of oil in an air path at the entrance to the compressor, on the contrary, will cause a high concentration of smoke in the air taken from the compressor to the pressurization of the aircraft cabin. This is due to the fact that the temperature of the air output from the compressor in modern gas turbine engines is much higher than the combustion temperature and coking oil.

For the radial holes between the slots under the blades along the middle lines of the projections of the disk space necessary for minimum reserves diminish the strength of the projections of the disk.

The diameter d of the radial holes, which determines the flow rate of the oil flowing in the setting of the compressor depends on the flow of air through the compressor, i.e., the dimension kompresory the path of the compressor at its input with the formation of smoke high concentration, also will be needless to diminish the protrusion of the disk between the slots for the blades.

When D/d>500 blockage of radial holes polluting particles entering the compressor inlet air.

The number of radial holes should be not less than 2, because otherwise the possible imbalance of the rotor of the compressor, and not more than 10 because of the possibility of plugging and coking holes due to their small diameter.

In Fig. 1. shows a longitudinal section of the compressor of the gas turbine engine of Fig. 2. shows the element I in Fig.1 in an enlarged view. In Fig.3. shows the section a-a in Fig.2.

The compressor 1 gas turbine engine consists of a stator 2 and a rotor 3 that is installed in the stator 2 on the radial bearing 4 on the input side 5 of the compressor 1 and the radial-thrust bearing 6 on the output side 7 of the compressor 1.

On the shaft 8 of the rotor 3 of the compressor 1 is installed disks 9, and the rim 10 of the first stage disc 11 from the inlet 5 to the compressor 1 has radial holes 12 of diameter d, connecting the annular cavity 13 formed by the fabric 14, the rim 10 and the annular radial rib 15 of the disc 11 with a flowing part 16 of the impeller 17 of the first stage, the sleeve 18 which is made of diameter D with a hundred is e shown).

Radial holes 12 of diameter d with the aim of minimum reserves diminish the strength of the projections 19 of the disk 11 is performed on the middle line 20 of each protrusion 19, between the slots 21 under the rotor blades 22.

The radial bearing 4 located at the side of entrance 5 to the compressor 1, is located in the oil cavity 23, which is separated from the air cavity 24 at the entrance to the impeller 17 using labyrinth seals 25 and 26 with the intermediate air cavity 27.

Does this device as follows.

When the compressor 1 of a gas turbine engine in some transitional modes, for example, when the discharge gas between the oil cavity 23 and the air cavity 24 there is a reverse differential pressure (pressure in the cavity 23 is larger than the cavity 24), causing the oil particles through labyrinth seals 25, 26 and an intermediate air chamber 27 are received in the air chamber 24 at the entrance to the impeller 17, where the centrifugal forces are separated in the annular cavity 13 formed annular radial rib 15 for installation of the balancing weights, and the rim 10 and the blade 14 of the first stage disc 11.

When the engine is running and after the engine stop occupationand the evaporating and burning in the compressor without the formation of smoke.

Sources of information 1. C. A. Vunow. Construction and design of aircraft gas turbine engines. M: mechanical engineering, 1981, page 76, Fig.3.18.

2. C. A. Vunow. Construction and design of aircraft gas turbine engines. M: mechanical engineering, 1981, page 66, Fig.3.10.

Claims

The compressor of the gas turbine engine with an annular cavity formed by the blade and the disk rim of the first impeller, and an annular radial rib mounted on the inner surface of the rim of the disk from the side of the entrance to the first impeller, characterized in that an annular cavity is connected with a flowing part of the impeller radial holes formed between the slots under the blades along the middle lines of the projections of the disk, while the number of holes is 2...10, and the ratio D/d lies in the range 50...500, where D is the diameter of the sleeve of the first impeller inlet to the compressor; d - diameter radial holes.

 

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