RU2439337C2 - Rotor assembly for gas turbine engine - Google Patents

Abstract

FIELD: machine building.

SUBSTANCE: rotor assembly for gas turbine engine includes a pair of metal discs with central hole for rotor shaft and many accurately processed through holes for a pin, uniformly distributed throughout the length of two concentric circles, and blade assembly arranged between metal discs. Blade assembly includes many double blades from composite laminated material, each of which has a pair of aerodynamically shaped feather parts and common root part in the form of U-seat the shape of which provides its seating in the pin. Blades are combined with two concentric load-carrying rings of the first and the second row relative to the shaft, and inter-blade inserts from composite laminated material. Inter-blade inserts are arranged on surface of load-carrying ring of the second row between each pair of feather parts of double blade and adjacent double blades. Inter-blade insert has box form with a hole for a pin on one end and aerodynamic shaping on the other end, constituting together with feather part of the blades the aerodynamic working surface of rotor assembly. Metal discs and blade assembly is attached to rotor assembly by means of pins of the first and the second row. Pins enter the holes of both discs and their ends protrude in order to fix discs on them. Pins of the first row enter the holes of the first circle of discs through the seat of double blade and conjugate its inner surface and load-carrying ring of the first row. Pins of the second row enter the holes of the second circle of discs through the hole of inter-blade inserts and conjugate feather parts of blades and load-carrying ring of the second row and are supported on the surface of load-carrying ring of the second row.

EFFECT: invention allows increasing service life and improving reliability of rotor of gas turbine engine.

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Description

The invention relates to turbomachines for gas turbine engines, and more particularly relates to a rotor assembly for a gas turbine engine equipped with a blade system.

Known wheel of a gas turbine engine equipped with a system of blades (RF patent No. 2333366, IPC F02C 1/00, publ. 2008). The blades are held in the peripheral annular groove of the impeller by means of a hammer-shaped attachment points. Each blade further comprises a platform whose radially outer surface limits the gas flow channel.

Known rotor of a fan of a turbojet engine (RF patent No. 2221169, IPC F04D 29/26, publ. 2004), containing many removable fan blades mounted on the rim of the disk. Each of the fan blades contains the root part of the blade, which is installed in the groove in the rim of the disk. Each of the blades is held in the groove with a key located under each root part of the scapula, and means for holding the keys and root parts of the scapula in the grooves. The dowels have a thickening at one of their ends, and the aforementioned means comprise a flange attached to the disk by means of fasteners, which contain many studs parallel to the axis of rotation of the given disk.

A known node of the rotor disk of a gas turbine engine containing a number of annular hubs, a plurality of dovetail grooves spaced around the circumference, a plurality of fan blades and a plurality of blade spacers (RF patent No. 2281403). The fan blades have dovetail shanks arranged in dovetail shaped grooves. Each of the spacer inserts is located inside each of the dovetail grooves between the bottom wall of the dovetail groove and the axially extending dovetail shank of the dovetail.

A known rotor assembly comprising a rotor impeller having a plurality of grooves and a plurality of blade assemblies, each of which has a blade and a dovetail section engaged with a corresponding impeller groove (RF application No. 2007134116, publ. 2009).

Known designs of rotor assemblies do not have high strength, since the material in the fastening devices (fittings) and blades, while holding the blades, experiences a load (works) on shear.

The basis of the invention is the task of increasing the durability and reliability of the rotors of light turbomachines.

The technical result is to increase the durability and reliability of the rotors of light turbomachines due to the creation of a structure in which the composite material in the attachment devices while holding the blades works for shear. Another technical result is improved torque transmission.

The problem is solved in that the rotor assembly for a gas turbine engine contains a pair of metal disks with a central hole for the rotor shaft and a plurality of precision machined through holes for the pin, uniformly distributed along the length of two concentric circles, and a blade assembly containing a plurality of twin blades of a layered composite material placed between metal disks, each of which has a pair of aerodynamically shaped feather parts and a single root part in the form of a U-shaped saddle, the shape of which ensures its landing in a pin, united by two concentric bearing rings of the first and second row relative to the shaft, interscapular inserts of layered composite material, placed on the surface of the bearing ring of the second row between each pair of feather parts of a double and adjacent twin blades moreover, the interscapular insert has a box shape with a pin hole at one end and aerodynamic profiling at the other end, which together with the rim part of the blades is the aerodynamic working surface of the rotor assembly, the metal disks and the blade assembly being fastened to the rotor assembly with pins of the first and second row, while the pins enter the holes of both disks and protrude by the ends to fix the disks on them, the pins of the first row enter these holes through the saddle of the double blade and match its inner surface and the bearing ring of the first row, and the pins of the second row through the hole of the interscapular inserts and match the feather parts of the blades and the bearing ring of the second of the second row and rest on the surface of the bearing ring of the second row.

In the future, the invention is illustrated by the description and drawings, in which

figure 1 - turbine wheel of a gas turbine engine according to the invention in cross section, General view,

figa is a fragment of a section of a turbine wheel (view A) on an enlarged scale,

figure 2 is an axial section of the turbine wheel of figure 1 according to the invention,

figure 3 - interscapular insert of the turbine wheel (side view),

figure 4 - interscapular insert of the turbine wheel (top view),

5 is a turbine wheel disk.

The rotor assembly for a gas turbine engine (Fig. 1) contains a pair of metal disks 11 with a central hole 12 (Fig. 5) under the rotor shaft 9. Each disk 11 has a plurality of precision machined through holes 13 for the pin. Through holes 13 are made and evenly distributed along the length of two concentric circles in two rows.

The rotor assembly also comprises a blade assembly located between the metal disks 11.

The blade assembly contains a plurality of twin blades 14 of a layered composite material.

Each such twin blade 14 has a pair of aerodynamically shaped feather parts 5 and a single root part 6 in the form of a U-shaped saddle. The shape of the root part 6 is adapted for the pin 1 placed therein.

Twin blades are combined by two concentric bearing rings of the first and second row (relative to the rotor shaft 9 installed in the central hole 12) - the bearing ring 3 of the first row and the bearing ring 4 of the second row.

The blade assembly also contains a plurality of interscapular inserts 7 of a laminated composite material.

The interscapular insert 7 (FIGS. 3 and 4) has a box shape with a hole 10 for a pin at one end. At the other end, the surface 8 of the interscapular insert is aerodynamically profiled and together with the feather part of the blades forms the aerodynamic working surface of the rotor assembly. Each insert 7 is placed on the surface of the bearing ring 4 of the second row between each pair of feather parts 5 of the twin blades, as well as between the feather parts of adjacent twin blades. The size of the insert 7 provides the necessary gaps between the branches of the blade and adjacent blades, and the interscapular inserts 7 located both inside the twin blades and between two adjacent twin blades have the same geometric parameters.

The interscapular inserts 7 provide a precise outer gap between adjacent twin vanes 11 and inside the vanes and an aerodynamically shaped outer surface.

The metal disks 11 and the blades are fastened to each other in a monolithic rotor assembly by pins of the first and second row entering the holes of both disks, and the pins 1 of the first row enter the holes 13 of the first concentric circle (relative to the central hole 12) through the saddle 6 of the twin blades and mate its inner surface and the bearing ring 3 of the first row, and the pins 2 of the second row enter the holes 13 of the second concentric circle through the holes of the interscapular insert 7 and match the feather parts of the blades and bearing e ring of the second row and rest on the surface of the bearing ring 4 of the second row.

The disks 11 (figure 5) contain through holes 13 made under the pins 1 and 2, respectively, of the bearing rings of the first 3 and second 4 rows.

During the operation of the gas turbine engine, the gas flow passing through the engine at the final stage enters the profiled parts 5 of the blade 11. When the pressure of the working gas flow to the surface of the twin blades 14 is transmitted. The twin blade 14 operates under a load of centrifugal forces, which is perceived by the bearing ring 3, the pins 1 of the first row and the metal disk 11 in the holes 13. Then the torque is transmitted to the metal disk 11, which, in turn, transmits the torque to the shaft 9, which inserted into the hole 12. The interscapular inserts 7 absorb the load with the support ring 4 and pins 2 of the second row. The lead torque is transmitted to the shaft 9, inserted into the hole 12.

In the claimed design, the fiber of the composite material of the inserts and blades operates on a cut.

In the known constructions of non-metallic turbine impellers, dovetail shank material fixed in the locking groove of the disk operates on an interlayer shift. The shear strength characteristics of layered composite materials are low. They vary from 1 to 3 kg / mm ² depending on the temperature level. Damage to the shank is the reason for the early loss of the bearing capacity of the blades. In the proposed twin blades, which bend around the pins of the first order and are fixed with their help, the material works in a shear in the transverse direction. The characteristics of composite materials for a slice are 5-10 times higher than the characteristics of interlayer shear strength. In addition, the dovetail type shank is located in an area closer than the pin 1 of the twin blades to the profile of the blade, where the temperature is the highest. Therefore, in previous designs, the effect of temperature on the strength characteristics is quite large. Thus, the proposed design allows to obtain a product with enhanced strength characteristics.

One of the difficult problems in the design of impellers made of non-metallic materials is the transmission of torque from the blades to the shaft 12. In the proposed design, the transmission of torque is improved due to the fact that the torque is transmitted to the shaft using two metal disks 11, mounted on the protrusions ( splines) of pins in precision machined holes 13.

A large number of pins unload each pin. However, in order for all the pins to be equally and simultaneously loaded, a sufficiently accurate fit in the holes 13 of the disks 11 is necessary.

A good fit is achieved by the precision machining of the holes 13 in the disks 11 and the precise execution of the pins.

It should be noted that when the blades rotate in the form of a rigid fit of interscapular inserts, the torque from the blades is transmitted to these inserts, which, in turn, transmit torque to the 2 row pins, thereby partially unloading the 1 row pins from the loads. The combination of these measures allows us to solve the problem with obtaining the specified technical result.

Claims (1)

A rotor assembly for a gas turbine engine, comprising a pair of metal disks with a central bore for the rotor shaft and a plurality of precision machined through holes for the pin, uniformly distributed along the length of two concentric circles, and a blade assembly located between the metal disks, containing a plurality of twin blades of a layered composite material , each of which has a pair of aerodynamically shaped feather parts and a single root part in the form of a U-shaped saddle, the shape of which It ensures its landing in a pin, united by two concentric bearing rings of the first and second row relative to the shaft, and interscapular inserts of layered composite material, placed on the surface of the bearing ring of the second row between each pair of feather parts of a double and adjacent twin blades, and the interscapular insert has a box-shaped a shape with a pin hole at one end and aerodynamic profiling at the other end, which together with the feather part of the blades are aerodynamic the working surface of the rotor assembly, the metal disks and the blade assembly being fastened to the rotor assembly with pins of the first and second row, while the pins enter the holes of both disks and protrude by the ends to fix the disks on them, the pins of the first row enter the holes of the first annular circumference of the disks through the saddle of the twin blades and match its inner surface and the bearing ring of the first row, and the pins of the second row into the holes of the second annular circle through the hole of the interscapular inserts and match the feathers The main parts of the blades and the bearing ring of the second row are supported by the surface of the bearing ring of the second row.

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