RU2506436C2 - Device for optimisation of radial clearances of aircraft gas turbine engine multistage axial-flow compressor - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed device for optimisation of radial clearances by compressed air bled from compressor comprises housing with blading. Compressed air flows successively through inner chambers of compressor stages. Rotor of every stage comprises multiple radial blades secured to the disc. This device comprises case fitted on the shaft under discs of at least three last stages, system of seals and slots between said discs and case, holes in said discs and case outlets. Device inlet communicates with compressor end area wherein cyclic loads at aircraft engine define maximum air heating. Seals, slots and holes are arranged to create loop-like airflows in said case from inlet along disc webs towards outlets, opposite the airflow direction in blading.

EFFECT: decreased thermal stress at rotor discs, minimized compressed air inner leaks by optimized radial clearances.

6 cl, 5 dwg

Description

The invention relates to gas turbine plants, and more specifically relates to a device for optimizing the radial clearances of a multi-stage axial compressor of an aircraft gas turbine engine.

It is known that in the compressor structure, the radial clearances in the flow path are the gap between the stator and the surface formed by the peripheral sections of the blades, and the gap between the inner surface of the guide vanes (HA) and the rotor element of the drum or ring spacer (N.I. Startsev. Design and design of a turbo-compressor GTE, Samara, publishing house STAu, 2006, pp. 24-30).

In aircraft engines, the problem is that, on the one hand, the radial clearances adversely affect the air compression in the compressor and increase with increasing operating time, and therefore they must be reduced, and on the other hand, the radial clearances should be increased, since they do not remain constant in the flight cycle of the aircraft engine, including such operating modes as starting, warming up, take-off, climbing, cruising, descent, approach, glide path, reverse thrust, stop. Radial clearances change both during the transition from one operating mode to another, and under constant flight conditions or in a steady state when the flight conditions change, and can either increase or decrease until the rotor touches the stator dangerously, which creates an emergency.

The negative function of radial clearances is also expressed in the fact that the loss of expended energy when air flows through the clearances under the influence of a pressure difference leads to a decrease in the efficiency and gas-dynamic stability of the compressor.

The selected mounting gap, guaranteeing that the rotor does not touch the stator in all operating modes, does not provide optimal maximum radial clearance values in cruise mode, the mode with the highest operating time, where a low specific fuel consumption is required.

In modern gas turbine engines with high pressure compressors, which include multi-stage axial compressors, gap optimization is a significant problem.

A known system for regulating the radial clearance in the flow part of a turbomachine (patent for PM No. 87213, 2009), containing a conical movable element with holes communicating with the flow part of the turbomachine, is located in the stator housing, the boost cavity formed by the installed bellows in the stator housing is hermetically sealed connected to the stator housing and the conical movable element. When changing the radial clearance, the pressure behind the conical movable element and in the bellows cavity changes, causing the conical element to move towards the restoration of the specified value of the radial clearance. The loss of energy expended when air flows through the gaps under the influence of the pressure difference leads to a decrease in the efficiency and gas-dynamic stability of the compressor

A device is known for optimizing the clearance between the blades and mating elements in a multi-stage axial compressor of a gas turbine power plant (US Pat. No. 4,795,307 of 1989). The device thermally controls the compressor rotor disks with compressed air taken from the compressor. Hotter air comes from the compressor flow path and from the gap between the rotor and stator after the last stage. Further air movement through the steps of the compressor flow path is organized in parallel: through holes in the disks and special structural elements for supplying air from one inter-disk cavity to another and through radial holes in the cylindrical parts of the disks for air supply across the disks of the disks. The thermal management of the gaps is carried out by adapting the local temperature to the temperature of the rim of the rotor disks. This patent relates to a gas turbine power plant. Compressors of aircraft engines due to differences in design, modes and operating parameters are significantly different from compressors for engines of steam turbines of power plants.

A device is known for ventilating a rotor of a gas turbine engine, which contains at least two rotors between which a cavity is formed (US patent 7,775,764). An opening for cooling air inlet is made in the first rotor, and a cooling air outlet in the second rotor, so that the cooling air passes radially between the disks. This device provides ventilation of the inner cavity of the rotor to prevent overheating of the rotor discs. It does not intensify heat transfer along the entire height of the rotor discs and does not optimize radial clearances.

A device for controlling the clearance of a multi-stage axial compressor of a gas turbine aircraft engine during engine operation, including maximum thrust and low thrust modes (US patent 4,576, 547, publ. 1986).

The device comprises a unit for taking compressed air from one of the stages in the middle part and from another of the stages downstream of the air flow through the stages, a unit for supplying selected air to the area under the rotating disks of the rotor of the axial compressors, which includes at least one hollow a stator blade, an anti-vortex tube extending from the inner diameter of the hollow stator blade radially inward to the axis of rotation, a control module for selecting a sampling point and the specified supply of compressed air from the flowing part inward torus to heat said disks so that the disks expand in the direction of the air seal and close the gap between the air seal and the peripheral sections of the blades during low thrust.

The well-known technical solution organizes the supply of compressed only under the disks, which does not provide an intense thermal effect on the entire disk blade and therefore does not allow changing the radial clearances to the cyclic loads of the aircraft engine and achieve uniform thermal state of the disks.

In transient GTE modes in a multistage axial compressor, the time to achieve a uniform temperature state of the rotor discs remains significant compared to the duration of the regime, which creates thermal stresses of the rotor discs.

The basis of the invention is the task of increasing the resource and increasing the efficiency of the compressor of an aircraft gas turbine engine (GTE).

EFFECT: reduction of thermal stresses of rotor disks and minimization of internal leaks of compressed air by optimizing changes in radial clearances that are adequate to the operating regimes of the flight cycle of an aircraft engine, which increase the resource and increase the efficiency of a gas turbine compressor.

The problem is solved in that a device for optimizing the radial clearances of a multi-stage axial compressor of an aircraft gas turbine by heating with compressed air drawn from a compressor mounted on a shaft and containing a housing with a flow part, where the compressed air passes through the internal cavities of the compressor stages, and the rotor of each stage includes many radially arranged blades mounted on the disk, contains a casing placed with a gap under the disks of at least three ice stages of the compressor, and a system of seals, slots and holes in the indicated disks and the casing, while the entrance to the device is connected to the region of the end of the compressor, where cyclic loads on the aircraft engine determine the maximum heating of the air, and seals, slots and holes are placed so that they create a loop-like the flow of the specified heated air in the casing from the entrance along the blades of the disks to the exit of the casing in the direction opposite to the direction of the air flow in the flow part in order to heat up the disk web and optimize ix radial clearances adequate mode of operation of the flight cycle. The system of seals, slots and holes is designed so that the total hydraulic resistance of the specified system is less than the hydraulic resistance of the compressor seal disk.

For a 7-stage compressor, the device comprises a casing located under the seventh, sixth and fifth stage disks, the system has a group of holes in the shelves of the seventh last disk, a group of holes in the sixth stage disk and in the casing for release, a gap between the casing and the hub of the seventh disk, and calculated so that the total hydraulic resistance of the specified system is less than the hydraulic resistance in the seal of the seal disk, while the entrance to the device is associated with a leakage gap between the rotor and the stator after the last seventh enyu compressor.

The holes of each group are evenly spaced around the circumference. The holes in the shelves of the seventh disk can be located on two concentric circles. The casing is mounted on the compressor shaft, which contains holes evenly spaced around the circumference, communicating the internal cavity with the cavity of the reduced air pressure in the compressor.

The invention is further explained in the description and graphic material, in which Fig. 1 shows a longitudinal section through the flow part of a multistage axial compressor of a gas turbine aircraft engine (circuit diagram);

figure 2 shows a view To figure 1 on a larger scale;

figure 3 shows a view of M of figure 1 on a larger scale;

figure 4 (a, b, c, d) shows the arrangement of groups of holes in the shelves of the disk of the last stage, holes in the penultimate disk, in the compressor shaft and in the casing of the device, respectively, sections HH of figure 2, P-P and PP of FIG. 1 and CC of FIG. 3;

figure 5 shows graphs of the calculated changes in the radial clearance of the last stage of a 7-stage axial compressor with and without the device according to the invention (for comparison).

Figure 1 shows in section a schematic diagram of a multi-stage axial compressor 1 of a gas turbine aircraft engine containing a flow part 2, where the incoming air is compressed, sequentially passing through the compressor stage. In Fig. 1, a multistage axial compressor contains 7 stages, but their number may be different depending on the purpose of the compressor.

We call the 1-2 stages along the air flow in the compressor flow path the first, the 3-4 steps - the middle, and the remaining 5-6-7 steps - the last compressor steps.

Compressor 1 contains a rotor (rotating part) and a stator (fixed part).

The rotor contains many radially arranged blades 6 of each stage, mounted on disks 3 and rotating during operation of the compressor.

The stator contains many blades 5 of each stage, mounted on the annular housing 4 of the stator and stationary during operation of the compressor.

In order to avoid grazing of the moving blades 6 and the fixed blades 5 about the counterparts during installation, the mounting values of the radial clearances are provided - a gap 7 between the stator housing 4 and the surface formed by the peripheral sections of the working blades 6, and a gap 8 between the inner surface of the stator blades 5 and the ring spacers 9 rotor.

The first steps are made of alloy materials, as They are located in a zone of low temperatures; therefore, they do not experience significant thermal expansions and do not need to regulate radial clearances.

The middle and last stages in modern gas turbine compressors are in the high temperature zone and their elements - rotor blades and disks, blades and stator housing are made of refractory materials with close values of thermal expansion coefficients. Especially the last stages experience significant thermal expansion and need to regulate radial clearances, the value of which varies in accordance with cyclic loads in accordance with the modes of operation of the aircraft engine.

Shown in figure 1, the 7-stage compressor 1 contains the gaps 7 and 8. However, it is possible that the compressor may have a different design, where only the gap 7 will be present.

The compressor is equipped with a radial clearance optimization device, which, according to the invention, comprises a casing 10 located under the disks of at least the last three stages of the compressor, in FIG. 1 these are 5-6-7 stages. The casing 10 is mounted on the shaft 15 of the compressor rotor.

The device also contains (Fig. 1, Fig. 3) a system of seals for slots and holes: a seal 14 between the casing 10 and the disks 3 of the 6th and 5th steps, a slit 16 between the casing 10 and the disks 3, a group of holes 17 in the disk shelves 3 of the last 7th stage, a group of holes 18 in the disk 3 of the penultimate 6th stage.

The input of the device into the casing 10 is connected with the region of the end of the compressor, where cyclic loads on the aircraft engine determine the maximum heating of the air in the compressor. For example, this input may be associated with a leak in the axial clearance 11 between the rotor and stator elements (behind the disk 3 of the seventh last stage of the compressor, FIG. 2) and the cavity 12 at the end of the compressor.

The exit from the casing 10 contains a group of outlet openings 13 (Fig. 1, Fig. 3 and Fig. 4) and is located in the area of the device opposite from the entrance.

The areas of bottlenecks, such as openings 17, openings 18, outlet openings 13, slit 16, are designed so that their combined hydraulic resistance is less than the hydraulic resistance of the seal disk 20 (in seal 19).

When the compressor is operating, compressed air from the leak through the axial clearance 11, which is most heated during the current operation of the aircraft engine, enters the cavity 12. Then this air, under the influence of the pressure difference created by the sealing element 19 between the cavity 12 and the cavity 21 with a lower air pressure, enters into the interdisk cavity between the seal disk 20 and the disk 7 of the 7th stage, then moves from top to bottom through the holes 17 in the shelves along the webs of the disks 3 and 20, passes through the slot 16 between the casing and the hub of the last disk to the seal 14 and rises up along the canvases of disks of the 6th and 7th steps, through the holes 18 in the disk 3 of the 6th step enters the next interdisc cavity, goes down along the canvases of the disks 3, exits through the outlet openings 13 and then through the openings 22 (FIG. 1, FIG. 4) in the shaft of the rotor 15 enters the cavity 21 of the compressor with a lower pressure, moving as a whole in the direction opposite to the direction of the main air flow in the compressor. Thus, the air moves in a loop, blowing the rotor discs.

Loop-like moving warmer / colder, depending on the operating mode, compressed air intensifies heat transfer, resulting in a uniform change in the temperature state of the disk web, respectively, increase / decrease in temperature leads to an increase / decrease in the thermal expansion of the disk and other contacting elements, which reduces / increases radial clearances, optimizing them throughout the flight cycle of an aircraft engine adequately to the modes: start, warm-up, take-off, climb, cr ysersky flight, descent, approach to landing, reverse thrust stop.

For example, when the engine enters the take-off mode after warming up with idle gas, the compressor rotor speed increases, thereby increasing the work in the compressor to the main air flow in the flow part 2. The work is aimed at increasing the pressure and air temperature, which reaches the compressor outlet maximum temperature of about 650 ... 750 ° C.

In accordance with the foregoing, there is a loop-like movement along the inter-disk cavities and intensive heat exchange between the hot air and the disks, as a result of which the temperature of the disks is equalized along the entire length and the thermal expansion decreases, which reduces the radial clearances - both the clearances 7 between the surfaces of the peripheral sections of the working blades and the stator, and the gaps 8 between the annular spacer of the rotor and the surface formed by the internal sections of the guide vanes.

In the case of the engine switching from cruise to low mode, the compressor rotor speed is reduced, thereby reducing the work in the compressor to the main air flow in the flow part. There is a decrease in pressure and air temperature along the flow part of the compressor, which at the outlet of the compressor reaches a maximum temperature of about 350 ... 450 ° C. With a decrease in air temperature, cooling similarly occurs, a decrease in thermal expansion, which reduces the indicated radial clearances.

Figure 5 shows the calculated changes in the radial clearance 8 between the surface formed by the peripheral sections of the rotor blades 6 of the 7th stage of the compressor and the stator housing 4 during the flight cycle of the aircraft engine including: start, warm-up, take-off, climb, cruise, decrease , approach, reverse thrust, stop with the device according to the invention with a dashed line, and changes in the same clearance under the same conditions without such a device (as exemplary for comparison and explanation of the achieved result) is solid iey.

The interval from zero to point A is the interval of operation of the compressor from starting the engine to reaching low gas. At this interval, the change in radial clearance is insignificant, not taken into account.

Interval A-B (from point A to point B) is the interval of operation of the compressor where the engine warms up. Although during heating the change in the radial clearance is also insignificant, a slight decrease in the radial clearance is already observed when using the radial clearance optimization device.

Interval B-B (from point B to point C) is the interval of the compressor in the take-off mode, there is a rapid change in the temperature of the air flow in the compressor flow path and, as a result, the different times of heating of the stator bodies and rotor disks begin to influence . It is important to note that for the group of the last compressor stages, the radial clearance value is significantly affected by the thermal expansion of the stator housings and compressor rotor disks, and the effect of disk deformation under the influence of centrifugal forces is insignificant for the group of last stages. Figure 5 shows that in the interval BB, the radial clearance is significantly reduced by about 95%.

The interval V-G (from point B to point G) is the interval of operation of the compressor during climb, here the air temperature in the flow part of the compressor changes smoothly, and the duration of the interval is large. Figure 5 shows that on the interval V-G, the device according to the invention can significantly reduce the radial clearance by about 70%.

The G-D interval (from point G to point D) is the interval of compressor operation in cruise mode, here the air temperature in the compressor flow section practically does not change and the interval is long. Figure 5 shows that although the radial clearance is small and practically unchanged in the GD interval, the device according to the invention allows to reduce the radial clearance.

The interval DE (from point D to point E) is the interval of operation of the compressor during descent and flight in a circle. Here, the air temperature in the flowing part decreases rapidly, the stator bodies are cooled, and their thermal expansion decreases. In the absence of the device according to the invention, the rotor disks do not have enough time to cool down, there is a sharp decrease in radial clearances in the group of the last compressor stages, which can lead to contact of the blades and damage to the compressor. It is this mode of operation that is decisive when choosing the compressor mounting clearances. In Fig. 5, it can be seen that in the interval DE, the device according to the invention allows to increase the radial clearance, which reduces the risk of dangerous impact of the rotor against the stator and the risks of an emergency.

The interval Е-Ж (from point Е to point Ж) is the interval of operation of the compressor when flying on a glide path, landing, reverse traction, taxiing and engine shutdown, there is a frequent change in short operating modes of the compressor with medium and low values of air temperature in the flow part. However, in this interval, although the value of the radial clearance varies, it remains average and therefore quite effective.

Thus, there is a change in the radial clearance adequate to the flight cycle operating modes (cyclic loads of the aircraft engine), which leads to an increase in the resource and an increase in the compressor efficiency.

The invention can be used in multistage axial compressors of aircraft gas turbine engines.

Claims (6)

1. A device for optimizing the radial clearances of a multi-stage axial compressor of a gas turbine aircraft engine with compressed air discharged from a compressor containing a housing with a flow part, where compressed air passes through the internal cavities of the compressor stages in series, and the rotor of each stage includes many radially arranged blades mounted on the disk, characterized in that it contains a casing (10) mounted on the shaft (15) and placed under the disks (3) of at least the last three stages of the compressor, and a system of seals (14) and slots (16) between these disks (3) and the casing (10), the holes (17, 18) in these disks and the outlet openings (13) of the casing, while the entrance to the device is connected with the region of the end of the compressor where cyclic loads on the aircraft engine determine the maximum heating of the air, and seals, slots and openings are placed so that they create a loop-like flow of the specified air in the casing from the inlet (11) along the blade webs to the outlet openings (13) in the opposite direction to the air flow in flow part to opt ize the change in radial clearance adequately mode of operation of the flight cycle. 2. The device according to claim 1, characterized in that the system of seals, slots and openings is made so that the total hydraulic resistance of the specified system is less than the hydraulic resistance of the compressor seal disk. 3. The device according to claim 1, characterized in that the device for the 7-stage compressor contains a casing located under the seventh, sixth and fifth stage disks, the system has a group of holes in the shelves of the seventh last disk, a group of holes in the sixth stage disk and in the casing for outlet, the gap between the casing and the hub of the seventh disk, and is designed so that the total hydraulic resistance of the specified system is less than the hydraulic resistance in the seal of the seal disk, while the entrance to the device is associated with a leakage gap between the mouth rum and the stator of the last seventh stage of the compressor. 4. The device according to claim 3, characterized in that the holes of each group are evenly spaced around the circumference. 5. The device according to claim 4, characterized in that the openings in the shelves of the seventh disk are arranged in two concentric circles. 6. The device according to claim 1, characterized in that the casing is mounted on the compressor shaft, which contains holes evenly spaced around the circumference, communicating the internal cavity with the cavity of the reduced air pressure in the compressor.

Images