RU2704056C2 - Turbine of double-flow gas turbine engine with active thermal control of radial clearance in turbine, method of active thermal control of radial clearance in turbine of double-flow gas turbine engine - Google Patents

Abstract

FIELD: machine building.

SUBSTANCE: group of inventions relates to aircraft gas turbine engines and gas turbine plants, namely to mechanical devices with thermal control of radial clearance between ends of rotor blades of compressor or turbine stage and housing of gas turbine engine. Single-stage high-pressure turbine of double-flow gas turbine engine with active thermal control of radial clearance in turbine includes one cooled stage with nozzle device and turbine rotor with cooled impeller, as well as a turbine stator, having two turbine housings with cavities between them, into which compressed air is supplied due to the last stage of the high pressure compressor, and a radial clearance control system comprising an annular insert, above the turbine wheel, enveloping the turbine rotor blades with an annular radial gap and elastically and hermetically attached to the parts forming the turbine inner housing, heater enveloping annular insert with possibility of its heating, air intake, cooling air flow rate controller with drive, onboard computer and sensors, and heater. Flow controller drive and transducers are electrically connected to the on-board computer. Annular insert is made hollow and annular projections made on its lateral sides, with tension is fixed with possibility of thermal expansion in mating annular grooves of vertical wall and flange, made on inner part of outer housing of HPT, nozzle assembly with annular projections is also tightly fitted in reciprocal annular grooves of vertical wall and housing of combustion chamber, in which and in vertical wall there are holes equally distributed in circumferential direction, through which secondary air is supplied from combustion chamber in cavity above nozzle device and annular insert. Disk and working blades of the HPT wheel are also cooled by secondary air swirled by the screwing device before entering the disc web. Circular microwave heater, either resistive or inductive, consists of two separate semi-rings made each in the form of a metal housing, inside which a heating element is fixed, and each semi-ring of heater is fixed on annular insert with possibility of radial thermal expansion together with annular insert and tangential thermal expansion relative to annular insert by means of bayonet connection with it and keys located with interference tension in reciprocal slots of bayonets of annular insert and each half-ring in its average cross section. Two branch pipes are screwed into annular insert diametrically opposite with tension along tubular thread, air intake branch pipe for supplying cooling air from the second circuit into inner cavity of annular insert and branch pipe for discharge of this air into the second circuit or for other purposes. Outlet of branch pipes into the second circuit is sealed by piston rings. On each branch the support is screwed up and to this support symmetrically to the branch pipe the central supports of two springs are fixed, made in the form of the multilayered package compressed by the distributed load, typed from steel, hardened or cold-worked, polished tapes made of stainless steel, coated with wear-resistant coating, and the spring by its end supports are fixed in the second outline on the outer housing of the HPT. Air discharge branch pipe is connected to pipeline, at outlet of which there is either normally open or normally closed electro pneumatic valve, and control of radial clearances is carried out by onboard computer commands according to proposed method, or spool distributor with electromagnetic drive, designed so that position of spool adjusting air flow, at cruising mode at altitudes higher than height limit on barostat H, was recorded with the electric drive de-energized, and the heater switching on and off and adjustment of its heating intensity, opening and closing the cooling air supply from the second circuit and intensity of this supply is performed by commands of onboard computer generated by program according to signals of sensors - engine rpm sensor and barostat according to another proposed method, or sensors measuring size of radial clearance along working blades.

EFFECT: sufficiently simple, repairable design, with possible quick replacement of worn-out radial clearance control system units, with good mass characteristic, efficient design.

15 cl, 17 dwg

Description

The group of inventions relates to aircraft gas turbine engines and gas turbine installations, namely to mechanical devices with thermal regulation of the radial clearance between the ends of the rotor blades of the compressor or turbine rotor stage and the casing of the gas turbine engine.

At various flight cycle modes: start, warm-up, taxiing, take-off, climb, cruise mode, descent, approach, glide path, thrust reverse, stop, and when switching from one mode to another, and even with a constant operating mode when changing flight conditions, the radial clearances between the ends of the rotor blades and the casing of the gas turbine engine are changed.

An increase in the radial clearance by 1% (see Kuznetsov ND, Danilchenko PD, Reznik VE Management of radial clearances in turbo-compressors of aircraft gas turbine engines. - State Committee of the RSFSR for science and higher education. SSAU. Textbook . Samara, 1991. -. 109) reduces the efficiency of the HP turbine and HPC approximately 2% and increases specific fuel consumption to 7%, the coefficient reduces the dynamic stability margin ΔK _at 3%, and engines with small dimensions the flow part, e.g. engine RB. 199, for which the height of the blades at the outlet of the HPC h _l = 19 mm, an increase in the radial clearance by 1% led to a decrease in ΔK _by 8%.

The choice of the value of the radial mounting clearance, which guarantees the absence of grazing of the rotor rotor blades on the stator at all engine operating modes, does not provide the optimal maximum radial clearance values in the mode with the highest operating time (cruising mode), at which the specific fuel consumption remains economically acceptable.

To enable the selection of appropriate values of radial clearances, the stator surfaces forming these clearances are covered with a layer of abradable material, and cellular seals are used.

After a period of time, aircraft engines in operation are often subject to the formation of large gaps, both due to the abrasion of the feathers of the blades, and due to the wear of the abrasive layer.

In addition, the radial clearances can vary during transient and stationary engine operating modes due to different values of working and thermal deformations of these units at each instant of time, arising due to the difference in their elastic properties and loading conditions, rotor and stator heating rates, determined by different thermophysical properties of their materials (thermal conductivity, heat capacity), geometry, overall dimensions, various conditions of convective heat transfer when flowing around these nodes, different coefficients of the thermal line Foot extension of their materials.

In FIG. Figure 1 shows the nature of the change in thermal deformations of the rotor and stator stages and the change in the radial clearance in the flight cycle in the case typical of turbofan engines. FIG. taken from work (see Kuznetsov ND, Danilchenko PD, Reznik VE Management of radial clearances in turbo-compressors of aircraft gas turbine engines. - State Committee of the RSFSR for Science and Higher Education. SSAU. Textbook. Samara, 1991 .-- 109 s).

In transition modes, idle gas, take-off mode, cruising mode, the heating rate of the thin-walled stator is higher than the heating rate of the massive rotor and the thermal deformation of the stator in all these modes is greater than the thermal deformation of the rotor and the radial clearance between the ends of the rotor blades of the rotor and stator increases in these modes, which can cause engine thrust to fail on take-off (see the same book). During the transition from cruising to low gas and engine shutdown, the stator cooling rate is higher than the rotor cooling speed for the same reasons and the radial clearance decreases sharply and there is a danger of cutting the blades into the engine stator.

This danger is further exacerbated by the extraction of the working blades and the locking part of the disks in the field of centrifugal forces and elastic deformations of the rotor and stator under the action of dynamic and static loads.

Note that when starting an unheated engine, at the initial stage of starting, there is also a danger of cutting rotor blades into the stator (due to the elastic drawing of the feather of the blades and the locking part of the disks under the action of the centrifugal force field) and to avoid this danger it is necessary to increase the mounting values radial clearances.

In modern aircraft gas turbine engines, these dangerous situations are overcome by thermal regulation of the radial clearances between the ends of the rotor blades of the rotor and the stator. The problem of matching the thermal expansions of the rotor and stator and the related task of managing radial clearances is one of the important tasks of developing aircraft gas turbine engines and there are many examples of successful solutions to them. Therefore, a fairly detailed overview of the known solutions to these problems is given below both in the design of modern gas turbine engines and in patents.

In modern gas turbine engines, two types of thermal regulation of radial clearances in compressors and turbines are used - passive and active.

Passive regulation is carried out automatically without any feedback signals and program control of the radial clearance values in accordance with these signals.

An example of the use of passive thermal regulation is the cooling system of the internal housing of the GT90 compressor of the gas turbine engine, in which at one of the intermediate stages of the compressor the cooling air is drawn, which enters the cavity formed by the internal and external compressor casings. The inner casing has massive elements located above the working blades of each compressor stage, to which hollow segments are mounted with bolts, combined with the blades of the guide vanes and forming the inner surface of the compressor stator, covered with abrasive material in places that form radial gaps at the ends of the working blades of each stage.

The presence of massive elements in the inner casing reduces its heating rate and brings it closer to the rotor heating rate. Due to this, the speed and magnitude of the increase in radial clearances in the flight cycle during acceleration of the engine to low gas, operation on low gas, acceleration to the cruiser mode and operation on it, and the speed and magnitude of the decrease in radial clearances during the transition from the cruiser mode to low gas, are reduced at idle and engine shutdown.

This increases the efficiency of the compressor in all engine modes and reduces the wear rate of the abraded material.

The advantage of such regulation is the simplicity of its implementation, the regulation of radial clearances immediately at the required number of compressor stages, as well as the use of air cooling the inner casing and for other purposes.

The disadvantages of this regulation include the insufficient accuracy of maintaining gaps, since their size has a complex nonlinear dependence on the engine operating modes and requires a more flexible response to the temperature state of the engine. In addition, during transient conditions, the temperature and flow rate of cooling air can be ambiguous, determined by the direction of the transient mode — gain or drop of engine speed.

Cold air taken from the first stages of the compressor to cool the housing parts to regulate radial clearances is ineffective for nearby stages. Therefore, the cold air of the first stages can rationally be used only for external blowing of the body parts of the last stages of the compressor, provided they are properly sealed. External airflow is characterized by increased inertia, which degrades the performance of the system to maintain minimum radial clearances. Uncontrolled external airflow can also be ineffective because it requires a large amount of air taken from the flow part for cooling (1.5-2 percent or more). Such air costs for airflow can be acceptable only if air after airflow of the housing is reused for other purposes, for example, for cooling a turbine, jet nozzle, etc.

Among the systems with passive regulation of radial clearances is the device for optimizing the radial clearances of a multi-stage axial compressor of an aircraft gas turbine engine (see RF patent, IPC F01D 11/24, 2506436 / EG Steshakov, AN Startsev, Yu.M. Temis, V.V. Novokreschenov, D.A. Yakushev, A.A. Mishakov, S.V. Kharkovsky - Publishing Patent 02.10.2014. - http // www / freepatent.ru / patent / 2506436) with compressed air discharged from the compressor into its internal cavity, where it successively passes through the internal cavities of the compressor stages, comprising a casing fixed to a shaft located under the rotor disks of at least the last three compressor stages, and a system of seals and slots between these disks and the casing, openings in the disks and exhaust openings in the casing. The inlet is connected to the end region of the compressor, where the air temperature in the compressor is maximum. The seals, slots and openings are arranged so as to create a loop-like flow of hot compressed air in the casing along the blade webs to the exhaust openings through which air enters the casing in the general direction opposite to the direction of the air flow in the flow part in order to optimize the change in radial clearances adequately the operating mode of the flight cycle. The system of seals, slots and holes is designed so that its combined hydraulic resistance is less than the hydraulic resistance of the compressor seal disk.

Loop-like moving warmer / colder, depending on the operating mode, compressed air intensifies heat transfer, as a result of which the temperature is equalized throughout the disk web. Accordingly, there is an increase / decrease in the thermal expansion of the disk and other contacting elements, which reduces / increases the radial clearances, optimizing them throughout the flight cycle of the engine.

A comparative analysis by the authors of the patent of the results of calculating the change in the radial clearance at the last 7th stage of the compressor during the flight cycle, including starting, warming up, taking off, climbing, cruising, lowering, landing, reverse thrust and stop at the proposed compressor with a device regulation of radial clearances with the most optimal system of seals, slots and openings, and a compressor without such a device under the same flight conditions showed a significant useful decrease in radial Azora the first stages of the flight cycle including cruise mode in which the radial clearance reduction reaches 70%.

Particularly interesting and useful are the results of the operation of the device in the modes of reduction and flight in a circle, decrease in glide path, landing, reverse thrust, taxiing and engine shutdown. If there is no device at the reduction stage, the radial clearance is reduced to zero and there is a danger of cutting the blades into the housing. In the case of a compressor with a device, the radial clearance is even slightly increased compared to the clearance in cruising mode, and the possibility of cutting blades is excluded.

It can be accepted that in both cases, changes in the thermal deformation of the body occur according to approximately the same laws. Consequently, a 70 percent reduction in radial clearance when using the proposed device in cruising mode occurred only due to the thermal expansion of the rotor. Then, for a time of 3000 seconds (see FIG. 5 of Patent 2506436), the thermal deformation of the compressor rotor with the proposed device is 0.4 mm larger than the thermal deformation of the compressor rotor without the proposed device. The air temperature in the flow part of the compressor in the reduction and flight mode in a circle drops sharply and the compressor disks are cooled. Moreover, in the case of a compressor with the proposed device, the cooling rate is higher than that of a compressor without a device and a decrease in thermal deformation of the compressor rotor with the device is 0.2 mm larger than that of the compressor rotor without a device. This confirms the physical reliability of the results of the authors' calculations.

In the modes of descent along the glide path, landing, reverse thrust, taxiing and engine shutdown, the disks in the case of a compressor with the proposed device are more warmed up. Here, a frequent change of the compressor operating modes with medium and low air temperatures in the flow part occurs, but due to the inertia of the system, the radial clearance value remains somewhat average, which eliminates the risk of cutting the blades into the casing.

The undoubted advantages of the device according to the patent of the Russian Federation 2506436 is that it is able, with an optimal choice of a system of seals, slots and openings, corresponding to the operating conditions of the flight cycle, to ensure changes in radial clearances, in which flight safety, increased resource and increased compressor efficiency are achieved, and the choice is optimal systems are carried out by calculation.

This is a good result with information very useful for designers of turbomachines, however, it may turn out here that with all the factors affecting the size of the radial clearance in potentially hazardous modes fraught with cutting blades, two tenths of the gap may be insufficient.

Active thermal regulation of the radial clearance between the stator and rotor is typical of modern aviation gas turbine engines.

In the majority of completed designs of aviation gas turbine engines, to control the radial clearances, systems are applied that affect the temperature of the stator parts by controlling the amount of air used to blow them. Moreover, the greatest successes were achieved in the low-pressure fuel pumps, the stators of which are executed without local geometric features that cause warpage and loss of concentricity in transient conditions when blowing (see Shustrov Yu.M. Aviation air conditioning systems / Yu.M. Shustrov, M.M.Bulaevsky. - M.: Mechanical Engineering, 1978. - 159 p.).

There is a known system for blowing the high pressure fuel pump of the CF6 engine and the VD turbine of the JT9D - 59/70 engine (see Kuznetsov N.D., Danilchenko PD, Reznik V.E. science and higher school. SSAU. Textbook. Samara, 1991. - 109 p.), containing an air intake, a pipeline with a valve, a jet cooling system made in the form of pipelines covering the stator, with holes distributed along the length of the pipelines through which blowing the stator, and the electrical circuit that controls the open ti and closing the valve so that the stator airflow is included only in cruise mode at heights exceeding barostat limits.

A known system for regulating the radial clearance in the internal combustion engine of an E ³ engine of the Pratt-Whitney company by blowing air through stator parts with double walls (see the same book), including a pipeline, a manifold, a cooling air intake point in front of the first compressor stage, which is blocked by a valve, and the place of selection of heating air behind the compressor, blocked by a valve. The system includes cooling the stator with cooling air in cruise mode and stator heating in transient conditions.

Known examples of the use of a cooling system for stator housings in transient engine operation to reduce radial clearance and draft failure (see Aleksandrov A.A. P. Danilchenko, G. M. Gorelov [et al.] // Industrial Engineering. Academy of Sciences of the Ukrainian SSR. - Kiev, 1989, vol. II, No. 6. - P. 57-61.).

An example of the application of active thermal regulation is also the cooling system of the compressor TPDD PS - 90A internal compressor case, through which the air volume regulated by the open / closed position of the damper is supplied through the hole in the compressor external power case rigidly and tightly at both ends with the compressor internal case from the second circuit, which is regulated according to the operating mode of the flight cycle. Getting into the receiver, formed by the outer power casing and the shell located inside the casing and mounted on it, the air is distributed throughout the receiver ring to uniformly blow the inner casing through the holes evenly located in the casing. The blades of the guide vanes and the annular segments are fixed on the inner casing, forming radial gaps above the working blades. Air cooling the inner case exits through an opening located outside the receiver in the outer power case for reuse for other purposes.

Consider the operation of the radial clearance control device of this engine during the flight cycle.

In the transition mode, from starting the engine to entering the idle mode, the radial clearance changes are small and may not be taken into account. In the interval from low gas to take-off mode and when switching to cruising mode, the damper is in the “open” position. In cruise mode and with a decrease in engine speed, when switching from cruising mode to idle and before the engine stops, the shutter is in the “closed” position. As a result, in these modes, the cooling rate of the inner case decreases faster than the cooling speed of the rotor, but, this ensures the required value of the radial clearance, excluding the incision of the working blades in the inner housing of the compressor.

Known systems for regulating radial clearances by acting on the rotor.

Serial foreign engines with air-blowing of disks of the rotor of the high-pressure rotor are gas turbine engines CF6 - 80 C2 and PW - 4000 (see the same book). In the design of the HPC of the PW - 4000 engine, during take-off mode, air is supplied to the rotor cavity from the fifth stage of the HPC through the hollow guide apparatus of this stage, and in cruising mode, from the XI stage. Moreover, the design of the HPC of these engines have the following features:

the blade webs are made thin or completely absent (the first stages of the HPC CF 6 - 80 C2), which ensures their accelerated heating or cooling;

the selection of the masses of the rotor and stator brings their thermal inertia closer;

the regulation of radial clearances is carried out at all stages of the HPC;

power and path rings of the compressor are separated, which ensures concentricity;

in accordance with the regulatory program, the implementation of transients is preceded by a change in the number of the stage from which air is taken for blowing.

Known systems for regulating radial clearances by acting on the rotor and stator.

So, the CFM - 56 - 5 engine (see the same book) has two systems for regulating radial clearances - a control system for radial clearances of the HPC by affecting the temperature of the rotor of the HPC and a system for controlling radial clearances of the HPC by blowing air from the HPC and the stator fan of the HPP. In the control system for the radial clearance of the HPC, air is taken in behind the V stage of the HPC and through a pipeline with a valve is supplied to the rotor of the HPC in front of the guide apparatus of the first stage of the HPC and further to the rotor cavity of the HPP. In the TND radial clearance control system, air is supplied through pipelines with valves to blow the stator of the low pressure pump either because of the V stage of the high-pressure valve or because of its last stage. Air from the fan is also mixed with air drawn from the compressor through a pipe with a valve. The valves are controlled by an electronic digital engine management system.

Known high-pressure compressor turbofan engine with active thermal regulation of the radial clearance between the stator and the rotor (see RF patent 2033563. High-pressure compressor of a turbojet engine / IV Maksimov, NM Oshkanov, AI Tunkin. - http: //www.ru/patent/203/2033563/html), comprising an inner casing with guiding devices and intermediate rings fixed therein and enveloping its outer casing, rigidly connected to the inner casing by radial flanges, a sealed annular chamber made between the outer and inner them with buildings. The annular chamber is in communication with the flow part of the compressor with nozzle openings, and the nozzles are provided with rotary shutters and are in communication with the channel of the outer motor circuit.

The design of the HPC of this patent is fundamentally different from the design of the HPC of the PS - 90A engine only by the presence of nozzle openings in the compressor inner casing with a flow rate very small compared to the flow rate through the compressor flow path. In the transition mode from small gas to take-off mode, the radial clearance is reduced, as in the PS-90A engine, due to the supply of cooling air from the secondary circuit, and also to an insignificant extent, by reducing the pressure drop across the inner casing. During prolonged operation in cruise mode, the shutters are closed and cooling air does not come from the secondary circuit. Due to air leaks through the nozzle openings, the pressure drop across the inner casing gradually becomes zero and the temperature of the rotor and casing are equalized and the radial clearance becomes minimal. When switching to the low-gas mode, the inner case cools faster than the rotor, the radial clearance decreases, and the blades can cut into the case. To eliminate this danger, the dampers open again in this mode, the pressure outside the inner casing drops sharply to the pressure in the second circuit, the pressure drop across the inner circuit rises sharply, the casing elastically expands, which increases the radial clearance.

In our opinion, the advantages of this proposal can be realized if a number of conditions are discussed, which are discussed below.

In aircraft gas turbine engines, when the engine is operating in cruise mode, the radial clearance is usually greater than the mounting radial clearance due to greater than the rotor thermal deformation of the body and the pressure drop on the inner body, and without blowing it with cooling air from the second circuit can unacceptably reduce the compressor efficiency.

The implementation of the patent proposal will be effective only if a decrease in the radial clearance due to a decrease in the pressure drop across the housing to zero is greater than a decrease in the radial clearance due to air cooling from the secondary circuit and such that even with the exclusion of blowing it will reduce the radial clearance to a value that provides an acceptable value of efficiency compressor. Note that otherwise you will have to apply airflow. In this case, the effect of the presence of nozzle openings with a very low air flow rate will be insignificant or even harmful, since it leads to a decrease in compressor efficiency due to air leaks from them from the compressor flow path.

When switching to small gas, a rapid decrease in the radial clearance due to a rapid pressure release outside the inner casing will only occur during this release, then the radial clearance will begin to decrease due to the flow of cooling air from the secondary circuit.

Therefore, in order to obtain a significant positive effect on this transient process when using the patent proposal, one will either have to severely limit the time of this transient process or change the damper control program.

Note that the authors of the patent claim that in the transition mode from small gas to take-off mode the radial clearance decreases inaccurately, at least when the thin-walled inner compressor housing is made of metals and alloys used in aircraft engine manufacturing for its manufacture. It can only be true if the inner case is made of composite material with a small or zero coefficient of thermal linear expansion of the material, which is not specifically stipulated in the patent.

An example of the use of a passive thermal device for regulating radial gaps between the working blades of a turbine rotor and a stator is a two-stage high-pressure turbine (HPT) of the PS - 90A engine. Each stage of the turbine contains a massive annular insert with a honeycomb seal mounted on the casing of the theater, covering the working blades with a radial clearance. The blades of the honeycomb apparatus of the first stage of the theater and its disk and rotor blades are cooled by secondary air from the combustion chamber, blowing around the blade web and entering these blades, and then all these blades are additionally cooled by a film created on the surface of their feathers by this air entering through many throttle openings, evenly distributed over the entire height of the feather blade at its leading edge.

A reference to a similar well-known design of a control device for a single-stage high-pressure turbine made in the form of a ring with a radial clearance covering the working blades located inside the elastic element and attached to its vertices is given in the description of the patent of the Russian Federation 2503822 (see below).

However, both of these solutions are considered insufficient if the support to which the ring is attached is also subject to uneven thermal deformation at its periphery when such deformation leads to deformation of the turbine ring.

A number of patents are known, offering turbines of gas turbine engines with active devices for regulating radial clearances between the working blades of the turbine rotor and the stator. The importance of the problem being solved has given rise to a large number of its patented solutions over the past decade, and these devices with the regulation of radial clearances by selective cooling or heating of the stator or rotor elements are even allocated to a separate international patent class - F01D 11/24.

We will consider of them only, in our opinion, the most technologically advanced.

So known is a high-pressure turbine with an improved chamber for regulating the radial clearance of moving blades (see RF patent 2503822. A high-pressure turbine with an improved chamber for regulating the radial clearance of moving blades and a turbomachine using such a turbine / J. A. Dominic (FR), D. Peru (FR), W - L. Le Strat (FR), P. Pasyonssal (FR). Patent holder SNEKMA (FR). - Publication of the patent 10.01.2014. - http://www.freepatent.ru/patents/2503822) comprising an outer casing, a distributor, an impeller, a ring forming unit and arranged circumferentially five rotating blades, a device for controlling the radial clearance between the rotating blades and wingtips ring, an annular support and an annular damping element. The distributor is formed by a series of guide vanes, and the impeller is mounted on its output side. A device for adjusting the radial clearance includes a control chamber with annular chambers attached to the outer casing at least in two places at a distance from each other. The control chamber covers an annular support, which supports a ring covering the blades with a radial clearance, and is attached to the outer casing of the turbine. An annular damping element is fixed at one end to an annular support, and rests on the input side of the control chamber with elastic petals with specified elasticity and interference

Depending on the frequency of rotation of the turbine rotor, the control chamber supplies heating / cooling compressed air to the ring covering the blades, due to which the radial clearance is regulated.

The authors of the patent indicate that the control chamber during operation of the engine is subjected to strong vibrations, which are damped due to the work of the dry friction forces during elastic slipping with friction of the elastic lobes on the surface of the control chamber.

Note that this patent sets priority only on the damping device of the control chamber; the control chamber itself is proposed in the application FR 2865237 previously filed by the applicant.

The patent DE 4309199 A1 is known, in which a method for changing the temperature of individual rings covering the working blades of the stages of a multi-stage turbine using air or another refrigerant or using electric heating by induction is proposed.

A known method and device for adjusting the gap at the ends of the blades of a turbine rotor in a gas turbine engine (see RF patent 2425985, IPC F01D 11/24, MPK F01D 25/10. Method (options) and a system for regulating the gap at the ends of the rotor blades in a gas turbine engine, and also a gas turbine engine containing such a system / F. Winsay (FR). Patentee SNEKMA (FR). - Published patent. 08/10/2011. - http://www.freepatent.ru/patents/2425985).

The proposed method for regulating the gap between the ends of the blades of the turbine rotor and the ring of the outer turbine casing surrounding the working blades of an aircraft gas turbine engine operating in all flight cycle modes is that the gap increases due to heating of the external turbine casing with short-term engine operation modes: low gas mode and in a mode with rpm exceeding the speed of the cruising mode, or only in one of these modes.

The proposed device for adjusting the gap at the ends of the blades of the turbine rotor consists of two vertical annular walls integral with the outer turbine casing, these segments are attached end-to-end to segments, forming a ring covering the working blades of the stage of the turbine rotor, a heater, made as a resistive element, for example, in the form of a sheet covering the outer turbine casing over a ring of segments mounted on it, in an insulating material, inside of which one or several only conductors, or an induction circuit composed of one or more coils embedded in insulation, connected to the outer casing of the turbine in such a way that it can create an inductive coupling with its part, or with a fraction of this part, a current collector fixed to the external turbine circuit, breakers , wires, an autonomous digital control system, and wiring connecting all of these electrical devices to an aircraft alternating current electric generator.

This method and device can only be useful in throttling modes of the engine, in which the turbine casing or casings are cooled faster than its disk, where it is necessary to heat it to prevent the blades from cutting into the casing. In engine boost modes, the heater should be turned off, and this method and device does not have a beneficial effect on the radial clearance, which increases due to heating of the housing or bodies, faster than heating the turbine disk, and the turbine efficiency may unacceptably decrease in these modes.

A known turbine of a gas turbine engine according to the patent of the Russian Federation 2435039, IPC F01D 11/24, containing an external, internal and intermediate housing, a stage with a nozzle apparatus and an impeller with an annular insert covering the impeller, and mounted on the working and external housing, and means for regulating radial clearances of at least one turbine stage.

The disadvantage of this turbine is a sharp increase in radial clearance during engine boosting due to rapid heating of the housing.

Known double-circuit gas turbine engine (see RF patent 2511860, IPC F01D 11/24. Double-circuit gas turbine engine and method for regulating the radial clearance in the turbine of a double-circuit gas turbine engine / NB Bolotin. - Published on 04/10/2014. - http: // www .freepatent.ru / paterits / 2511860) comprising a compressor, a combustion chamber, a turbine 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, as well as a turbine stator, soda holding at least one turbine housing and an annular insert above the turbine impeller and a clearance control system, characterized in that the turbine stator is made of cooled air of the second circuit, while the supply system of this air is made in the form of an air intake installed in the second circuit, and a flow regulator with a drive, and contains an on-board computer and radial clearance measurement sensors, a flow regulator drive and radial clearance measurement sensors are connected by electrical connections to the on-board computer m

The advantage of this proposal can be considered the use of sensors for measuring radial clearance and its regulation by commands (program) of the on-board computer in accordance with the data on the actual value of the radial clearance measured by the sensors.

Note that in many turbojet engines the turbine stator is cooled by secondary air, and the rotor is cooled by air due to some intermediate stage of the compressor. New in this proposal is that the air of the second circuit enters the cavity between the outer and inner turbine casing in portions controlled by the on-board computer.

The disadvantage of this method is that the outer rings of the nozzle apparatus are subjected to the tensile effect of a large pressure drop equal to the pressure difference in front of the turbine stage and the air pressure in the second circuit, which leads to the need to increase the mass of these rings.

In addition, in our opinion, to create a highly efficient system for regulating the radial clearance in the turbine in all flight cycle modes only due to the cooling of the turbine rotor by air due to some intermediate stage of the compressor and the controlled cooling of its stator by cold air from the second engine circuit will fail the fact that in the throttle modes of the engine, for the above reasons, the air flow will be required to cool the turbine rotor, which unacceptably reduces the compressor efficiency, or in general it turns out physically impossible.

Known turbine of a double-circuit gas turbine engine with regulation of the radial clearance in the turbine (see RF patent 2519127, IPC F01D 11/24. Turbine of a gas turbine engine and method for regulating the radial clearance in the turbine / NB Bolotin. - Publish. 06/10/2014. - http : //www.freepatent.ru/patents/2519127), containing at least one cooled stage with a nozzle apparatus with cavities above and below it and a turbine rotor with a cooled impeller, as well as a turbine stator containing at least at least two turbine bodies with cavities between them and a control system a radial clearance comprising an annular insert above the impeller of the turbine, 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, the radial clearance control system includes an on-board computer, a radial clearance measurement sensor and microwave radiation sources, mounted above the ring insert. In this case, the sensor and sources are electrically connected to the on-board computer. Moreover, the sources of microwave radiation are configured to heat the annular insert.

This patent also proposes a method for regulating the radial clearance in this turbine, including cooling the rotor and heating the stator, characterized in that the radial clearance is measured and, depending on its magnitude, the inclusion of microwave radiation sources for heating the annular insert.

New in this method is only the use of microwave sources for rapid heating of the annular insert.

In addition, a new one in the proposed turbine design is the setting of a flow regulator in a system supplying air to cool the turbine stator, taken at the compressor outlet, and controlling the flow of this air according to the commands of the on-board computer, according to the signals of sensors measuring the gap.

Note that the casing of the high-pressure turbine, the blades of the nozzle apparatus, the disk, and the working blades of its first stage for known aircraft gas turbine engines are cooled by air taken from the last stage of the high-pressure compressor (HPC) (secondary air from the combustion chamber). The temperature Tr in front of a high-pressure turbine in modern gas turbine engines reaches 1800 ° K. With this method of cooling the disk and rotor blades of the first stage of such a turbine, in our opinion, with the use of this air, it is permissible to reduce the turbine's efficiency, none of the known aircraft engines had optimal control of radial clearances along the working blades of such a turbine has been achieved. Those. in order to ensure optimal regulation, in our opinion, too high air consumption with high potential will be required, possibly even in cruising mode, and, therefore, will lead to an unacceptable decrease in the efficiency of the turbine and the entire engine, including this mode of operation.

We also note that at present no sensors have been created that measure the gap, capable of operating at Tr = 1800 ° K, although such work is underway.

In our opinion, the regulation of the radial clearance of a turbine according to RF patent 2511860, IPC F01D 11/24 in the engine boosting modes can be more flexible and more accurate than that of this turbine, and, therefore, more efficient. In the modes of boosting the engine of a turbine according to the patent of the Russian Federation 2519127, IPC F01D 11/24, microwave sources are turned off. The annular insert according to the patent of the Russian Federation 2519127 is cooled by much hotter air than the annular insert according to the patent of the Russian Federation 2511860. In addition, the conditions for its cooling here are still worsened by the location of the microwave radiation sources directly above the annular insert, which negatively affects the flow of cooling air around it. Although microwave sources heat up quickly, they cool as slowly as other heat sources. This especially reduces the effectiveness of the method for regulating the radial clearance of a turbine according to the patent of the Russian Federation 2511860, IPC F01D 11/24 in engine boosting modes with repeated engine throttle response when the microwave source is hot.

The thermal inertia of gas turbine engine turbines is high, and therefore, in many cases, the thermal inertia of resistive and induction heating sources may be quite suitable for use in systems for regulating the radial clearance of turbine engines.

Turbine designs with these heat sources are simpler than turbines with microwave sources, as these sources can be powered by an airplane generator, and special equipment is required to power microwave sources. Therefore, it is advisable to use microwave sources only in those cases where the use of resistive and induction sources is not effective.

In addition, a common drawback of RF patents: 2425985, 2435039, 2511860, 2519127 is the lack of effective damping in these proposals of possible oscillations of the proposed radial clearance control system.

Almost all leading aviation companies developing new designs of aircraft engines are engaged in the development of systems for the thermal active regulation of the radial clearances of aircraft turbine engines. But at present, none of the known patented designs of turbines with a system of thermal active regulation of radial clearances with special heaters provides an indisputable advantage in all flight cycle modes compared to other proposed such systems. Apparently, this and the fact that such systems in known cases are designed to regulate the radial clearances of only one stage of the turbine, is still one of the reasons for the absence of such systems in the design of aircraft gas turbine engines in operation.

We emphasize once again that attempts to introduce an active system for regulating the radial clearance of a VD turbine have so far not been particularly successful (see Danilchenko VP, Lukachev SV, Kopylov Yu.L. [et al.]. Design of aircraft gas turbine engines. - Samara, Publishing House of the SSC RAS, 2008 .-- 619 s).

Therefore, the task is to propose the design of a single and two-stage VD turbine of a dual-circuit aircraft engine with a system of thermal active regulation of radial clearances, which is structurally simple, maintainable, with the possible quick replacement of worn-out components of the radial clearance control system, with a good mass characteristic, effective in all flight cycle modes, moreover, one whose effectiveness in each mode would be at least not lower than the best result, is hypothetically achieved on this mode, any of the known proposals, and it is combined with the known successfully applied to in-service and in the development of gas turbine engine turbochargers with controls radial clearances. Moreover, the position of the structural element of the flow regulator, which regulates the flow of cooling air in the cruise mode, would be fixed with a deenergized drive of the flow regulator, and possible oscillations of the radial clearance control system would be effectively damped, and the cooling air of the second circuit after cooling the ring insert would be reused useful.

The turbine of a double-circuit gas turbine engine with regulation of the radial clearance in the turbine according to the patent of the Russian Federation 2519127, MPK F01D 11/24 is by its technical essence closest to the proposed one and adopted as a prototype.

The problem is solved in that a one-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, comprising one cooled stage with a nozzle apparatus with cavities above and below it and a turbine rotor with a cooled impeller, as well as a turbine stator, comprising at least two turbine bodies with cavities between them, into which compressed air enters due to the last stage of the high pressure compressor, and the system is regulated a radial clearance containing an annular insert above the turbine impeller, covering the rotor blades of the turbine rotor with an annular radial clearance, and elastically and hermetically fastened to the parts forming the turbine’s inner casing, a heater enclosing the annular insert with the possibility of heating, an air intake, a flow regulator cooling air with a drive, an on-board computer and sensors, and a heater, a flow regulator drive and sensors are electrically connected to an on-board computer, characterized in that about the annular insert is hollow and placed in a cavity separated from the second circuit of the engine by a housing and a vertical wall, jointly mounted on the housing of the combustion chamber, and the annular insert with annular protrusions made on its sides, is tightened with the possibility of thermal expansion in the reciprocal ring grooves vertical wall and flange made on the inner part of the outer housing of the theater, the nozzle apparatus with ring protrusions is also tightly fixed in the reciprocal ring grooves the vertical wall and the housing of the combustion chamber, in which holes are equally distributed around the circumference in the vertical wall, through which secondary air enters the cavity above the nozzle apparatus and the annular insert, and the disk and rotor blades of the turbine engine wheel are also cooled by secondary air swirling device before entering the disk blade, and the outer ring of the nozzle apparatus, the annular insert and the inner part with the flange of the outer housing of the theater form an internal tight a TVD case, a microwave ring heater, or resistive or induction, consists of two separate half rings, each made in the form of a metal case, inside which a heating element is fixed, and each half of the heater is mounted on the ring insert with the possibility of radial thermal expansion together with the ring insert and tangential thermal expansion relative to the annular insert using a bayonet connection with it and dowels located with interference in the mating grooves of the annular insert and each two half rings in its middle cross section, two nozzles are screwed into the annular insert diametrically opposite with an interference fit through a conical pipe thread - a nozzle - an air intake for supplying cooling air from the second circuit to the inner cavity of the annular insert and a pipe for removing this air to the second circuit or for cooling the nozzle , or for other purposes, and each nozzle passes through an opening in the outer theater of the theater and in the case, hermetically mounted on the theater, the connection of each nozzle to this it is sealed with a two-piece piston ring pair, and in each pair of piston rings the cuts of these rings are diametrically opposed, a support is screwed on each pipe, or this support is integral with each of the pipes, and central to this pipe symmetrically secures the pipe with lock washers and screws supports of two springs, made in the form of a multilayer package, compressed by a distributed load, recruited from steel, red-hot or cured, polished tapes made of stainless steel, hardened by a wear-resistant coating, and the springs themselves with their end supports are fixed in the second circuit on the outer theater housing, so that the longitudinal axis of the package is perpendicular to the axis of the engine, and due to the elastic deformation of the package, the required additional elastic additional fixing force of the ring insert is created, and the package in the central support and two end supports of the spring is fixed motionless with the help of two rivets located on the sides of the package without a gap in the semicircular recesses made in tapes, or the package at the end the support is fixed with the possibility of displacement of its ends along its longitudinal axis within the straight section, made in this case in the central part of the semicircular recesses, the free end of the pipe - air intake from the side of the air flow in the second circuit is obliquely cut, and the pipe of the air outlet is connected to the pipeline, at the output of which either a normally open or a normally closed electro-pneumatic valve is installed, and radial clearance control is performed according to the on-board computer commands according to the method of claim 14 of the formula a spool, or a spool valve with an electromagnetic drive, designed so that the position of the spool regulating the air flow in cruise mode at heights exceeding the height of the barostat limit H is fixed with a de-energized electromagnetic drive, and the electromagnetic drive is made with rotation protection and with a spring return, turning on and off the heater, and regulating the intensity of its heating, opening and closing the supply of cooling air from the second circuit and the intensity This approach occurs according to the on-board computer commands generated by the program according to the signals of the sensors — engine speed sensor and barostat according to the method of claim 13, or sensors measuring the size of the radial clearance by the blades according to the method of claim 15 of the claims / rectangular engine hatches, hermetically closed by covers, are made over the housing of the engine’s second circuit above the spring arrangement, and a nozzle is made in the outer theater assembly connected by a pipeline to the intermediate stage KV D, supplying cooling air for cooling the front nozzle apparatus of the medium pressure turbine.

In a one-stage theater of operations, a variant of the method according to claim 15 of the claims is used, in which the sensors measuring the radial clearance are fixed in the annular insert of the proposed one-stage theater.

A new and fundamentally important, in our opinion, is the proposal for cooling the annular insert with air supplied to the internal cavity of the annular insert, as well as the use of this proposal in combination with the supply of secondary air from the combustion chamber (CC) in the cavity above the nozzle apparatus of the theater of manufacture and above the annular insert and for cooling the disk and rotor blades of the theater rotor. Moreover, the first sentence provides an extension of the range of control of radial gaps, allows more flexible and precise control of the values of these gaps. The use of the second sentence almost completely relieves the internal theater housing from pressure differences, since the air pressure in front of the turbine impeller is equal to the pressure in the indicated cavities, equal to the secondary air pressure (approximately equal to the air pressure behind the last stage of the HPC). This greatly simplifies the constructive solution to the problem of ensuring the strength and tightness of the internal housing of a theater, and reduces the mass of its parts, especially the mass of the outer ring of the nozzle apparatus.

Note that the use of an annular insert, controlled selection of cooling air from the second circuit and its supply into the internal cavity of the annular insert with the simultaneous supply of secondary air from the combustion chamber to the cavity above the annular insert, a heater with an adjustable speed and heating temperature, a computer, and a speed sensor and barostat and their use according to the method of claim 13 of the claims, or sensors measuring radial clearance, and their use according to the method of claim 15 of the claims in one system radial clearances, in our opinion, is new, although the use of these devices individually, or their combination, different from the proposed one, is known from published literature and patent literature (see above).

We also note that the fulfillment of the condition for maintaining the position of the spool on cruising mode with a de-energized electromagnetic drive, although it reduces energy costs, but at the same time increases the stroke of the spool and drive, which in turn increases the mass and dimensions of the spool distributor and drive.

The application of the proposed single-stage turbine engine allows organizing a more accurate and flexible than the modern aircraft engine (see above) control system for the radial clearances of the turbine VD, which, in turn, will reduce the mounting radial clearance in the theater, and reduce the air consumption with high potential taken due to the last stage of the compressor (due to its use only in volumes sufficient for unloading from the pressure drop of the internal housing of the theater), to ensure the operation of the engine in all modes of the flight cycle with less radial clearances, and, therefore, increase the efficiency of the theater of operations and the engine as a whole, reduce the specific fuel consumption C _p , avoid the danger of cutting the ends of the working blades of the theater rotor when starting the engine and in the throttle modes of the engine when it stops and the thrust fails to take off.

Moreover, in the most typical cases, the heater is used to preheat the annular insert when starting the engine to avoid cutting the rotor blades of the theater rotor into the stator due to the elastic pulling of the blades and the castle part of the disk under the action of centrifugal forces, and during throttle operation of the engine and its stop to increase radial clearances and exclusion of insertion of rotor blades, and cooling with air supplied from the second circuit is used in boosted engine modes to reduce radial clearance and elimination failure thrust at takeoff and with a smaller flow rate of the cooling air at cruise mode.

The implementation of the heater in the form of two half rings, securing them to the annular insert using a bayonet connection and fixing them from displacement as a solid in the circumferential direction using a key located in the middle section of each half ring provides free thermal expansion of the rings relative to the annular insert in the circumferential direction and their displacement, as a solid, without separation from the annular insert in radial directions with its temperature expansion in these directions, which eliminates the appearance of tamper tural tensile stress in the heater during the heating of the heater and the annular insert.

The radial cross-sectional displacements of the annular insert in the radial directions during its thermal expansion can be large, and the increase in the diameter of the annular insert can exceed 4-5 mm (see Kuznetsov N.D., Danilchenko P.D., Reznik V.E. Control of radial gaps in turbochargers of aviation gas turbine engines. - State Committee of the RSFSR for science and higher education. SSAU. Textbook. Samara, 1991. - 109 p.) and is provided by the size of the gap between the outer surface of the protrusions of the annular insert and the reciprocal surface of the grooves of parts in which the ring Vai insert.

The use of a bayonet joint provides a symmetrical design of the annular insert in the circumferential direction and, therefore, the isotropy of its thermal expansion in radial directions. The use of piston rings provides the possibility of displacement of the nozzles under the action of thermal expansion of them and the ring insert and the static and dynamic loads acting on these parts without depressurization of the joint between the nozzle and the body part. The use of pairs of piston rings and the placement of the joints of the piston rings of the pair in diametrically opposite radial sections overlaps all possible leakage channels and seals this joint.

An oblique cut of the end section of the pipe - air intake increases the area of its inlet.

The elastic suspension of the nozzles on the springs, made in the form of a multilayer package, was chosen because they are calculated (see Eskin I.D. Study of the generalized elastic-friction characteristics of dampers and shock absorbers of aircraft engines: dis ... Candidate of Technical Science / I.D. Eskin. - Kuibyshev: KuAI, 1973. - 150 p.) And calculation it is possible to determine the parameters of the spring, which, with its great flexibility and large deflection equal to the maximum possible displacement of the nozzle in radial directions, are quite constructive and at the same time the spring strength is ensured ora, and the pressure that these springs exert on the annular insert during this deflection will be quite acceptable. Fixed sealing of a multilayer package in the spring supports should be used when ensuring the spring strength does not lead to an unconstructively long spring, and when the forces compressing along the longitudinal axis of the spring due to the thermal elongation of the spring do not lead to loss of stability of the spring when exposed on it the maximum possible bending load. Note that the action of such a compressive load can be useful, as in this case, when a spring compresses along the longitudinal axis and a cyclic bending load is applied, the dispersion coefficient of the spring will increase over the entire range of amplitudes of its deformations (see Eskin I.D. , dis ... candidate of technical sciences). The use of springs with termination of the multilayer package in the end supports with the possibility of slipping its ends in the direction of the longitudinal axis of the spring allows to reduce the length of the spring.

In addition, the proposed springs have very high damping characteristics, which are discussed in detail below, and effectively dampen vibrations of parts of the radial clearance control system of the proposed theater.

The symmetrical arrangement of the springs relative to each nozzle reduces the possibility of misalignment of the nozzles when they are subjected to dynamic loading and the annular insert, and, therefore, increases the reliability of sealing the joint sealed with piston rings and reduces wear on the piston rings.

The presence of hatches in the outer casing of the second circuit provides good access to the springs - parts that are most susceptible to wear, and their replacement in airfield conditions.

An alternative installation at the outlet of the secondary cooling air exhaust pipe of a normally open, normally closed electro-pneumatic valve, or spool valve has been proposed because the conditions for efficient control of radial clearances for various aircraft gas turbine engines are very diverse and some functions have enough functions to achieve such control “Turning on” and “turning off” the heater and “opening” and “closing” the supply of cooling air from the secondary circuit, while for other engines In order to achieve such a result, in addition to these functions, you will also need the functions “change the heating intensity and“ change the intensity of the supply of cooling air of the second circuit ”.

Proposal of two alternative methods of controlling radial clearances based on the use of different measuring instruments - one method uses an engine speed sensor and a barostat to enable intensive airflow during flight at altitudes exceeding the height of the barostat limit, and the other uses radial clearance sensors fixed in an annular insert directly above the working blades, despite the fact that it is known in advance that the second method is more accurate and more effective according to the achieved results of the first, about because there are currently no sensors for measuring radial clearance, capable of operating at Tr = 1800 ° K. The operating range of such sensors is limited by the temperature Tr = 1500 ° K. But at present, many leading aviation companies are developing radial clearance sensors capable of operating at Tr = 1800 ° K. Therefore, it is recommended to use the first of these methods for aviation gas turbine engines with a single-stage turbine engine with a temperature of Tr = 1800 ° K, and the second of these methods for aviation gas-turbine engines with a single-stage turbine engine with a temperature of Tr≤1500 ° K.

In order to reduce the inertia of the control system of the proposed single-stage turbine engine and increase the efficiency of this system, a single-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, characterized in that the cooling air supply system to the blade web and the working blades of the turbine rotor contains a pipeline with a shutter connecting the cavity behind the last stage of the HPC with the cavity in front of the winding device, and the control system itself radial clearances of the turbine is controlled according to process variant n. 13 drive web with active cooling, or by the method of claim. 15 claims.

This design of the theater should be used in the case when the increment of the radial clearance in take-off mode is large with the radial clearance control system turned off (see below).

For use in a DDRT with a temperature of Tr = 1800 ° K with a radial clearance control system using a speed sensor and a barostat and methods for controlling these clearances, a single-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed characterized in that a layer of abradable material is applied to the inner working surface of the annular insert.

In order to improve the elastic-frictional characteristics (UVC) of the turbine, a one-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, characterized in that each spring is assembled from straight tapes and is compressed by a uniformly distributed load - air pressure in the second circuit of the engine.

The spring parameters of this turbine can be determined by calculation (see Eskin I.D. Study of the generalized elastic-friction characteristics of dampers and shock absorbers of aircraft engines: diss ... candidate of technical sciences, appendix / I.D. Eskin. - Kuibyshev: KuAI, 1973. - 315 from). As the number of ribbons n in the spring increases, its dispersion coefficient ψ grows and, for n = 10, it reaches a very large value ψ _max > 6, and starting from n = 15 with a further increase in n, it grows asymptotically, and for n = 15 its dispersion coefficient ψ _max = 7 . Therefore, it is recommended to use springs with n = 10 ÷ 15 tapes with a thickness h = 0.4 ÷ 0.5 mm.

The pressure in the second circuit of the engine is not large and may not be enough to obtain the amount of energy dissipated by the springs necessary for the effective damping of the oscillations of the parts of the radial clearance control system.

Therefore, in this case, in order to increase the energy dissipated by the springs during the oscillation cycle, we propose a single-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine, characterized in that each spring is assembled in the following arrangement: its package is assembled from the same ribbons thickness, in the center of it are two or more smooth ribbons, on both sides are packages assembled “corrugation into corrugation” of two or more corrugated ribbons in this way m, that the vertices of the corrugations of one packet rest on a packet of smooth ribbons in the same sections as the vertices of the second packet, and the step of the corrugations of the corrugated ribbons is chosen such that only one vertex is located in the span of the packet, resting on the packet of smooth ribbons in the middle of the span, and under each spring support there is only one vertex of the corrugation, and packages made of one, two or more smooth ribbons are installed on the packages of corrugated ribbons, and in the assembled spring corrugations of the corrugated ribbons are completely straightened.

Although these springs have a lesser dispersion coefficient than springs compressed by a uniformly distributed load, it is still quite high ψ _max = 5, and the energy dissipated during the oscillation cycle with the same design parameters and the same deflection is larger for this spring, since In addition to the pressure in the second circuit, the tapes are additionally compressed by the load obtained by elastic straightening of the corrugated packets. The parameters of these springs are also calculated (see Eskin I.D., disc ... Cand. Tech. Sciences).

, где k=5÷10, h - толщина внутренних лент пакета, и шаг гофрированных лент пакета выбран таким, что в каждом пролете располагаются одна, две и более вершин гофров, и в собранной рессоре гофры гофрированных лент полностью выпрямлены.A one-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is also proposed, characterized in that one smooth tape with a thickness of one is installed on packages of smooth tapes or directly on packages of corrugated tapes

, where k = 5 ÷ 10, h is the thickness of the inner ribbons of the bag, and the step of the corrugated ribbons of the bag is chosen so that in each span there are one, two or more vertices of the corrugations, and in the assembled spring the corrugations of the corrugated ribbons are completely straightened.

With an increase in the value of k, the scattering coefficient ψ of the spring decreases, but at k = 5 ÷ 10 it drops slightly compared to a spring drawn from tapes of the same thickness (k = 2) (see Eskin I.D., diss. ) Installation outside the package of tapes with such a greater thickness allows guaranteed to fully straighten the ribbons of the corrugated bags in the assembled spring, which improves its elastic friction properties.

From the above analysis of known radial clearance control systems, it can be concluded that in operating aircraft engines (as opposed to published patents), only systems are introduced that control radial clearances of not one, but all of the device steps at once - for example, low-pressure, high-pressure and high-pressure pumps.

Modern aircraft gas turbine engines are performed with one or two-stage high-pressure turbine.

Therefore, the task is to create a two-stage high-pressure turbine of a double-circuit gas turbine engine with an effective control system for the radial clearance system of both stages of the VD turbine, which is quite compatible with other radial clearance control systems of the turbocharger of the engine, which is quite simple in design with good mass characteristics.

The turbine of a double-circuit gas turbine engine with regulation of the radial clearance in the turbine according to the patent of the Russian Federation 2519127, IPC F01D 11/24 is in technical essence closest to the proposed one and is also taken as a prototype.

The problem is solved in that a two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, comprising two cooled stages with nozzle devices with cavities above and below them and a turbine rotor with a cooled impeller of the first stage, and a turbine stator containing at least two turbine bodies with cavities between them, and into the cavity of the first stage of which compressed air enters due to the last stage of the compressor a high pressure nozzle, and a system for regulating the radial clearance of each turbine stage, containing an annular insert, above the turbine impeller, covering the rotor blades of the turbine rotor with an annular radial clearance, and elastically and hermetically fastened to the parts forming the turbine’s inner casing, a heater covering the annular insert with the possibility of heating, an air intake, a cooling air flow regulator with a drive, and sensors, and heaters, flow regulator drives are connected by electrical connections to computer mouth, characterized in that the annular insert of the first stage of the turbine is made hollow and placed in a cavity separated from the second circuit of the engine by a housing and a vertical wall, jointly mounted on the housing of the combustion chamber, and the annular insert with annular protrusions made on its sides, with interference fixed with the possibility of thermal expansion in the reciprocal annular grooves of the vertical wall and flange made on the inner part of the outer casing of the first stage of the theater, the nozzle apparatus lug projections are also tightly fixed in the reciprocal annular grooves of the vertical wall and the combustion chamber body, in which the holes are equally distributed around the circumference in the vertical wall, through which secondary air flows into the cavity above the nozzle apparatus and the annular insert, and the disk and working the blades of the wheels of the first stage of the theater are also cooled by secondary air, twisted by a twisting device before entering the disk blade, a microwave ring heater, or resistive, or induction consists of two separate half rings, each made in the form of a metal casing, inside which a heating element is fixed, and each half of the heater is mounted on an annular insert with the possibility of radial thermal expansion together with the annular insert and tangential thermal expansion relative to the annular insert using a bayonet connection with it and dowels located with interference in the mating grooves of the annular insert and each half ring in its average cross section, in the ring the insert is diametrically opposite with an interference fit on a conical pipe thread, two nozzles are screwed - a nozzle - an air intake for supplying cooling air from the second circuit to the inner cavity of the annular insert and a pipe for removing this air to the second circuit or for cooling the nozzle, or for other purposes, and each nozzle passes through an opening in the outer casing of the first stage of the theater and in the casing part, hermetically mounted on this casing of the theater, the connection of each pipe with this casing part is sealed with two pairs piston rings, moreover, in each pair of piston rings, the cuts of these rings are diametrically opposite, a support is screwed on each nozzle, or this support is integral with each of the nozzles, and the central bearings of the two springs of the first stage are attached symmetrically to the nozzle with lock washers and screws TVD, made in the form of a multilayer package, compressed by a distributed load, recruited from steel, hardened or cured, polished tapes made of stainless steel, coated with wear-resistant coating, and these springs themselves with their end supports are fixed in the second circuit on the outer casing of the combustion chamber and on the outer casing of the first stage of the theater, so that the longitudinal axis of the package is parallel to the axis of the engine, and due to the elastic deformation of the package created the required elastic force additional fixation of the annular insert, and one of the end supports is made common for both springs of the first stage, and the other is common for four springs - two springs of the first stage and two structurally similar p second step essor, in which the common second end support is fixed on the outer casing of the first stage of the theater of operations, and the second end support, common for two springs of the second stage, is fixed on the outer casing of the second stage of the theater of operations, and the package of each spring in the central support and two end supports is fixed motionlessly using two rivets located on the sides of the bag without a gap in semicircular recesses made in ribbons, or the bag in the end supports is fixed with the possibility of displacement of its ends along its longitudinal axis within the direct a run made in this case in the central part of semicircular recesses, the free end of the air inlet pipe on the side of the air flow in the second circuit is obliquely cut, and the air exhaust pipe is connected to the pipeline, at the outlet of which either a normally open or normally closed electro-pneumatic valve is installed, and in the outer casing of the first stage of the theater of operations, a nozzle is made, connected by a pipe to the intermediate stage of the HPC, supplying cooling air to cool the nozzle apparatus of the second stage of the theater of operations and through the equalizer The holes located around the circumference in a vertical wall, fixed together with the outer casing of the second stage of the theater of operations, into the cavity above the annular insert of the second stage of the theater of operations, and the intermediate stage of the cylinder is considered to provide a slight pressure drop on the outer ring of the nozzle device of the second stage of the theater of operations the apparatus is mounted in the stator of the theater of operations with the help of annular protrusions that fit with interference into the reciprocal ring grooves made in the flange of the inner part of the outer casing of the first stage of the theater of operations and vert on the inner wall, the annular insert of the second stage is attached to the same vertical wall and the vertical wall made integrally with the outer casing of the second stage of the theater, and the outer ring of the nozzle apparatus of the first stage, its ring insert, the outer ring of the nozzle apparatus of the second stage, its ring insert, the elements of the external TVD housings to which these parts are attached form an internal sealed TVD housing, an annular insert, an air inlet, an air outlet, a heater, body parts, and a central The spring supports and connections of these parts of the second stage of the theater are structurally similar to these parts and their connections of the first stage, but can differ from them only in parameters, and the branch pipe of the air exhaust of the second stage is connected to the pipeline, at the outlet of which is installed either normally open or normally closed the electro-pneumatic valve, or the cooling air exhaust pipes of the first and second stages are each connected to their entrance to the spool valve with an electromagnetic drive, designed so that each stage either has its own cooling air outlet from the spool distributor, or both steps have a common outlet from it, and the position of the spool regulating the air flow in cruise mode, at heights exceeding the height of the barostat H limit, was fixed with a de-energized electromagnetic drive, and the electromagnetic drive is made with protection against cranking and with a return spring, turning the heaters on and off and regulating the intensity of their heating, opening and closing the cooling supply the air from the second circuit and the intensity of this supply of both stages occur according to the on-board computer commands generated by the program according to the signals of the sensors measuring the size of the radial clearance by the blades according to the method of claim 15, fixed in the ring insert of the second stage, and in the outer casing of the second circuit engine over the arrangement of springs made rectangular hatches, hermetically sealed with covers.

With a fairly high degree of certainty, it can be accepted that for both stages of the turbine, for the effective control of radial clearances at each moment of engine operation, the same active process is required - cooling or heating, although the intensity of these processes at the first and second stages of the theater of operations can be different.

This assumption is the basis of the proposed control system for the radial gaps of the first and second stages of a two-stage turbofan engine using the signals of sensors measuring the radial clearance of the second stage of the turbine. The working temperatures of the second stage make it possible to use current eddy and capacitive sensors already available for measuring the gap in mm (see Danilchenko VP, Lukachev SV, Kopylov Yu.L. [et al.]. Design of aviation gas turbine engines. - Samara, Publishing House - SSC RAS, 2008. - 619 p.).

In order to reduce the inertia of the control system of the proposed two-stage turbine engine and increase the efficiency of this system, a two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, characterized in that the cooling air supply system to the blade web and rotor blades of the first stage of the turbine engine contains a pipeline with a shutter connecting the cavity behind the last stage of the HPC with the cavity in front of the tightening device of the first penalties, and the control system of the radial clearance of the turbine is controlled according to a variant of the method of claim 15 of the claims with active cooling of the disk web.

This design of a two-stage theater of operations should also be used when the increment of the radial clearance of the first stage of the theater of operations is large during take-off mode when the cooling of the disk of the impeller of the first stage is turned off (see below).

, где k=5÷10, и шаг гофрированных лент пакета выбран таким, что в каждом пролете располагаются одна, две и более вершин гофров, и в собранной рессоре гофры гофрированных лент полностью выпрямлены.A two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, characterized in that each spring is assembled from direct tapes and compressed by a uniformly distributed load - air pressure in the second circuit of the engine, or each spring is assembled in the following layout: its package assembled from tapes of the same thickness, in the center of it are two, three or more smooth tapes, on them on both sides are packages assembled “corrugations into corrugations” of two and more corrugated tapes in such a way that the corrugated vertices of one packet rest on a packet of smooth ribbons in the same sections as the vertices of the second packet, and the corrugated corrugation pitch of the corrugated ribbons is selected so that only one vertex is supported in the span of the packet, resting on the packet of smooth ribbons in the middle of the span, and under each spring support there is only one vertex of the corrugation, and packages of one, two or more smooth ribbons are installed on the packages of corrugated ribbons, and in the assembled spring of the corrugation of the corrugated ribbons are completely straightened, or on packages of smooth tapes or directly on packages of corrugated tapes, one smooth tape with a thickness of

, where k = 5 ÷ 10, and the step of the corrugated tapes of the packet is chosen so that in each span there are one, two or more vertices of the corrugations, and in the collected spring the corrugations of the corrugated tapes are completely straightened.

A two-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, characterized in that a layer of abradable material is deposited on the inner working surface of the annular insert of the first stage, or a layer of abrasive material is deposited on the inner working surface of the ring inserts of both stages of the turbine engine. In this case, the sensors measuring the radial clearance are shifted in the sockets of the annular insert of the second stage in radial directions by the thickness of the abradable layer.

With the proposed method for controlling radial gaps, the accuracy of controlling these gaps in the first stage is obviously lower than in the second stage and in some cases it may not be possible to guarantee the complete exclusion of the impellers of the wheels of the first stage of the theater of operations in the annular insert. Therefore, in these cases, it is advisable to apply a layer of abradable material to the inner working surface of the annular insert of the first stage. A layer of abradable material deposited on the inner working surface of the annular inserts of both stages of the theater will be useful in emergency situations, for example, when the working blade is broken, when it becomes impossible to prevent the cutting of the working blades.

A two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is proposed, characterized in that the hydraulic diameters of the path supplying the cooling air of the second circuit inside the annular insert of the first stage, and its discharge, are larger than the hydraulic diameter of a similar path of its second stage, or vice versa.

The increment of the radial clearance due to the thermal expansion of the stator and the turbine rotor may not be the same for the first and second stages of the theater. If this increment is larger at the first stage, then it is advisable to make the hydraulic diameter of its path supplying the cooling air of the second circuit inside the annular insert larger than that of a similar path of the second stage, if less, then vice versa.

For the same reason, in order to increase the efficiency of the proposed radial clearance control system, a two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine is also proposed, characterized in that the electrical parameters of the half rings with resistive heater conductors or induction heater coils are different for the first and the second stage of the turbine and the heater of each stage of the theater operates according to its program.

To ensure fuel economy in cruising flight conditions, many gas turbine engines require a strong effect on the temperature of the mating parts. The monograph (see Danilchenko V.P., Lukachev S.V., Kopylov Yu.L. [et al.]. Design of aircraft gas turbine engines. - Samara, Publishing House - SSC RAS, 2008. - 619 pp.) a method that solves this problem, which consists in combining two systems: blowing the stator and partially turning off the air used to cool the turbine rotor, which come into operation by one control signal. The stator cooling is turned on, and the rotor is turned off at high pressure cascade revolutions _nd = 9200 rpm and turned off at _ndd = 8800 rpm. During transient engine operation, the operation of the twin theater cooling system is not controlled by the radial clearance control system. The radial clearance control system is switched on at heights exceeding the barostat limit height (N≥9 km) and is controlled by the electronic engine control system.

Turning off the air cooling the turbine rotor improves the efficiency of the engine cycle and is advisable in connection with a significant decrease in the temperature of the gas and turbine blades in cruise mode compared to take-off.

The combined systems provide a counter radial temperature extension of the stator and rotor, which allows to reduce the radial clearance while reducing the cooling air flow to a level that provides the design life of the theater.

The radial clearance control tools used in the proposed theater of operations allow us to propose radial clearance control methods that are more effective than the known ones described above.

Therefore, we propose a method of thermally active regulation of radial clearances, in our opinion, quite simple and could be useful in many modern gas turbine engines, used in the proposed designs of a single-stage high-pressure turbine with Tr = 1800 ° K.

The method for managing radial clearances described above, comprising combining two theater cooling systems, is selected as the prototype.

The proposed method of thermal active control of the radial clearances of a single-stage high-pressure turbine contains operations for controlling radial clearances performed by the engine electronic system according to the signals of the rotor speed sensor of the VD turbocompressor and the barostat, and the disk and rotor blades of the turbine engine rotor are cooled by air because of the last stage of the HPC - without commands from these sensors, or actively - the cooling of these parts is turned off during take-off mode and transitional modes from take-off mode cruise mode, characterized in that all radial clearances theater management operations are performed according to the flight cycle on commands board according to the program computer, developed on experimental basis for the entire flight cycle, providing good efficiencies, C _p, and the characteristics of its traction and safety operation of the engine, excluding insertion of the working blades and failure of the takeoff thrust, which is performed in the following sequence: before starting the engine, the heater is turned on, which is turned off after a certain For a certain period of time, the engine starts and in forced modes the cooling of the annular insert is turned on by the air of the second circuit supplied to its internal cavity with the prescribed law of change in revolutions and time, with an intensity that reaches its maximum value during take-off mode, with passive cooling of the disk and rotor blades A theater of operations is performed at all operating modes of the engine and is not regulated by the radial clearance control system, and at a regime transitioning from take-off mode to cruising mode the annular insert is cooled by the air of the second circuit with an intensity that gradually decreases according to the law specified by the computer program to the level required in cruising mode at heights exceeding the height of the barostat limit N, and with the indicated active cooling of the disk and rotor blades of the theater rotor, the specified cooling law of the annular insert air of the second circuit, also decreasing to the same level in cruise mode, but with large gradients of intensity reduction in engine speed, and in throttle The desired flight cycle modes turn off the cooling of the annular insert by the air of the second circuit and turn on its heating with a heater with an intensity that gradually decreases according to a given law with a decrease in engine speed, and turns off the heater and stops the engine, and the excited oscillations of the annular insert and parts attached to it in all operating modes the engine is effectively damped by special dampers.

Here, the cooling intensity is understood as the value of the second flow rate of cooling air, which can be determined by calculation.

The time interval for heating the annular insert before starting the engine is determined as experimentally as possible, providing thermal expansion of the annular insert, eliminating the incision of the working blades due to the lengthening of their feathers and pulling their locks from the disk under the action of centrifugal forces and due to the dynamic effect of the turbine rotor load.

Note that the use of active regulation of the cooling of the disk and rotor blades of the turbine engine rotor allows the cooling of the annular insert with secondary air with a lower total flow rate of this air at the specified transition mode due to the above-described effect of the counter action of both cooling methods.

The operating conditions of radial clearance control systems are varied, diverse and quite numerous are known systems and methods for managing radial clearances (see above). But the task of developing a fairly simple method that provides more opportunities for effective control of radial clearances than the known ones remains relevant.

Therefore, a significantly simpler than the above proposed method of controlling radial clearances is proposed.

The radial clearance control method according to the patent of the Russian Federation 2425985, is selected for the prototype of this proposed method.

The proposed method of thermal active control of the radial clearances of a single-stage high-pressure turbine comprises adjusting the radial clearance by heating the annular insert with a heater in low gas and transition modes with revolutions higher than the cruise speed, characterized in that all operations of controlling the radial clearances of the theater of operations are performed according to the flight cycle on-board computer teams according to a program developed on the basis of an experiment for the entire flight cycle, providing higher values of efficiency, С _p and its traction characteristics and engine operation safety, excluding the insertion of rotor blades and failure of take-off thrust, which is performed in the following sequence: before starting the engine, the heater turns on, which turns off after a certain period of time, the engine starts, if installed at the outlet of the cooled air supplied from the second circuit into the internal cavity of the annular insert, a normally open electro-pneumatic valve, this air is supplied to the internal cavity to face insert in transient forced modes, in the idle mode as a part of engine boost modes, take-off mode, transient modes from take-off to cruising and in cruising mode, then the electropneumatic valve closes and the cooling air supply from the second circuit stops and the heater heating the ring insert is turned on at throttling transitional modes, the heater is turned off until the engine stops at coasting speed, eliminating the possibility of cutting the blades, and in case of installation at the outlet of the cooled air supplied from the second circuit to the inner cavity of the annular insert, normally closed electro-pneumatic valve, after the heater is turned off when the engine is started, the electro-pneumatic valve opens and keeps open in all transient forced modes, in the idle mode as part of the engine boost modes, take-off mode, transient conditions from take-off to cruising, then the electro-pneumatic valve deenergizes and closes and the cooling air of the second circuit does not enter into the internal cavity of the annular insert until the engine is completely stopped, and in cruising mode the annular insert is cooled passively only by the air coming from the last stage of the HPC, the heater is turned on, heating the annular insert on the throttling transition modes, the heater is turned off until the engine completely stops at coasting speed excluding the possibility of cutting the working blades, and the excited vibrations of the annular insert and the parts attached to it at all engine operating modes are effectively damped I have special dampers.

A method using a normally open electro-pneumatic valve can be used if it is not possible to ensure good coordination of the temperature expansions of the stator and turbine rotor when a sufficiently high stator cooling rate is required in cruising mode.

A method using a normally closed electro-pneumatic valve is advisable to apply when the temperature expansion of the stator and turbine rotor is well coordinated, and in the cruise mode passive cooling of the stator and turbine rotor by air is sufficient due to the last stage of the compressor.

As a prototype of the following proposed method for controlling the radial clearances of a high-pressure turbine, the method according to RF patent 2519127 is selected.

The proposed method of thermal active control of the radial clearances of a high-pressure turbine comprises adjusting the clearance according to the instructions of the on-board computer receiving signals from sensors measuring the radial clearance, characterized in that the computer issues commands based on the signals of the sensors measuring the radial clearance, fixed in the ring insert of a single-stage turbine engine or in ring insert of the second stage of a two-stage theater of operations, before starting the engine, the computer determines the value of the smallest of the gaps measured by the sensors, in Luciano heaters and the annular insert before starting the engine are heated by heaters to obtain the value of the radial clearance δ _i, the least possible, but a larger total quantity blades drawing and the hinge portion of the disc the high pressure turbine by the centrifugal force of the turbine stage, in which this sum is greater, the engine starts, and if the measured radial clearances increase during forced flight cycle modes, the heater turns off, the computer at each step of the change in cooling intensity After the ring inserts from the recorded signals from each sensor, it forms its own group of n = 1, 2, 3, ... periods of change in the size of the gaps, and in each period calculates the difference between the largest and smallest values of these gaps, separately selects from all groups for turbine stage sensors the largest difference, determines the corresponding smaller gap and compares its value with the minimum allowable and possible value of this gap δ _min for this engine, if this smaller gap is greater than the minimum allowable value If this gap is δ _min , then, at the command of the computer, the heater, if it was turned on, is turned off, and the element that regulates the flow of cooling air coming from the second circuit into the internal cavities of the ring inserts - the spool, moves at each step of regulation and opens a certain area of the holes, through which this cooling air is discharged, in transient forced modes and in stationary mode - low gas, these operations are continuously repeated while the smaller clearance defined in this operation is at the second stage of the tour the bins will not reach the minimum permissible and possible value for a given engine, the maximum intensity of cooling of the annular inserts is achieved during take-off operation when these holes are fully open, and thrust failure is avoided during engine take-off, and cooling of the disk and rotor blades of a stage of a single-stage turbine engine or the first stage of the rotor because of the last stage of the HPC, a two-stage theater of operations by air occurs passively at all engine operating modes - without computer control commands, or at take-off mode IU and transient conditions from takeoff to cruise mode is active - cooling off these parts, and in cruising mode, at heights greater than the height H of the barostat restriction valve position required to maintain the range of δ _min ≤δ≤δ _min +0, 1 mm of the smallest measured radial clearance value is maintained until the end of the cruise mode in the absence of power to the electromagnetic drive, if there is no such extreme change in flight conditions that unacceptably change the smaller clearance, computer, in this case, increase / decrease this gap by switching on the heaters or shifting the spool to a new saved position, ensuring equality with a tolerance of +0.1 mm of the smaller gap to the minimum allowable and possible for a given engine, and on throttled engine modes, the spool is fully blocks the openings supplying cooling air from the second circuit to the internal cavities of the ring inserts, the heaters and the computer are switched on at each step of the change in the same way as in forced modes, selects a smaller gap, controls the on / off of heaters and changes in the intensity of heating until a smaller gap is set in the interval δ _min ≤δ≤δ _min + 0.1 ÷ 0.2 mm, heaters they are switched off until the engine completely stops at coasting speed, which excludes the possibility of cutting the blades, and at repeated start-up, at its beginning, the computer compares the smaller clearance chosen in a known manner between the annular insert and the blades of a single-stage turbine or between the annular insert and the rotor blades of the second stage two-stage turbines, the magnitude of the radial clearance δ _i, the least possible, but at most 0.1 ÷ 0.2 mm total drawing of rotor blades and the hinge portion of the disc by centrifugal force single-stage turbine wheel or the wheels of the second stage of a two-stage turbine, and if the smaller clearance is greater than this value, then the radial clearance is controlled by changing the cooling intensity, as well as in forced modes of the engine, and if less, then at the beginning, before the engine is restarted, the heaters turn on, and when the gap reaches a value of 0.1 ÷ 0.2 mm, a larger value δ _c , the radial clearance is further controlled by changing the cooling intensity, as well as in forced engine modes, and the excited oscillations of the radial control system Clearances at all engine operating modes are effectively damped by special dampers.

In our opinion, with the design values of the parameters of the proposed high-pressure turbines and reasonably acceptable values of the parameters δ _min and δ _{c, the} use of the proposed method of thermal active control of the radial clearance of the theater of high pressure in many modern gas turbine engines is able to eliminate dangerous situations at all engine operating modes, such as insertion of rotor blades rotor into the stator, failure of thrust on take-off and ensure high engine efficiency and good values of C _{p of the} engine at all engine operating modes. Moreover, these parameter values for the proposed turbines can be determined by calculation.

Note that in some cases when the proposed methods turn out to be unsuitable, in our opinion, it will be possible to use them after some editing, or some of the ideas and constructive solutions suggested above may be useful.

The proposed design of high pressure turbines, the proposed methods of thermal active regulation of the radial clearances of high pressure turbines are illustrated by illustrations. In FIG. cooling of the working blades of the first stages of the proposed turbines is not shown, and electrical, electronic equipment, electrical communications, gas equipment, pipelines and gas fittings are shown conventionally.

In FIG. _Figure 1 shows the change in the radial dimensions of the rotor δ _rp and stator δ _rc , as well as the radial clearance δ _rz in the turbine ND of the NK - 8 - 2U engine in transient conditions: a - low gas, b - throttle response, c - stop, d - repeated throttle response, e - take-off mode, ƒ - cruising mode; 1, ..., 8 - characteristic points, 3 - failure of thrust on take-off, 8 - incision of working blades.

In FIG. 2 shows a longitudinal section along AA in FIG. 3 single-stage high-pressure turbines of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine.

In FIG. 3 shows a section along BB in FIG. 2 and FIG. 14.

In FIG. 4 is a view along page B in FIG. 2.

In FIG. 5 shows the fastening of the spring package in the end supports in cross-section along DG in FIG. 4. In FIG. 84 - spring end support bracket.

In FIG. 6 shows a section along DD in FIG. 5.

In FIG. 7 shows a section along DD in FIG. 5 end supports of springs in the case when the semicircular recesses on the sides of the package of springs are made with a straight section.

In FIG. 8 shows a spool valve and an electromagnetic actuator in longitudinal section, the spool shown in its position in cruising mode at heights exceeding the height of the barostat N. The position of the spool in take-off mode is shown in FIG. dash-dot line with two dots. The position of the spool, completely blocking the cooling air outlet, is depicted by a dotted line. In FIG. 85 - input channels, 86 - output channel, 87 - plug, 88 - swivel, 89 - rod, 90 - guide key, 91 - induction coil, 92 - heat-insulating gasket.

In FIG. 9 is a schematic diagram of a radial clearance control system based on the signals of the engine speed sensor and the barostat according to the method of claim 13. In FIG. 93 - microswitch.

In FIG. 10 shows a schematic diagram of a radial clearance control system based on the signals of sensors 17 measuring the size of the radial clearance by rotor blades according to the method of claim 15. In FIG. 86 - output channel, 93 - microswitch, 94 - the last stage of the compressor, from which air is taken to cool the disk of the first stage of the theater, 95 - tightening device, 96 - damper drive, 97 - damper switch, 98 - impeller of the second stage of the theater, 99 - regulator of heating intensity.

In FIG. 11 shows a section along EE in FIG. 3.

In FIG. 12 shows the layout of a spring package assembled from smooth and corrugated tapes of the same thickness h.

, где k=5÷10.In FIG. 13 shows the layout of the spring package, assembled from smooth and corrugated tapes of the same thickness h and two plates with a thickness

where k = 5 ÷ 10.

In FIG. 14 is a longitudinal section along AA in FIG. 3, but a two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine.

In FIG. 15 is a view along page B in FIG. fourteen.

In FIG. 16 shows a cross section along EE in FIG. 3, but a two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine.

In FIG. 17 shows the mounting of sensors measuring radial clearance in the sockets of the annular insert of the second stage of the theater.

The proposed single-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 2) contains one cooled stage 1 with a nozzle apparatus 2, a turbine rotor 3 with a cooled impeller 4, a turbine stator 5, containing at least at least two turbine bodies 6 and 7 and a radial clearance control system comprising an annular insert 8 above the impeller of the turbine 4, covering working blades 10 of the turbine rotor with an annular radial clearance 9, and elastically and hermetically fastened to the parts forming the inner casing of the turbine 6, a heater 11, covering the annular insert 8 with the possibility of heating, an air intake 12, a cooling air flow regulator 13 with a drive (see Fig. 8), an on-board computer 14 and sensors 15, 16 and 17 (see Figs. 9 and 10). The heater 11, the drive of the flow regulator 13 and the sensors 15, 16 and 17 are electrically connected to the on-board computer 14. The annular insert 8 (see Fig. 2) is hollow and placed in the cavity 18, separated from the second circuit of the engine by the housing 7 and a vertical wall 19, jointly mounted on the housing 20 of the combustion chamber 21. The annular insert 8 annular protrusions 22, made on its sides, is tightly fixed with the possibility of thermal expansion in the reciprocal annular grooves 23 of the vertical wall 19 and the flange 24, made on the inner part 25 of the outer casing of the theater 7. The nozzle apparatus 2 with annular protrusions 26 is also tightly fixed in the reciprocal annular grooves 27 of the vertical wall 19 and the housing 20 of the combustion chamber, in which the holes 28 are equally distributed around the circumference in the vertical wall 19, through which the combustion chamber 21 receives secondary air in the cavity 29 and 18 above the nozzle apparatus 2 and the annular insert 8. The disk 30 and the blades 10 of the turbine engine wheel 4 are also cooled by the secondary air swirling the twist device (in FIG. not shown) before entering the disk blade 30. The outer ring 31 of the nozzle apparatus 2 (see Fig. 2), the annular insert 8 and the inner part 25 with the flange 24 of the outer housing of the fuel assembly 7 form the inner sealed housing of the fuel assembly 6. The ring heater 11 is a microwave or resistive or induction consists of two separate half rings 32 (see Fig. 3), each made in the form of a metal casing 33, inside of which a heating element 34 is fixed. Each half ring 32 of the heater 11 is mounted on the ring insert 8 with the possibility of radial thermal expansion together with the annular insert 8 and the tangential thermal expansion relative to the annular insert 8 by means of a bayonet connection 35 and dowels 36 located with interference in the mating grooves of the annular insert 8 and each half ring 32 in its average cross section. Two nozzles are screwed into the annular insert 8 (see Fig. 2) diametrically opposite with an interference fit on a conical pipe thread 37 — a nozzle — an air intake 12 for supplying cooling air from the second circuit to the inner cavity 38 of the annular insert 8 and a nozzle 39 for discharging this air into the second circuit for cooling the nozzle, or for other purposes. Each of the nozzles 12 and 39 passes through an opening 40 in the outer housing of the fuel assembly 7 and in the housing part 41, which is hermetically fixed to the housing of the fuel assembly 7 using the gasket 42, screws 43 and lock washers 44 (see Fig. 2 and 4). The connection of each nozzle with this body part is sealed with two pairs of piston rings 45 (see Fig. 2), and in each pair of piston rings 45, the cuts of these rings are diametrically opposed. A support 47 is screwed onto each of the nozzles 12 and 39 along the thread 46, or this support is integral with each of the nozzles (not shown in Fig.). To the support 47, symmetrically to each branch pipe (see Figs. 2 and 4), lock washers 44 and screws 43 fasten the central supports 48 of two springs 49, made in the form of a multilayer package 50, compressed by a distributed load, recruited from steel, red-hot or caked, polished tapes 51, made of stainless steel, coated with a wear-resistant coating, and the springs 49, with their end supports 52, are mounted in the second circuit on the outer housing of the theater 7, so that the longitudinal axis of the package 50 is perpendicular to the axis of the engine, and when the volume due to the elastic deformation of the package created the required elastic additional fixing force of the annular insert 8. Package 50 (see Fig. 2 and 5) in the Central support 48 and two end supports 52 of the spring 49 is fixed motionless with two rivets 53 located on the sides of the package without a gap in the semicircular recesses 54 (see Fig. 6) made in the tapes 51, or the package 50 in the end supports 52 is fixed with the possibility of displacement of its ends along its longitudinal axis within the straight section 55 (see FIG. 7), made in this case in the central part of the semicircular recesses 54. The free end of the pipe — the air intake 12 (see FIG. 2) is obliquely cut from the side of the air flow in the second circuit, and the pipe 39 of the air outlet is connected to the pipe 56, at the outlet of which is installed either normally open 57 or normally closed solenoid valve 58 (shown in Fig. 2 conventionally), and the radial clearance control is performed according to the instructions of the on-board computer 14 (see Fig. 9) according to the method of claim 14 of the claims, or a cooling flow regulator air ha - a spool valve 13 with the electromagnetic actuator 60, which differs from the one shown in FIG. 8 and 10 only by the fact that it is made with only one inlet channel (not shown in FIG.), Designed so that the position of the spool 59 regulating the air flow in cruising mode at heights exceeding the height of the limit on the barostat H is fixed when de-energized electromagnetic drive 60. Moreover, the electromagnetic drive 60 is protected against rotation and with a return spring 61. Turning the heater 11 on and off and adjusting the intensity of its heating, opening and closing the supply of cooling air from the second The tour and the intensity of this supply are based on the instructions of the on-board computer 14, generated by the program according to the signals of the sensors - engine speed sensor 15 and barostat 16 according to the method of claim 13, or sensors 17 measuring the size of the radial clearance by working blades according to the method of claim 15 inventions (see FIG. 10). In the outer casing 62 (see Fig. 2) of the second engine circuit above the arrangement of the springs 49, rectangular hatches 63 are made, hermetically sealed with covers 64, and in the outer casing of the theater 7 there is a pipe 65 (see Fig. 11) connected by a pipeline to the intermediate stage An HPC (not shown in FIG.) Supplying cooling air for cooling the front nozzle apparatus 66 of the medium pressure turbine (see FIG. 2).

A one-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 2) is proposed, in which the cooling air supply system to the disk blade 30 and the working blades 10 of the turbine engine 3 includes a pipe 67 with a shutter 68 (see Fig. .10), connecting the cavity 69 behind the last stage of the HPC with the cavity 70 in front of the winding device, and the turbine radial clearance control system itself is controlled according to the method of claim 13 with active cooling of the disk web, and whether according to the method of claim 15 of the claims.

A one-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 2) is proposed, in which a layer of abradable material 71 is deposited on the inner working surface of the annular insert 8.

A one-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 2) is proposed, in which each spring 49 is assembled from direct tapes 51 and is compressed by a uniformly distributed load - air pressure in the second engine circuit.

A one-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 2) is also proposed, in which each spring 49 is assembled in the following arrangement (see Fig. 12): its package 50 is assembled from tapes of the same thickness , two, three or more smooth ribbons 51 are installed in its center, packages 72 are installed on them on both sides, assembled “corrugations into corrugations” of two or more corrugated ribbons 73 so that the corrugation vertices of one package 72 rest on a smooth package 74 tapes 51 in those In cross sections, as the vertices of the second packet 72, and the step of the corrugations of the corrugated ribbons 73 is chosen so that in the span of the packet 50 there is only one vertex resting on the packet 74 of smooth ribbons in the middle of the span, and under each spring support there is only one vertex of the corrugation, and on packages 72 of corrugated tapes 73, packages 75 are assembled from one, two or more smooth tapes 51, and in the assembled spring corrugations of corrugated tapes 73 are fully straightened.

, где k=5÷10, h - толщина внутренних лент пакета 50, и шаг гофрированных лент 73 пакета 50 выбран таким, что в каждом пролете располагаются одна, две и более вершин гофров, и в собранной рессоре гофры гофрированных лент 73 полностью выпрямлены.In addition, a one-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 2) is proposed, in which for packages 75 of smooth tapes 51 or directly on packages of 72 corrugated tapes 73 (see Fig. 13) set on one smooth tape 76 with a thickness

, where k = 5 ÷ 10, h is the thickness of the inner tapes of the bag 50, and the step of the corrugated tapes 73 of the bag 50 is chosen so that in each span there are one, two or more vertices of the corrugations, and in the assembled spring the corrugations of the corrugated tapes 73 are completely straightened.

A two-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 14) is proposed, comprising two cooled stages 1 with nozzle devices 2 with cavities above and below them and a turbine rotor 3 with a cooled impeller 4 the first stage, as well as the turbine stator 5, containing at least two turbine bodies 6 and 7 with cavities between them, and into the cavity of the first stage of which compressed air flows due to the last stage of the high-pressure compressor niya, and a system for regulating the radial clearance of each stage of the turbine, containing an annular insert 8, above the impeller of the turbine, covering working blades 10 of the rotor of the turbine 3 with an annular radial clearance 9, and elastically and hermetically fastened to the parts forming the inner casing of the turbine 6, heater 11 covering an annular insert 8 with the possibility of heating, an air intake 12, a flow rate regulator for cooling air 13 with a drive 60, and sensors 17 (see FIG. 10). The heaters 11, the drive 60 of the flow controller are electrically connected to the on-board computer 14. The annular insert 8 (see Fig. 14) of the first stage of the turbine is hollow and placed in the cavity 18, separated from the second circuit of the engine by the housing 7 and a vertical wall 19, jointly fixed on the housing 20 of the combustion chamber 21. An annular insert 8 with annular protrusions 22 made on its lateral sides is tightened with thermal expansion in the reciprocal annular grooves 23 of the vertical wall 19 and the flange 24, flax on the inner part 25 of the outer casing 7 of the first stage of the theater. The nozzle apparatus 2 with annular protrusions 26 is also tightly fixed in the reciprocal annular grooves 27 of the vertical wall 19 and the housing 20 of the combustion chamber 21, in which the holes 28 are equally distributed around the circumference in the vertical wall 19, through which secondary air enters the cavity from the combustion chamber 21 29 and 18 above the nozzle apparatus 2 and the annular insert 8. The disk 30 and rotor blades 10 of the wheel 4 of the first stage of the theater are also cooled by secondary air swirling by a twisting device (not shown in Fig.) Before entering and a cloth of a disk. The ring heater 11 — microwave, either resistive or induction, consists of two separate half rings 32 (see FIG. 3), each made in the form of a metal body 33, inside of which a heating element 34 is fixed. Each half ring 32 of the heater 11 (see FIG. 3 and 14) is mounted on the annular insert 8 with the possibility of radial thermal expansion together with the annular insert 8 and tangential thermal expansion relative to the annular insert using the bayonet connection 35 and the dowels 36 located with an interference fit in the mating grooves of the count the front insert 8 and each half ring 32 in its average cross section. In the annular insert 8 (see Fig. 14) two nozzles — a nozzle — an air intake 12 for supplying cooling air from the second circuit to the inner cavity 18 of the annular insert 8 and a pipe 39 for discharging this air to the second loop are screwed diametrically opposite with an interference fit on the conical thread or to cool the nozzle, or for other purposes. Each nozzle passes through an opening 40 in the outer housing 7 of the first stage of the theater and in the housing part 41, which is hermetically fixed to this theater by means of a gasket 42, screws 43 and lock washers 44. The connection of each nozzle to this housing part is sealed with two pairs of piston rings 45, and in each pair of piston rings 45, the cuts of these rings are diametrically opposed. A support 47 is screwed onto each of the nozzles 12 and 39, or this support is integral with each of these nozzles (not shown in Fig.). To this support, symmetrically, the pipe with lock washers 44 and screws 43 (see Fig. 5) are attached to the central supports 48 of two springs 49, made in the form of a multilayer package 50, compressed by a distributed load, recruited from steel, red-hot or caked, polished tapes 51 made stainless steel coated with wear-resistant coating. The springs 49 themselves (see FIGS. 14 and 15) of the first stage of the theater are mounted with their end supports 52 in the second circuit on the outer casing of the combustion chamber 77, the outer casing 7 of the first stage of the theater and the outer casing of the second stage of the theater 78 so that the longitudinal axis of the packet 50 is parallel to the axis of the engine, and due to the elastic deformation of the package, the required additional elastic additional fixing force of the annular insert 8 is created, one of the end supports 52 being made common for both springs 49 and the other common for four springs (see Fig. 15 ) - two springs 49 of the first stage and two structurally similar springs 49 of the second stage, in which a common second end support 52 is mounted on the outer casing 78 of the second stage of the theater. The bag 50 (see Fig. 5) in the central support 48 and two end supports 52 of each spring 49 is fixedly fixed with two rivets 53 located on the sides of the bag without a gap in the semicircular recesses 54 (see Fig. 6) made in tapes 51, or the package 50 in the end supports 52 is fixed with the possibility of displacement of its ends along its longitudinal axis within the straight section 55 (see Fig. 7), made in this case in the central part of the semicircular recesses 54. The design of the ring insert 8, pipes 12 and 39, body parts 41 of the second stage of the theater (see Fig. 14), their crepe pouring and compaction is similar to these elements of the first stage of a theater. The free end of each pipe — the air intake 12 from the side of the air flow in the second circuit — is obliquely cut off, and each pipe 39 of the air outlet is connected to a pipe 56, at the outlet of which either a normally open 57 or a normally closed electro-pneumatic valve 58 is installed (fig. . A nozzle 65 (see FIG. 16) is made in the outer casing 7 of the first stage of the theater of operations; it is connected by a pipe to the intermediate stage of the HPC (not shown in FIG.), Which supplies cooling air to cool the nozzle apparatus 2 of the second stage of the theater of operations (see Fig. 14) and through the holes 79 equally spaced around the circumference in the vertical wall 80, fixed together with the outer casing 78 of the second stage of the theater, into the cavity 18 above the annular insert 8 of the second stage of the theater. Moreover, the intermediate stage of the HPC was selected, which provides a slight pressure drop on the outer ring 31 of the nozzle apparatus 2 of the second stage of the theater. And the nozzle apparatus 2 itself is mounted in the stator of the theater of operations 5 with the help of annular protrusions 26, which fit with interference into the reciprocal ring grooves 27, made in the flange 24 of the inner part 25 of the outer housing 7 of the first stage of the theater and the vertical wall 80. The ring insert 8 of the second stage is attached the same vertical wall 80 and vertical wall 81, made integral with the outer casing 78 of the second stage of the theater. The outer ring 31 of the nozzle apparatus 2 of the first stage, its annular insert 8, the outer ring 31 of the nozzle apparatus 2 of the second stage, its annular insert 8, the elements of the outer housings 7 and 78 of the theater, to which these parts are attached, form an internal sealed housing of the theater. The second stage air exhaust pipe 39 is connected to a pipe 56, at the outlet of which either a normally open 57 or a normally closed electro-pneumatic valve 58 is installed, or the cooling air pipes 39 of the first and second stages are each connected to their entrance to the spool valve 13 with an electromagnetic drive 60 (see Fig. 8), designed so that both stages either have a common outlet from the spool distributor 13, or each stage has its own outlet of cooling air from it (not shown in Fig.) and the floor The cruise control valve 59, which regulates the air flow, at heights exceeding the barostat H limit height, was fixed when the electromagnetic drive 60 was de-energized. Moreover, the electromagnetic drive 60 is protected against rotation and with a return spring 61. Turning on and off the heaters 11 ( see Fig. 10) and the regulation of the intensity of their heating, the opening and closing of the supply of cooling air from the second circuit and the intensity of this supply of both stages occur according to the commands of the on-board computer 14, vyr abutable program according to the signals of the sensors 17, measuring the size of the radial clearance 9 by the working blades, according to the method of claim 15, fixed in the ring insert 8 of the second stage (see FIG. fourteen). An annular insert, a nozzle - an air intake, an air outlet pipe, a heater, body parts, central springs, and connections of these parts of the second stage of the theater are structurally similar to these parts and their connections of the first stage, but can only differ in parameters. In the outer casing 62 of the second motor circuit above the location of the springs 49, rectangular hatches 63 are made, hermetically sealed with covers 64.

A two-stage high-pressure turbine of a dual-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Figs. 14 and 10) is proposed, characterized in that the cooling air supply system to the disk blade 30 and the rotor blades 10 of the first stage of the turbine engine contains a pipe 67 s a shutter 68, connecting the cavity 69 behind the last stage of the HPC with the cavity 70 in front of the twisting device (not shown in Fig.) of the first stage, and the control system of the radial clearance of the turbine is controlled by NTU method of claim. 15 claims active leaf disc cooling.

A two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 14) is proposed, characterized in that its springs 49 are made according to any of the above-described variants of its design (see Figs. 14, 12 and 13 )

A two-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Figs. 14 and 17) is proposed, characterized in that a layer of abradable material is applied to the inner working surface of the annular insert 8 of the first stage, or a layer of abrasive material 71 is applied on the inner working surface of the ring inserts 8 of both stages of the theater, in which case the sensors 17 measuring the radial clearance are offset in the sockets 82 of the ring insert 8 of the second stage in the radial directions to the thickness of the abrasive layer.

A two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 14) is also proposed, characterized in that the hydraulic diameters of the duct supplying the cooling air of the second circuit inside the annular insert of the first stage and discharging it are larger than the hydraulic the diameter of a similar tract of its second stage, or vice versa.

In addition, a two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine (see Fig. 14) is proposed, characterized in that the electrical parameters of the half rings 32 with conductors of the resistive heater or coils of the induction heater 11 are different for the first and second stages turbines and heater 11 of each stage of the theater operates according to its own program.

The proposed method of thermal active control of the radial gaps of a single-stage high-pressure turbine (see Fig. 2) comprises radial clearance control operations performed according to the commands of the engine electronic system using signals from the engine speed sensor 15 of the turbo-compressor rotor VD (see Fig. 9) and barostat 16, cooling of the disk 30 and rotor blades 10 of the turbine rotor 3 by air due to the last stage of the HPC is passive - without commands from these sensors, or actively - the cooling of these parts is turned off during take-off mode and transitional zhimah from takeoff to cruise mode (see. Fig. 10). All operations of control of the radial clearances 9 of the theater are performed according to the flight cycle according to the instructions of the on-board computer 14 according to the program developed on the basis of the experiment for the entire flight cycle, providing good values of efficiency, C _p and its traction characteristics and engine operation safety, excluding the incision of the working blades 10 ( see Fig. 2) and failure of the takeoff thrust, which is performed in the following sequence: before starting the engine, heater 11 is turned on, which turns off after a certain period of time, two the igniter starts and in forced modes the cooling of the ring insert 8 is turned on by the air of the second circuit supplied to its internal cavity 37 with a predetermined law of change in speed and time, with an intensity that reaches its maximum value in take-off mode. Moreover, passive cooling of the disk 30 and rotor blades 10 of the turbine rotor 3 is carried out at all engine operating modes and is not regulated by the radial clearance control system. In the regime that is transitioning from the take-off mode to the cruising mode, the ring insert 8 is cooled by the air of the second circuit with an intensity that gradually decreases according to the law specified by the computer program 14 (see Fig. 10) to the level required in the cruise mode at heights exceeding the height of the barostat restriction N, and with the indicated active cooling of the disk 30 and rotor blades 10 of the turbine rotor 3 (see Fig. 2), the predetermined law of cooling the annular insert 8 by the air of the second circuit is fulfilled, which also decreases to the same level at cruising mode, but with large gradients of intensity reduction in engine speed. In the throttled flight cycle modes, the cooling of the annular insert 8 is turned off by the air of the second circuit and its heating is turned on by the heater 11 with an intensity that gradually decreases according to a given law with a decrease in engine speed, and the heater 11 is turned off and the engine is stopped. The excited vibrations of the annular insert 8 and the parts attached to it at all engine operating modes are effectively damped by special dampers - springs 49.

A method of thermal active control of the radial gaps of a single-stage high-pressure turbine (see Fig. 2) is proposed, comprising adjusting the radial clearance by heating the ring insert 8 by the heater 11 in idle and transition modes with revolutions higher than the cruise mode revolutions, which has all control operations the radial clearances 9 of the theater are performed according to the flight cycle according to the instructions of the on-board computer 14 according to the program developed on the basis of the experiment for the entire flight cycle, providing which provides good values of efficiency, С _p and the characteristics of its thrust and engine operation safety, excluding the insertion of working blades 10 and failure of the thrust on take-off, which is performed in the following sequence: before starting the engine, heater 11 is turned on (see Fig. 9), which is turned off through a certain period of time, the engine starts, if the cooled air supplied from the second circuit to the inner cavity 38 of the ring insert 8 is installed at the outlet, a normally open solenoid valve 57, this air is supplied into the inner cavity 38 of the annular insert 8 in transient forced modes, in the idle mode as a part of engine boosting modes, take-off mode, transitional modes from take-off to cruising and cruising. Then, the electro-pneumatic valve 57 is closed and the supply of cooling air from the second circuit is stopped. The heater 11 is turned on, heating the annular insert 8 on the throttling transient modes. The heater 11 is turned off until the engine stops at coasting speed, eliminating the possibility of cutting the blades 10. And if the cooled air supplied from the second circuit to the inner cavity 37 of the ring insert 8 is installed at the outlet, the normally closed electro-pneumatic valve 58 is turned off after the heater 11 is turned off at startup engine electric pneumatic valve 58 opens and keeps open in all transient forced modes, in the idle mode as part of the engine boost modes, take-off mode, transition s mode from takeoff to cruising. Then, the electro-pneumatic valve 58 is de-energized and closed and the cooling air of the second circuit does not enter the internal cavity 38 until the engine is completely stopped, and in cruising mode the annular insert 8 is cooled passively only by the air coming from the last stage of the HPC. The heater 11 is turned on, heating the annular insert 8 on the throttling transient modes. The heater 11 is turned off until the engine stops at coasting speed, eliminating the possibility of cutting the blades 10. The vibrations of the ring insert 8 and the parts attached to it are effectively damped by special dampers - springs 49 (see Fig. 2) at all engine operating modes.

In addition, a method of thermal active control of the radial clearances of a high pressure turbine (see Fig. 14) is proposed, comprising adjusting the clearance 9 by the commands of the on-board computer 14 receiving signals from sensors 17 measuring the radial clearance 9 (see Fig. 10), different the fact that the computer 14 issues commands from the signals of the sensors 17 measuring the radial clearance, fixed in the ring insert 8 of a single-stage theater or in the ring insert 8 of the second stage of a two-stage theater. Before starting the engine, the computer 14 determines the smallest of the clearances measured by the sensors 17, includes heaters 11 (see Fig. 14) and the ring inserts 8 are heated by the heaters 11 before starting the engine until this radial clearance δ _{c is} obtained, possibly smaller but larger extracting the blades 10 and the locking part of the disk 30 of the high pressure turbine under the action of centrifugal forces of that stage of the turbine, in which this total value is greater. The engine starts, and if the measured radial clearances 9 increase on forced flight cycle modes, the heater 11 turns off. Computer 14 at each step of changing the cooling intensity of the ring inserts 8 from the recorded signals from each sensor 17 forms its own group of n = 1, 2, 3, ... periods of change in the size of the gaps 9, and in each period calculates the difference between the largest and smallest values of these gaps , of all groups for sensors 17 of the turbine stage, selects the largest one separately; the difference, determines the corresponding smaller gap and compares its value with the minimum allowable and possible for this engine the value of this gap δ _min . If this smaller gap 9 is greater than the minimum permissible value of this gap δ _min , then, on the command of computer 14, heater 11, if it was turned on, is turned off, and the element that regulates the flow of cooling air coming from the second circuit into the internal cavities 38 of the ring inserts 8 is a spool 59 (see Fig. 8), at each step of regulation, it shifts and opens a certain area of holes 83 through which this cooling air is discharged. In transient forced modes and in stationary mode - low gas, these operations are continuously repeated until the smaller clearance 9 determined in this operation at the second stage of the turbine reaches the minimum allowable and possible value for this engine. The maximum cooling intensity of the annular inserts 8 is achieved in the take-off mode when these openings 83 are fully open, and the failure of the thrust during the take-off mode of the engine is ensured. Moreover, the cooling of the disk 30 and rotor blades 10 of the stage of a one-stage theater of operations or the first stage of the rotor of a two-stage theater of operation by air due to the last stage of the HPC (see Fig. 10) occurs passively at all engine operating modes - without control commands from computer 14, or in take-off mode and transient conditions from takeoff to cruise mode is active - cooling off these parts, and in cruising mode, at heights greater than the height H of the barostat limit 16, the position of the slide 59, required to maintain the range of δ _min ≤δ δ _min +0,1 mm smallest of the measured values of the radial clearance 9 is maintained until the end of the cruise mode in the absence of supply of an electromagnetic actuator 60, if there is no such an extreme change of flight conditions which change unacceptably smaller gap 9 defined by the computer 14. In this case according to the increase / decrease of this gap, the heaters 11 are turned on or the spool 59 is displaced to a new maintained position, ensuring equality with a tolerance of +0.1 mm of a smaller gap 9 to the size of the mini cial possible and permissible for the motor. In throttled engine modes, the spool 59 (see FIG. 10) completely covers the openings 81 supplying cooling air from the second circuit to the internal cavities 38 of the ring inserts 8, the heaters 11 and the computer 14 are turned on at each step of changing the heating intensity in the same way as in forced modes, selects a smaller gap 9, controls the on / off of heaters 11 and changes in the intensity of heating until a smaller gap 9 is established in the interval δ _min ≤δ≤δ _min + 0.1 + 0.2 mm. The heaters 11 are turned off until the engine stops at coasting speed, which excludes the possibility of cutting the blades 10 (see Fig. 14), and when the throttle response is repeated, at its beginning, the computer 14 compares the smaller gap 9 chosen in a known manner between the annular insert 8 and the blades a single-stage turbine or between an annular insert 8 and the working blades of the second stage of a two-stage turbine, with the value of this radial clearance δ _c , possibly smaller, but 0.1–0.2 mm larger than the total exhaust of the working blades 10 and the lock h part of the disk 30 under the action of the centrifugal forces of the wheels of a single-stage turbine or the wheels of the second stage of a two-stage turbine, and if the smaller clearance 9 is greater than this value, then the radial clearance 9 is controlled by changing the cooling intensity, as well as in forced modes of the engine, and if less, first before re-starting the engine heaters 11 are included, and when the lower clearance value at 0.1 ÷ 0.2 mm greater than the value δ _n further regulation by changing the radial clearance Intensity of cooling, as well as for the forced motor modes. The excited vibrations of the radial clearance control system at all engine operating modes are effectively damped by special dampers - springs 49.

The assembly of the proposed theater for the specialist is clear from the description of the proposed structures of the theater and FIG. 1 and 12 and is not specifically described.

The operation of gas turbine turbines is widely known, and the operation of the proposed radial clearance control systems is described in detail in the proposed methods for controlling these gaps, and therefore, the operation of the proposed theater is not specifically described.

The advantages of the proposed turbofan engines and methods for controlling radial clearances are mainly described above and, in our opinion, allow us to draw the following conclusions:

The proposed turbofan engine turbofan engines and radial clearance control methods can significantly expand the range of successful solutions for creating single-stage and two-stage turbofan engine turbofan engines with an active radial clearance control system that is effective in all flight cycle modes, to achieve good efficiency values for these modes, C _p with good mass characteristics High-pressure turbojet, including aircraft turbofan engines with Tr = 1800 ° K.

In the proposed turbofan engines, turbofan engines can be used cell seals, film cooling of the blades of the nozzle apparatus and the working blades of the first stage. These devices are widely known and the options for the proposed theater with these devices are not described.

All of the turbofan engines offered by the turbofan engine can be used in aircraft engines together with other well-known systems for controlling radial clearances KVD, TST, TNT, etc.

In addition, even in cases where the application of the proposed designs and methods will be inappropriate or impossible, it may be useful to use individual original ideas of this development.

Claims (15)

1. A single-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine, comprising one cooled stage with a nozzle apparatus with cavities above and below it and a turbine rotor with a cooled impeller, and a turbine stator containing at least two turbine bodies with cavities between them, into which compressed air enters due to the last stage of the high-pressure compressor, and a radial clearance control system containing an annular insert y, above the turbine impeller, covering the rotor blades of the turbine rotor with an annular radial clearance and elastically and hermetically fastened to the parts that form the turbine’s inner casing, a heater covering the annular insert with the possibility of heating, an air intake, a cooling air flow regulator with a drive, an on-board computer and the sensors, and the heater, the drive of the flow controller and the sensors are electrically connected to the on-board computer, characterized in that the ring insert is hollow and placed on the cavity, separated from the second circuit of the engine by the housing and a vertical wall, jointly mounted on the combustion chamber housing, and the annular insert with annular protrusions made on its sides, is tightened with thermal expansion in the mating annular grooves of the vertical wall and the flange made on the inner part of the outer theater, the nozzle device with annular protrusions is also tightly fixed in the mating annular grooves of the vertical wall and the housing of the combustion chamber, in which holes and a vertical wall are made equally distributed around the circumference through which secondary air enters the cavity above the nozzle apparatus and the annular insert from the combustion chamber, and the disk and impellers of the turbine engine wheel are also cooled by the secondary air twisted by a twisting device before entering the disk blade, and the outer ring of the nozzle apparatus, the annular insert and the inner part with the flange of the outer theater block form an internal sealed theater block, a microwave ring heater, or resistive The induction or induction consists of two separate half rings, each made in the form of a metal case, inside which a heating element is fixed, and each half of the heater is mounted on the ring insert with the possibility of radial thermal expansion together with the ring insert and tangential thermal expansion relative to the ring insert using a bayonet connections with it and dowels located with an interference fit in the reciprocal grooves of the annular insert and each half-ring in its average cross section, in the annular insert is diametrically opposite with an interference fit on a conical pipe thread; two nozzles are screwed in - a nozzle-air intake for supplying cooling air from the second circuit to the inner cavity of the annular insert and a pipe for removing this air to the second circuit or for cooling the nozzle, or for other purposes, and each nozzle passes through an opening in the outer housing of the theater and in the housing part, hermetically mounted on the housing of the theater, the connection of each pipe with this housing part is sealed with two pairs of piston rings ec, and in each pair of piston rings, the cuts of these rings are diametrically opposed, a support is screwed on each nozzle, or this support is made integrally with each of the nozzles, and the central supports of two springs made of the springs are symmetrically mounted to this support with lock washers and screws in the form of a multilayer bag, compressed by a distributed load, recruited from steel, hardened or caked, sanded tapes made of stainless steel, coated with a wear-resistant coating, and the springs themselves end supports are fixed in the second circuit on the outer casing of the theater, so that the longitudinal axis of the package is perpendicular to the axis of the engine, and due to the elastic deformation of the package, the required force of elastic additional fixation of the annular insert is created, and the package in the central support and two end supports of the spring fixed motionless with two rivets located on the sides of the package without a gap in the semicircular recesses made in tapes, or the package in the end supports is fixed with the possibility of displacement to cc along its longitudinal axis within the straight section, made in this case in the central part of the semicircular recesses, the free end of the air intake pipe on the side of the air flow in the second circuit is obliquely cut, and the air exhaust pipe is connected to the pipeline at the outlet of which is installed either normally open or normally closed electro-pneumatic valve, and rectangular hatches are hermetically closed by covers in the outer casing of the second engine circuit above the springs, and in the outer casing of the theater A nozzle connected to the intermediate stage of the HPC by the pipeline supplying cooling air for cooling the front nozzle apparatus of the medium-pressure turbine is filled, and all operations of the control of the radial clearances of the turbine engine are performed according to the flight cycle according to the on-board computer commands according to the program developed on the basis of the experiment for the entire flight cycle, good efficiencies, C _p and its traction characteristics and safety of operation of the engine, eliminating the infeed of rotor blades and rods fail during takeoff, Kotor I am executed in the following sequence: before starting the engine, the heater turns on, which turns off after a certain period of time, the engine starts, and if the cooled air supplied from the second circuit to the inner cavity of the annular insert, a normally open electro-pneumatic valve is installed at the outlet, this air is supplied to the inner cavity of the annular insert in transient forced modes, in the idle mode as part of the engine boost modes, take-off mode, transitional presses from the take-off to cruising and cruising mode, then the electro-pneumatic valve closes and the cooling air supply from the second circuit is stopped, and the heater is turned on, heating the annular insert on the throttled transition modes, the heater is turned off until the engine stops at coasting speed, eliminating the possibility of cutting the blades, and in the case of installation at the outlet of the cooled air supplied from the second circuit into the internal cavity of the annular insert, normally closed electropneum valve, after the heater is turned off when the engine is started, the electro-pneumatic valve opens and keeps open in all transient forced modes, in the idle mode as part of the engine boost modes, take-off mode, transient modes from take-off to cruising, then the electro-pneumatic valve deenergizes and closes and the secondary cooling air does not enters the internal cavity until the engine is completely stopped, and in cruising mode the annular insert is passively cooled only by air, due to the last stage of the HPC, the heater is turned on, heating the annular insert on the throttled transition modes, the heater is turned off until the engine stops at coasting speed, eliminating the possibility of cutting the blades, and in the case when the radial clearance control operations are also performed according to the on-board computer commands according to the signals of the rotor speed sensor of the VD turbocompressor and the barostat, cooling of the disk and rotor blades of the turbine engine rotor with air due to the last stage of the HPC is passive - without commands of these sensors, or actively - the cooling of these parts is turned off during take-off mode and transitional modes from take-off mode to cruising mode, and the spool valve with electromagnetic drive is designed so that the position of the spool regulating the air flow in cruising mode at altitudes exceeding the limit height according to the barostat H, it is fixed with a de-energized electromagnetic drive, and the electromagnetic drive is made with anti-rotation protection and with a return spring, the commands are executed in running sequence: before starting the engine, the heater also turns on, which turns off after a certain period of time, the engine starts and, in forced modes, the cooling of the annular insert is turned on by the air of the second circuit supplied to its internal cavity with a given law of change in speed and time, the intensity that reaches the maximum values in the take-off mode, and passive cooling of the disk and rotor blades of the turbine engine rotor is performed at all operating modes of the engine and not it is controlled by the radial clearance control system, and in the regime that is transitioning from the take-off mode to the cruising mode, the ring insert is cooled by the air of the second circuit with an intensity that gradually decreases according to the law set by the computer program to the level required in the cruise mode at heights exceeding the height of the barostat limit N and, with the indicated active cooling of the disk and rotor blades of the theater rotor, the specified law of cooling the annular insert by the air of the second circuit is also satisfied, which also decreases at the same level in cruise mode, but with large gradients of intensity reduction in engine speed, and in throttled flight cycle modes, the cooling of the annular insert is turned off by the air of the second circuit and its heating is turned on by a heater with an intensity that gradually decreases according to a given law with a decrease in engine speed, and off heater and engine shutdown, and in the case when the radial clearance of the high pressure turbine is controlled by the commands of the on-board computer receiving the signal s from sensors which measure the radial clearance embodied in an annular insert theater, computer before starting the engine determines the value of the smallest of the measured air gaps includes a heater and an annular insert before starting the engine is warming heaters to obtain the value of the radial clearance δ _i, the least possible, but more the total amount of exhaust of the blades and the locking part of the high-pressure turbine disk under the action of centrifugal forces, the engine starts, and if in forced flight modes of the cycle, the measured radial gaps grow, the heater turns off, the computer at each step of changing the cooling intensity of the ring inserts from the recorded signals from each sensor forms its own group of n = 1, 2, 3, ... periods of change in the size of the gaps, and in each period calculates the difference between the largest and the smallest values of these gaps, from all groups for sensors separately selects the largest difference in magnitude, determines the corresponding smaller gap and compares its value with the minimum allowable and possible for I of this engine with the value of this gap δ _min , if this smaller gap is greater than the minimum permissible value of this gap δ _min , then, on a computer command, the heater, if it was turned on, is turned off, and the element that regulates the flow of cooling air coming from the second circuit into the internal cavity annular insert - spool, at each step of regulation it shifts and opens a certain area of openings through which this cooling air is discharged, in transient forced modes and in stationary mode - small ha e these operations are continuously repeated until the smaller clearance defined in this operation reaches the minimum allowable and possible for a given engine, the maximum cooling rate of the annular insert is reached in take-off mode when these holes are fully open, and thrust failure is prevented during take-off mode of the engine moreover, the cooling of the disk and the working blades of the stage of the theater of air with air due to the last stage of the HPC is passive at all engine operating modes - without control commands a, or in the take-off mode and transitional modes from the take-off mode to the cruising mode, it is active - the cooling of these parts is turned off, and in the cruising mode at heights greater than H barostat restrictions, the position of the spool required to maintain in the interval δ _min ≤δ≤δ _min +0,1 mm smallest of the measured values of the radial gap is maintained until the end of the cruise mode in the absence of supply of an electromagnetic actuator, if there is no such an extreme change of flight conditions which change unacceptably m the smallest clearance determined by the computer, in this case, the heater is turned on / off according to the increase / decrease, or the slide valve is shifted to a new maintained position, ensuring equality with a tolerance of +0.1 mm of the smaller clearance to the minimum allowable and possible for a given engine, and in throttle modes the engine spool completely covers the holes supplying cooling air from the second circuit to the internal cavities of the annular inserts, the heater and computer are turned on at every step and changes the heating intensity in the same way as in forced modes, selects a smaller gap, controls the on / off heater and changes the heating intensity until a smaller gap is set in the interval δ _min ≤δ≤δ _min + 0.1 ÷ 0.2 mm , the heaters are turned off until the engine stops at coasting speed, eliminating the possibility of cutting the rotor blades, and at repeated start-up, at its beginning, the computer compares the smaller gap chosen in a known manner between the annular insert and the turbine rotor blades s with the magnitude of the radial clearance δ _i, the least possible, but at most 0.1 ÷ 0.2 mm total drawing of rotor blades and the hinge portion of the disc by centrifugal force of the turbine wheel, and if a smaller clearance is more than this value, the regulation of the radial clearance is carried out by changing the cooling intensity, as well as in forced modes of the engine, and if less, then first, before restarting the engine, the heaters turn on, and when the smaller gap reaches a value of 0.1 ÷ 0.2 mm greater than δ _c , further regulation Radial clearance is controlled by changing the cooling intensity, as well as in forced engine modes, and the excited vibrations of the annular insert and the parts attached to it at all engine operation modes are effectively damped by spring dampers. 2. A single-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to claim 1, characterized in that the cooling air supply system to the blade web and the working blades of the turbine rotor contains a pipeline with a shutter connecting the cavity behind the last stage of the HPC with cavity in front of the swirling device in the case when radial clearance control operations are performed according to the on-board computer commands according to the signals of the turbocharger rotor speed sensor VD and the barostat, and in the case when the radial clearance of the high-pressure turbine is controlled by the commands of the on-board computer, receiving signals from the sensors measuring the radial clearance, fixed in the ring insert of the theater, the cooling air supply due to the last stage of the HPC to the disk blade and workers the blades of the rotor of the theater are active. 3. A single-stage high pressure turbine of a dual-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to any one of paragraphs. 1 and 2, characterized in that on the inner working surface of the annular insert is applied a layer of abradable material. 4. A single-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to claim 3, characterized in that each spring is assembled from straight tapes and is compressed by a uniformly distributed load - air pressure in the second circuit of the engine. 5. A single-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to claim 3, characterized in that each spring is assembled in the following arrangement: its package is assembled from tapes of the same thickness, two, three or more are installed in its center smooth ribbons, on both sides of them are packages assembled “corrugations into corrugations” of two or more corrugated ribbons in such a way that the vertices of the corrugations of one packet rest on the packet of smooth ribbons in the same sections as the vertices of the second package, and the step of the corrugations of the corrugated tapes is chosen so that only one vertex is located in the span of the package, resting on the package of smooth ribbons in the middle of the span, and under each spring support there is only one vertex of the corrugation, and packets assembled from one , two or more smooth tapes, and in the assembled spring the corrugations of the corrugated tapes are fully straightened. 6. A single-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to claim 5, characterized in that one smooth tape with a thickness of one is installed on packages of smooth tapes or directly on packages of corrugated tapes

, where k = 5 ÷ 10, h is the thickness of the inner ribbons of the bag, and the step of the corrugated ribbons of the bag is chosen so that in each span there are one, two or more vertices of the corrugations, and in the assembled spring the corrugations of the corrugated ribbons are completely straightened. 7. A two-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine, comprising two cooled stages with nozzle devices with cavities above and below them, and a turbine rotor with a cooled impeller of the first stage, as well as a turbine stator containing at least at least two turbine bodies with cavities between them, into the cavity of the first stage of which compressed air enters due to the last stage of the high-pressure compressor, and a radio control system the gap of each stage of the turbine, containing an annular insert, above the impeller of the turbine, covering the rotor blades of the turbine rotor with an annular radial clearance and elastically and hermetically fastened to the parts forming the turbine’s inner casing, a heater covering the annular insert with the possibility of heating, an air intake, a regulator flow rate of cooling air with a drive, and sensors, and heaters, drives of the flow regulator are electrically connected to the on-board computer, characterized in that the annular the rate of the first stage of the turbine is hollow and placed in a cavity separated from the second circuit of the engine by a housing and a vertical wall jointly fixed to the combustion chamber body, and the annular insert with annular protrusions made on its lateral sides is tightened with thermal expansion in the reciprocal annular grooves of the vertical wall and flange made on the inner part of the outer casing of the first stage of the theater, the nozzle apparatus with ring protrusions is also tightened in response ring grooves of the vertical wall and the housing of the combustion chamber, in which holes are equally distributed around the circumference in the vertical wall, through which secondary air enters the cavity from the combustion chamber above the nozzle apparatus and the annular insert, and the disk and rotor blades of the wheel of the first stage of the turbine engine are also cooled the secondary air, twisted by a tightening device before entering the disk blade, the microwave ring heater, either resistive or induction, consists of two separate half rings, you each filled in the form of a metal casing, inside which the heating element is fixed, and each half of the heater is mounted on the annular insert with the possibility of radial thermal expansion together with the annular insert and tangential thermal expansion relative to the annular insert using a bayonet connection with it and dowels located with an interference fit mating grooves of the annular insert and each half-ring in its average cross section, into the annular insert is diametrically opposite with an interference fit two branch pipes are screwed into the tapered thread - a pipe-air intake for supplying cooling air from the second circuit to the internal cavity of the annular insert and a pipe for removing this air to the second circuit or to cool the nozzle or for other purposes, and each pipe passes through an opening in the outer casing of the first stages of the theater and in the housing part, hermetically mounted on this housing of the theater, the connection of each nozzle with this housing part is sealed with two pairs of piston rings, and in each pair of piston rings the cuts of these rings are diametrically opposed, a support fixed by a lock washer and a round nut is screwed on each nozzle, and the central bearings of the two springs of the first stage of the theater are mounted symmetrically to the nozzle with lock washers and screws, made in the form of a multilayer packet compressed by a distributed load, recruited from steel, red-hot or caked, polished tapes made of stainless steel, coated with a wear-resistant coating, and these springs themselves are fixed with their end supports len in the second circuit on the outer casing of the combustion chamber and on the outer casing of the first stage of the theater, so that the longitudinal axis of the package is parallel to the axis of the engine, and due to the elastic deformation of the package created the required force of additional elastic fixation of the annular insert, one of the end the supports are made common for both springs of the first stage, and the other is common for four springs - two springs of the first stage and two structurally similar springs of the second stage, which have a common second end support beyond is replicated on the outer casing of the first stage of the theater, and the second end support, common to the two springs of the second stage, is fixed on the outer casing of the second stage of the theater, and the package of each spring in the central support and two end supports is fixed motionless with two rivets located on the sides of the package without a gap in the semicircular recesses made in tapes, or the package in the end supports is fixed with the possibility of displacement of its ends along its longitudinal axis within the straight section made in this case in the central part of the semicircles of the recesses, the free end of the air intake pipe on the side of the air flow in the second circuit is obliquely cut, and the air exhaust pipe is connected to the pipeline, at the outlet of which either a normally open or normally closed electro-pneumatic valve is installed, and a pipe is made in the outer casing of the first stage of the theater, connected by a pipeline to the intermediate stage of the HPC, supplying cooling air for cooling the nozzle apparatus of the second stage of the HPT and through openings equally spaced around the circumference in the vertical wall, replicated together with the outer casing of the second stage of the theater of operations, into the cavity above the annular insert of the second stage of the theater of operations, and the intermediate stage of the HPC is selected, which provides a slight pressure drop on the outer ring of the nozzle device of the second stage of the theater of operations, and the nozzle device is mounted in the stator of the theater of operations with ring protrusions included with interference in the reciprocal ring grooves made in the flange of the inner part of the outer casing of the first stage of the theater and the vertical wall, the annular insert of the second stage is attached to it the same vertical wall and vertical wall, made in one piece with the outer casing of the second stage of the theater, and the outer ring of the nozzle apparatus of the first stage, its annular insert, the outer ring of the nozzle apparatus of the second stage, its ring insert, the elements of the outer casing of the theater to which are attached these parts form the internal sealed housing of the theater, the annular insert, the air intake pipe, the exhaust pipe, the heater, the body parts, the central springs and the connections of these parts of the second stage of the theater structurally similar to these parts and their first stage connections, but can only differ from them in parameters, and the second stage air outlet pipe is connected to a pipeline at the outlet of which either a normally open or normally closed electro-pneumatic valve or cooling air outlet pipes of the first and second are installed each stage is connected to its entrance to the spool valve with an electromagnetic drive, designed so that each stage either has its own cooling air outlet from the ash otnikovy distributor, or both stages have a common exit from it, and the position of the spool regulating the air flow, at cruising mode at heights exceeding the height of the barostat limit H, was fixed with a de-energized electromagnetic drive, and the electromagnetic drive is made with anti-rotation and with a spring return, and in the outer casing of the engine’s second circuit above the springs, rectangular hatches are made, hermetically sealed with covers, and control of the radial clearances of the high-turbine Aviation is carried out according to the instructions of the on-board computer, which receives signals from the sensors measuring the radial clearance, fixed in the ring insert of the second stage of the two-stage turbine engine, before starting the engine, the computer determines the smallest of the clearances measured by the sensors, turns on the heaters and the ring inserts are heated by the heaters before starting the engine to obtain the value of this radial clearance δ _c , possibly a smaller, but larger total amount of drawing of the blades and the locking part of the high turbine disk under the action of centrifugal forces of that stage of the turbine, for which this total value is greater, the engine starts, and if the measured radial clearances increase during forced flight cycle modes, the heaters turn off, the computer at each step of changing the cooling intensity of the ring inserts from the recorded signals from each sensor forms its group from n = 1, 2, 3, ... periods of change in the size of the gaps, and in each period calculates the difference between the largest and smallest values of these gaps, from all groups for sensors of the turbine stage selectors separately selects the largest difference, determines the smaller gap corresponding to it, and compares its value with the minimum allowable and possible value of this gap δ _min for a given engine, if this smaller gap is larger than the minimum allowable value of this gap δ _min , then by computer command the heaters, if they were turned on, are turned off, and the element that regulates the flow of cooling air coming from the second circuit into the internal cavities of the ring inserts is a spool, at each step of the pressure is shifted and opens a certain area of the holes through which this cooling air is discharged, in transient forced modes and in stationary mode - low gas, these operations are continuously repeated until the smaller clearance defined in this operation at the second stage of the turbine reaches the minimum permissible value and possible for this engine, the maximum cooling rate of the annular inserts is achieved in the take-off mode when these holes are fully open, and thrust failure is eliminated during the engine’s take-off mode, moreover, cooling of the disk and rotor blades of the first rotor stage of the two-stage turbine engine with air due to the last HPC stage occurs passively at all engine operating modes - without computer control commands, or in the take-off mode and transition modes from the take-off mode to the cruising mode - the cooling of these parts is turned off, and in cruise mode at heights greater than H the barostat limits, the position of the spool required to maintain the minimum in the interval δ _min ≤δ≤δ _min +0.1 mm from the measured radial clearance values, it remains until the end of the cruising mode in the absence of power to the electromagnetic drive, if there is no such extreme change in flight conditions that unacceptably change the smaller clearance determined by the computer, in this case, to increase / decrease this clearance, the heaters are turned on or the spool is shifted to a new maintained position, ensuring equality with a tolerance of +0.1 mm of a smaller gap to the minimum acceptable and possible for I of this engine, and on throttled engine modes, the spool completely blocks the holes supplying cooling air from the second circuit to the internal cavities of the ring inserts, the heaters are turned on and the computer at each step of changing the heating intensity in the same way as in forced modes, selects a smaller gap, controls the on / off of the heaters and the change in the intensity of heating, until a smaller gap is established in the interval δ _min ≤δ≤δ _min + 0.1 ÷ 0.2 mm, the heaters are turned off to a complete stop the engine at a run-out speed that excludes the possibility of cutting the rotor blades, and with repeated throttle response, at its beginning, the computer compares the smaller gap between the annular insert and the rotor blades of the second stage of the two-stage turbine, which is selected in a known manner, with the value of this radial clearance δ _c , possibly smaller, but greater by 0.1 ÷ 0.2 mm of the total exhaust of the working blades and the locking part of the disk under the action of centrifugal forces of the wheel of the second stage of the two-stage turbine, and if the smaller clearance is greater than this value, then the radial clearance is controlled by changing the cooling intensity, as well as in forced engine modes, and if it is less, then first, the heaters are turned on before restarting the engine, and when the smaller clearance reaches a value that is 0.1 ÷ 0.2 mm larger than δ _c , further the radial clearance is controlled by changing the cooling intensity, as well as in forced engine modes, and the excited vibrations of the radial clearance control system in all engine operation modes are effective Actively damped by spring dampers. 8. A two-stage high pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to claim 7, characterized in that the cooling air supply system to the disk blade and the rotor blades of the first stage of the turbine engine contains a pipeline with a shutter connecting the cavity after the last stage HPC with a cavity in front of the twisting device of the first stage, moreover, control of the supply of cooling air due to the last stage of the HPC for cooling the disk and rotor blades of the first The penalties of the rotor of a two-stage theater of operations is carried out actively. 9. Two-stage high pressure turbine of a dual-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to any one of paragraphs. 7 and 8, characterized in that each spring is assembled from straight tapes and compressed by a uniformly distributed load - air pressure in the second circuit of the engine, or each spring is assembled in the following layout: its package is assembled from tapes of the same thickness, two, three are installed in its center and smoother ribbons, on both sides of them are packages assembled “corrugations into corrugations” of two or more corrugated ribbons in such a way that the vertices of the corrugations of one packet rest on the packet of smooth ribbons in the same sections as the vertices of the second packet, and corrugation pitch in corrugated tapes is chosen such that in the span of the packet there is only one vertex resting on the packet of smooth ribbons in the middle of the span, and under each spring support there is only one vertex of the corrugation, and packets collected from one, two or more are installed on the packets of corrugated tapes smooth ribbons, and in the assembled spring the corrugations of the corrugated ribbons are completely straightened, or one smooth ribbon with a thickness of one is installed on packages of smooth ribbons or directly on packages of corrugated ribbons

where k = 5 ÷ 10, and the step of the corrugated tapes of the packet is chosen such that in each span there are one, two or more vertices of the corrugations, and in the collected spring the corrugations of the corrugated tapes are completely straightened. 10. A two-stage high pressure turbine of a dual-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to claim 9, characterized in that a layer of abradable material is applied to the inner working surface of the annular insert of the first stage, or a layer of abradable material is deposited on the inner working surface of the annular inserts both stages of the theater, and in this case, the sensors measuring the radial clearance are offset in the sockets of the ring insert of the second stage in radial directions by a thickness of iraemogo layer. 11. A two-stage high-pressure turbine of a double-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to claim 10, characterized in that the hydraulic diameters of the duct supplying the cooling air of the second circuit inside the annular insert of the first stage and exhausting it are larger than the hydraulic diameter of a similar duct second stage, or vice versa. 12. Two-stage high pressure turbine of a dual-circuit gas turbine engine with active thermal regulation of the radial clearance in the turbine according to any one of paragraphs. 7, 8, 9, 10 and 11, characterized in that the electrical parameters of the half rings with the conductors of the resistive heater or the coils of the induction heater are different for the first and second stages of the turbine and the heater of each stage of the theater operates according to its own program. 13. The method of thermal active control of the radial clearances of a single-stage high pressure turbine comprises radial clearances control operations performed according to the commands of the engine electronic system based on the signals of the rotational speed sensor of the VD turbocompressor and the barostat, and the disk and rotor blades of the turbine engine rotor are cooled by air due to the last stage of the HPC passively - without commands from these sensors, or actively - the cooling of these parts is turned off during take-off mode and transitional modes from take-off mode to cruising th mode, characterized in that all radial clearances theater management operations are performed according to the flight cycle on commands board according to the program computer, developed on experimental basis for the entire flight cycle, providing good efficiencies, C _p, and the characteristics of its traction and safety operation of the engine, excluding insertion of rotor blades and failure of takeoff thrust, which is performed in the following sequence: before starting the engine, a heater is turned on, which is turned off after a certain interval a terrible time, the engine starts and in forced modes the cooling of the annular insert is turned on by the secondary air supplied to its internal cavity with a predetermined law of change in speed and time, with an intensity that reaches its maximum value in take-off mode, and passive cooling of the disk and rotor blades of the turbine engine performed at all engine operating modes and is not regulated by the radial clearance control system, and in the mode transitioning from take-off mode to cruising mode, the ring the insert is cooled by the air of the second circuit with an intensity that gradually decreases according to the law specified by the computer program to the level required at cruising mode at heights exceeding the height of the barostat limit N, and with the indicated active cooling of the disk and rotor blades of the theater rotor, the predetermined law of cooling the annular insert by air is fulfilled the second circuit, also decreasing to the same level in cruise mode, but with large gradients of intensity reduction in engine speed, and in throttling In the flight cycle, the cooling of the annular insert is turned off by the air of the second circuit and its heating is turned on by an heater with an intensity that gradually decreases according to a given law with a decrease in engine speed, and the heater is turned off and the engine is stopped, and the excited vibrations of the annular insert and parts attached to it are applied at all engine operating modes effectively damped by spring dampers. 14. The method of thermal active control of the radial clearances of a single-stage high pressure turbine comprises adjusting the radial clearance by heating the annular insert with a heater in idle and transient modes with revolutions greater than the cruise speed, characterized in that all the operations for controlling the radial clearances of the theater of operations are performed according to the flight cycle on board computer commands according to a program developed on the basis of an experiment for the entire flight cycle, providing good values efficiency, С _p and the characteristics of its thrust and engine operation safety, excluding the insertion of rotor blades and failure of the thrust during take-off, which is performed in the following sequence: before starting the engine, the heater turns on, which turns off after a certain period of time, the engine starts, if installed on the outlet of the cooled air supplied from the second circuit to the inner cavity of the annular insert, a normally open electro-pneumatic valve, this air is supplied to the inner cavity of the annular sun avki in transient forced modes, in idle mode as a part of engine boost modes, take-off mode, transient modes from take-off to cruising and in cruising mode, then the electropneumatic valve closes and the supply of cooling air from the secondary circuit stops and the heater heating the ring insert to throttling transition modes, the heater is turned off until the engine stops at coasting speed, eliminating the possibility of cutting the blades, and if not installed the outlet of the cooled air supplied from the second circuit to the inner cavity of the annular insert, a normally closed electro-pneumatic valve, after the heater is turned off when the engine is started, the electro-pneumatic valve opens and keeps open in all transient forced modes, in the idle mode as part of the engine boost modes, take-off mode, transient modes from take-off to cruising, then the electro-pneumatic valve deenergizes and closes and the cooling air of the second circuit does not enter the internal the cavity until the engine is completely stopped, and in cruise mode the ring insert is cooled passively only by the air coming from the last stage of the HPC, the heater is turned on, heating the ring insert on the throttle transition modes, the heater is turned off until the engine stops completely at coasting speed, eliminating the possibility of the insertion of the working blades, and the excited vibrations of the annular insert and the parts attached to it at all engine operating modes, are effectively damped by spring dampers. 15. The method of thermal active control of the radial gaps of a high-pressure turbine comprises adjusting the gap by the commands of the on-board computer receiving signals from sensors measuring the radial clearance, characterized in that the computer gives commands by signals from sensors measuring the radial clearance, fixed in the ring insert of a single-stage turbine engine or in the ring insert of the second stage of the two-stage theater of operations, before starting the engine, the computer determines the smallest of the gaps measured by the sensors, includes heat registers and annular insert before starting the engine are heated by heaters to obtain the value of the radial clearance δ _i, the least possible, but a larger total quantity blades drawing and the hinge portion of the disc the high pressure turbine by the centrifugal force of the turbine stage, in which this sum is greater, the motor starts up, and if the measured radial clearances increase during forced flight cycle modes, the heater turns off, the computer at each step of changing the cooling intensity The main insertion of the recorded signals from each sensor forms its own group of n = 1, 2, 3, ... periods of change in the size of the gaps, and in each period calculates the difference between the largest and smallest values of these gaps, of all groups for the turbine stage sensors separately selects the largest largest difference, determines the corresponding smaller gap and compares it to the minimum allowable value and possible for a given engine size of this gap δ _min, if this smaller gap greater than the minimum permissible value of Azora δ _min, then the computer heater, if it was on command, turns off, and the regulating member of the cooling air flow supplied from the second circuit in the internal cavity of the annular inserts - the control piston is displaced and opens a certain area of the holes in each regulation step, through which this cooling air is discharged, in transient forced modes and in stationary mode - low gas, these operations are continuously repeated until the smaller clearance defined in this operation at the second stage of the turbine reaches the minimum permissible and possible value for a given engine, the maximum intensity of cooling of the annular inserts is achieved in the take-off mode when these holes are fully open, and thrust failure in the take-off mode of the engine is eliminated, moreover, the disk and rotor blades of the stage of a one-stage theater stage or the first stage of a rotor of a two-stage theater are cooled air due to the last stage of the HPC occurs passively at all engine operating modes - without computer control commands, or at take-off mode and not ehodnyh modes from takeoff to cruise mode is active - cooling off these parts, and at cruising altitudes at greater height limit H barostat, spool position required to maintain the range of δ _min ≤δ≤δ _min +0,1 mm lower from the measured values of the radial clearance, remains until the end of the cruise mode in the absence of power to the electromagnetic drive, if there is no such extreme change in flight conditions that unacceptably change the smaller clearance determined by by a computer, in this case, when the clearance is increased / decreased, the heaters are turned on or the slide valve is shifted to a new maintained position, ensuring equality with a tolerance of +0.1 mm of the smaller clearance to the minimum allowable and possible value for a given engine, and on throttled engine modes, the slide valve completely covers holes supplying cooling air from the second circuit to the internal cavities of the ring inserts, heaters and a computer are turned on at each step of changing the intensity of the heat in the same way as in forced modes, selects a smaller gap, controls the on / off of the heaters and changes in the intensity of heating until a smaller gap is set in the interval δ _min ≤δ≤δ _min + 0.1 ÷ 0.2 mm, heaters they are switched off until the engine completely stops at coasting speed, which excludes the possibility of cutting the rotor blades, and with repeated throttle response, at its beginning, the computer compares the smaller clearance chosen in a known manner between the annular insert and the rotor blades of a single-stage turbine or between at an annular insert and rotor blades of the second stage of a two-stage turbine with the value of this radial clearance δ _c , possibly smaller, but 0.1–0.2 mm larger than the total exhaust of the rotor blades and the locking part of the disk under the action of centrifugal forces of a wheel of a single-stage turbine or a second wheel stages of a two-stage turbine, and if the smaller gap is greater than this value, then the radial clearance is controlled by changing the cooling intensity, as well as in forced engine modes, and if less, then first restarting the engine unit includes a heater, and when the lower clearance value to 0.1 ÷ 0.2 mm greater value δ _i, the further control of radial clearance by changing the cooling rate, as well as for the forced motor modes excited oscillation and the radial control system Clearances at all engine operating modes are effectively damped by spring dampers.

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