The invention relates to low-pressure turbines of gas turbine engines for aviation applications. The low-pressure turbine of a gas turbine engine includes a rotor, a stator with a rear support, a labyrinth seal with internal and external flanges on the rear support of the stator. The labyrinth seal of the turbine is made in two levels. The inner tier is formed by two labyrinth sealing combs directed towards the turbine axis, and the working surface of the labyrinth seal inner flange directed towards the turbine flow path. The outer tier is formed by the sealing combs of the labyrinth directed towards the flow path of the turbine, and the working surface of the outer flange of the labyrinth seal directed towards the axis of the turbine. The sealing combs of the labyrinth of the inner tier of the labyrinth seal are made with parallel inner walls, between which a damping ring is installed. The outer flange of the labyrinth seal is made with an outer closed annular air cavity. Between the flow path of the turbine and the outer flange of the labyrinth seal there is an annular barrier wall mounted on the rear support of the stator. The working surface of the inner flange of the labyrinth seal is located in such a way that the ratio of the inner diameter at the outlet of the flow path of the turbine to the diameter of the working surface of the inner flange of the labyrinth seal is 1.05 1.5. The invention improves the reliability of the low-pressure turbine of a gas turbine engine. 3 ill. 
Drawings to the RF patent 2507401
The invention relates to low-pressure turbines of gas turbine engines for aviation applications.
A low-pressure turbine of a gas turbine engine with a rear support is known, in which the labyrinth seal separating the rear discharge cavity of the turbine from the flow path at the outlet of the turbine is made in the form of a single tier. (S.A. Vyunov, "Design and design of aircraft gas turbine engines", Moscow, "Engineering", 1981, p. 209).
The disadvantage of the known design is the low stability of the pressure in the unloading cavity of the turbine due to the unstable value of the radial gaps in the labyrinth seal, especially at variable engine operating modes.
Closest to the claimed design is a low-pressure turbine of a gas turbine engine, including a rotor, a stator with a rear support, a labyrinth seal with inner and outer labyrinth flanges mounted on the rear support of the stator (US patent No. 7905083, F02K 3/02, 03/15/2011).
The disadvantage of the known design, taken as a prototype, is the increased value of the axial force of the turbine rotor, which reduces the reliability of the turbine and the engine as a whole due to the low reliability of the angular contact bearing, which perceives the increased axial force of the turbine rotor.
The technical result of the claimed invention is to increase the reliability of the low-pressure turbine of a gas turbine engine by reducing the magnitude of the axial force of the turbine rotor and ensuring the stability of the axial force when operating in transient conditions.
The specified technical result is achieved by the fact that in a low-pressure turbine of a gas turbine engine, including a rotor, a stator with a rear support, a labyrinth seal made with inner and outer flanges mounted on the rear support of the stator, the labyrinth seal of the turbine is made in two tiers, while the inner tier of the labyrinth seal formed by two sealing combs of the labyrinth directed to the axis of the turbine, and the working surface of the inner flange of the labyrinth seal directed to the flow path of the turbine, and the outer tier of the labyrinth seal is formed by sealing combs of the labyrinth directed to the flow path of the turbine, and the working surface of the outer flange of the labyrinth seal directed to the axis of the turbine, and the sealing combs of the labyrinth of the inner tier of the labyrinth seal are made with parallel inner walls, between which a damping ring is installed, and the outer flange of the labyrinth seal is made with an external closed annular air cavity, while between the flow path of the turbine and the outer flange of the labyrinth seal there is an annular barrier wall mounted on the rear support of the stator, and the working surface of the inner flange of the labyrinth seal is located in such a way that the following condition is met:
where D is the inner diameter at the outlet of the flow path of the turbine,
The labyrinth seal at the outlet of the low-pressure turbine is two-tier, arranging the seal tiers in such a way that the inner tier is formed by two labyrinth sealing scallops directed towards the turbine axis and the working surface of the labyrinth seal inner flange directed towards the flow path of the turbine, and the outer tier is formed directed to the flow path turbine sealing combs of the labyrinth and directed to the axis of the turbine working surfaces of the outer flange of the labyrinth seal, allows you to ensure reliable operation of the labyrinth seal in transient modes of operation of the turbine, which ensures the stability of the axial force acting on the turbine rotor, and increases its reliability.
The implementation of the sealing scallops of the labyrinth of the inner seal tier with parallel inner walls, between which a damping ring is installed, reduces vibration stresses in the labyrinth and reduces the radial gaps between the scallops of the labyrinth and the flanges of the labyrinth seal.
The execution of the outer flange of the labyrinth seal with an external closed air cavity, as well as the placement of an annular barrier wall installed on the rear stator support between the flow path of the turbine and the outer flange of the labyrinth seal, can significantly reduce the rate of heating and cooling of the outer flange of the labyrinth seal in transient modes, bringing it closer thus to the rate of heating and cooling of the outer tier of the labyrinth seal, which ensures the stability of the radial clearances between the stator and the rotor in the seal and increases the reliability of the low-pressure turbine by maintaining a stable pressure in the unloading after-turbine cavity.
The choice of the ratio D/d=1.05 1.5 is due to the fact that at D/d<1,05 снижается надежность работы лабиринтного уплотнения из-за воздействия на уплотнение высокотемпературного газа, выходящего из турбины низкого давления.
When D/d>1.5 reduces the reliability of the gas turbine engine by reducing the axial unloading force acting on the rotor of the low-pressure turbine.
Figure 1 shows a longitudinal section of a low-pressure turbine of a gas turbine engine.
Figure 2 - element I in figure 1 in an enlarged view.
Figure 3 - element II in figure 2 in an enlarged view.
The low-pressure turbine 1 of the gas turbine engine consists of a rotor 2 and a stator 3 with a rear support 4. To reduce the axial forces from gas forces acting on the rotor 2 at its outlet, an unloading cavity 6 of increased pressure, which is inflated with air due to the intermediate stage of the compressor (not shown) and is separated from the flow path 7 of the turbine 1 by a two-tier labyrinth seal, and the labyrinth 8 of the seal is fixed by a threaded connection 9 on the disk of the last stage 5 of the rotor 2, and the inner flange 10 and the outer flange 11 of the labyrinth seal are fixed on the rear support 4 of the stator 3. The inner tier of the labyrinth seal is formed by the working surface 12 of the inner flange 10, directed (facing) towards the flow path 7 of the turbine 1, and two sealing combs 13, 14 of the labyrinth 8 directed towards the axis 15 of the turbine 1. The inner walls 16,17 respectively of the scallops 13, 14 are made parallel to each other. A damping ring 18 is installed between the inner walls 16 and 17, which helps to reduce vibration stresses in the labyrinth 8 and reduce the radial gaps 19 and 20, respectively, between the labyrinth 8 of the rotor 2 and the flanges 10, 11. The outer tier of the labyrinth seal is formed by the working surface 21 of the outer flange 11, directed (facing) towards the axis 15 of the turbine 1, and the sealing scallops 22 of the labyrinth 8 directed towards the flow path 7 of the turbine 1. The outer flange 11 of the labyrinth seal is made with an outer closed annular air cavity 23 bounded from the outside by the wall 24 of the outer flange 11. Between the wall 24 of the outer flange 11 of the labyrinth seal and the flow path 7 of the turbine 1 there is an annular barrier wall 25 mounted on the rear support 4 of the stator 3 and protecting the outer flange 11 from the high-temperature gas flow 26 flowing in the flow path 7 of the turbine 1.
The working surface 12 of the inner flange 10 of the labyrinth seal is located in such a way that the condition is met:
where D is the inner diameter of the flow part 7 of the turbine 1 (at the outlet of the flow part 7);
d is the diameter of the working surface 12 of the inner flange 10 of the labyrinth seal.
The device works as follows.
During operation of the low-pressure turbine 1, the temperature state of the outer flange 11 of the labyrinth seal can be affected by a change in the temperature of the gas flow 26 in the flow path 7 of the turbine 1, which could significantly change the radial clearance 19 and the axial force acting on the rotor 2 due to a change in air pressure in the unloading cavity 6. However, this does not happen, since the inner flange 10 of the inner tier of the labyrinth seal is inaccessible to the influence of the gas flow 26, which contributes to the stability of the radial clearance 20 between the inner flange 10 and the labyrinth combs 13, 14, as well as the stability of the pressure in the cavity 6 and the stability of the axial force acting on rotor 2 of turbine 1.
CLAIM
A low-pressure turbine of a gas turbine engine, including a rotor, a stator with a rear support, a labyrinth seal with inner and outer flanges mounted on the rear support of the stator, characterized in that the labyrinth seal of the turbine is made in two tiers, while the inner tier of the labyrinth seal is formed by two labyrinth seal combs, directed to the axis of the turbine, and the working surface of the inner flange of the labyrinth seal directed to the flow path of the turbine, and the outer tier of the labyrinth seal is formed by the sealing combs of the labyrinth directed to the flow path of the turbine, and the working surface of the outer flange of the labyrinth seal directed to the axis of the turbine, and the sealing the scallops of the labyrinth of the inner tier of the labyrinth seal are made with parallel inner walls, between which a damping ring is installed, and the outer flange of the labyrinth seal is made with an outer closed annular air cavity, while between the flow path of the turbine and the outer flange of the labyrinth seal there is an annular barrier wall installed on the rear stator support, and the working surface of the inner flange of the labyrinth seal is located in such a way that the following condition is met:
D/d=1.05 1.5, where
D is the inner diameter at the outlet of the flow path of the turbine,
d is the diameter of the working surface of the inner flange of the labyrinth seal.
Turbine
The turbine is designed to drive the compressor and auxiliary units of the engine. Engine turbine - axial, jet, two-stage, cooled, two-rotor.
The turbine assembly includes sequentially arranged single-stage high and low pressure axial turbines, as well as a turbine support. Support - an element of the power circuit of the engine.
high pressure turbine
SA HPT consists of an outer ring, an inner ring, a cover, a swirling device, nozzle blade blocks, labyrinth seals, nozzle blade joint seals, spacers with honeycomb inserts and fasteners.
The outer ring has a flange for connection with the flange of the rim of the LPT nozzle apparatus and the VVT body. The ring is telescopically connected to the VVT body and has a cavity for supplying secondary air from the OCS to cool the outer shelves of the nozzle blades.
The inner ring has a flange for connection with the cover and the inner body of the OKS.
SA TVD has forty-five blades, combined into fifteen cast three-blade blocks. The block design of the SA blades makes it possible to reduce the number of joints and gas overflows.
Nozzle blade - hollow, cooled two-cavity. Each blade has a vane, outer and inner flanges, which together with the vane and flanges of adjacent blades form the flow path of the HPT SA.
The TVD rotor is designed to convert the energy of the gas flow into mechanical work on the rotor shaft. The rotor consists of a disk, pins with labyrinth and oil sealing rings. The disk has ninety-three slots for fastening the HPT rotor blades in “Christmas tree” locks, holes for tight-fitting bolts tightening the disc, pin and HPT shaft, as well as inclined holes for supplying cooling air to the rotor blades.
HPT working blade - cast, hollow, cooled. In the inner cavity of the blade to organize the cooling process, there is a longitudinal partition, turbulent pins and ribs. The shank of the blade has an elongated leg and a herringbone type lock. In the shank there are channels for supplying cooling air to the blade airfoil, and in the trailing edge there is a slot for air outlet.
The trunnion shank contains an oil seal and a race of the radial roller bearing of the rear support of the high-pressure rotor.
Low pressure turbine
SA LPT consists of a rim, blocks of nozzle blades, an inner ring, a diaphragm, and honeycomb inserts.
The rim has a flange for connection with the VVT housing and the outer ring of the HPT, as well as a flange for connection with the turbine support housing.
SA TND has fifty-one blades soldered into twelve four-blade blocks and one three-blade block. Nozzle blade - cast, hollow, cooled. The feather, outer and inner shelves form with the feather and shelves of neighboring blades the flow part of the SA.
A perforated deflector is placed in the inner part of the cavity of the blade airfoil. On the inner surface of the pen there are transverse ribs and turbulence pins.
The diaphragm is designed to separate the cavities between the HPT and LPT impellers.
The LPT rotor consists of a disk with working blades, a trunnion, a shaft and a pressure disk.
The LPT disc has fifty-nine grooves for fastening the working blades and inclined holes for supplying cooling air to them.
The working vane of TND - cast, hollow, cooled. On the peripheral part, the blade has a shroud with a labyrinth seal comb, which seals the radial gap between the stator and the rotor.
From axial movements in the disc, the blades are fixed by a split ring with an insert, which, in turn, is fixed by a pin on the disc rim.
The trunnion has internal splines in the front part for transmitting torque to the LPT shaft. On the outer surface of the front part of the trunnion, there is an inner race of the roller bearing of the rear support of the HPT, a labyrinth and a set of sealing rings, which, together with the cover installed in the trunnion, form the front seal of the oil cavity of the HPT support.
A set of sealing rings is installed on the cylindrical belt in the rear part, which, together with the cover, form a seal for the oil cavity of the LPT support.
The TND shaft consists of three parts. The connection of the shaft parts to each other is forked. The torque at the joints is transmitted by radial pins. At the rear of the shaft there is an oil pump for the turbine support.
In front of the LPT there are splines that transmit torque to the low-pressure compressor rotor through the spring.
The pressure disc is designed to create an additional backwater and provides an increase in the pressure of the cooling air at the inlet to the working blades of the LPT.
The turbine support includes a support housing and a bearing housing. The support housing consists of an outer housing and an inner ring connected by power racks and forming a power circuit for the turbine support. The structure of the support also includes a screen with fairings, a defoaming mesh and fasteners. Inside the racks there are pipelines for supplying and pumping oil, venting oil cavities and draining oil. Air is supplied through the cavities of the racks to cool the LPT and air is removed from the pre-oil cavity of the support. Racks are covered with fairings. An oil sump pump and an oil collector are installed on the bearing housing. An elastic-oil damper is placed between the outer race of the LPT rotor roller bearing and the bearing housing.
A cone fairing is fixed on the turbine support, the profile of which ensures gas entry into the afterburner combustion chamber with minimal losses.
In 2006, the management of the Perm Engine Building Complex and OAO Territorial Generating Company No. 9 (Perm Branch) signed an agreement for the manufacture and supply of a GTES-16PA gas turbine power plant based on a GTE-16PA with a PS-90EU-16A engine.
We asked Daniil SULIMOV, Deputy General Designer-Chief Designer of Aviadvigatel JSC, to tell us about the main differences between the new engine and the existing PS-90AGP-2.
The main difference between the GTE-16PA plant and the existing GTU-16PER is the use of a power turbine with a rotation speed of 3000 rpm (instead of 5300 rpm). Reducing the rotational speed makes it possible to abandon the expensive gearbox and increase the reliability of the gas turbine plant as a whole.
Specifications of GTU-16PER and GTE-16PA engines (under ISO conditions)
Optimization of the main parameters of the power turbine
The basic parameters of a free turbine (ST): diameter, flow path, number of stages, aerodynamic efficiency are optimized to minimize direct operating costs.
Operating costs include the cost of purchasing ST and costs for a certain (acceptable for the customer as a payback period) period of operation. The choice of a payback period that is quite visible for the customer (no more than 3 years) made it possible to implement an economically sound design.
The choice of the optimal variant of a free turbine for a specific application as part of the GTE-16PA was made in the engine system as a whole based on a comparison of direct operating costs for each variant.
Using one-dimensional modeling of the ST, the achievable level of aerodynamic efficiency of the ST was determined by the average diameter for a discretely given number of stages. The optimal flow part for this variant was chosen. The number of blades, taking into account their significant impact on the cost, was chosen from the condition of ensuring the Zweifel aerodynamic load factor equal to one.
Based on the selected flow path, the weight of the SP and the production cost were estimated. The turbine options in the engine system were then compared in terms of direct operating costs.
When choosing the number of stages for ST, the change in efficiency, acquisition and operation costs (fuel cost) are taken into account.
The cost of acquisition increases evenly with the growth of the cost price with an increase in the number of steps. In a similar way, the realized efficiency also grows - as a result of a decrease in the aerodynamic load on the stage. Operating costs (fuel component) fall with increasing efficiency. However, the total costs have a clear minimum at four stages in the power turbine.
The calculations took into account both the experience of our own developments and the experience of other companies (implemented in specific designs), which made it possible to ensure the objectivity of the estimates.
In the final design, by increasing the load per stage and reducing the efficiency of the ST from the maximum achievable value by about 1%, it was possible to reduce the total costs of the customer by almost 20%. This was achieved by reducing the cost and price of the turbine by 26% relative to the variant with maximum efficiency.
Aerodynamic design ST
The high aerodynamic efficiency of the new ST at a sufficiently high load was achieved by using the experience of JSC Aviadvigatel in the development of low-pressure turbines and power turbines, as well as the use of multi-stage spatial aerodynamic models using the Euler equations (without viscosity) and Navier-Stokes (taking into account viscosity ).
Comparison of the parameters of power turbines GTE-16PA and HPP Rolls-Royce
Comparison of the parameters of the ST GTE-16PA and the most modern Rolls-Royce TRD family TRD (Smith diagram) shows that in terms of the angle of rotation of the flow in the blades (approximately 1050), the new ST is at the level of Rolls-Royce turbines. The absence of a strict weight limit inherent in aircraft structures made it possible to somewhat reduce the load factor dH/U2 by increasing the diameter and circumferential speed. The value of the output speed (typical of ground structures) made it possible to reduce the relative axial speed. In general, the potential of the designed ST to realize efficiency is at the level characteristic of the stages of the Trent family.
The peculiarity of the aerodynamics of the designed ST is also to ensure the optimal value of the turbine efficiency at partial power modes, which are typical for operation in the base mode.
While maintaining the rotational speed, a change (decrease) in the load on the ST leads to an increase in the angles of attack (deviation of the direction of the gas flow at the inlet to the blades from the calculated value) at the inlet to the blade rims. Negative angles of attack appear, the most significant in the last stages of the turbine.
The design of ST blade rows with high resistance to changes in angles of attack is ensured by special profiling of the rows with additional verification of the stability of aerodynamic losses (according to 2D/3D Navier-Stokes aerodynamic models) at high inlet flow angles.
As a result, the analytical characteristics of the new ST showed significant resistance to negative angles of attack, as well as the possibility of using the ST to drive generators that produce current at a frequency of 60 Hz (with a rotation speed of 3600 rpm), that is, the possibility of increasing the rotational speed by 20 % without noticeable loss of efficiency. However, in this case, loss of efficiency is practically inevitable at low power modes (leading to an additional increase in negative angles of attack).
ST design features
To reduce the material consumption and weight of the ST, proven aviation approaches to turbine design were used. As a result, the mass of the rotor, despite the increase in the diameter and number of stages, turned out to be equal to the mass of the rotor of the GTU-16PER power turbine. This ensured a significant unification of the transmissions, the oil system, the pressurization system of the supports and the cooling system of the ST were also unified.
The amount and quality of air used to pressurize transmission bearings has been increased, including its cleaning and cooling. The quality of lubrication of transmission bearings has also been improved by using filter elements with a filtration fineness of up to 6 microns.
In order to increase the operational attractiveness of the new GTE, a specially developed control system has been introduced, which allows the customer to use turbo-expander (air and gas) and hydraulic launch types.
The weight and size characteristics of the engine make it possible to use serial designs of the GTES-16P packaged power plant for its placement.
The noise and heat insulating casing (when placed in capital premises) ensures the acoustic characteristics of the GTPP at the level provided for by sanitary standards.
The first engine is currently undergoing a series of special tests. The gas generator of the engine has already passed the first stage of equivalent-cyclic tests and has started the second stage after the revision of the technical condition, which will be completed in the spring of 2007.
The power turbine as part of a full-size engine passed the first special test, during which 7 throttle characteristics and other experimental data were taken.
According to the test results, a conclusion was made about the operability of the ST and its compliance with the declared parameters.
In addition, according to the test results, some adjustments were made to the design of the ST, including a change in the cooling system of the hulls to reduce heat release into the station room and ensure fire safety, as well as to optimize the radial clearances to increase efficiency, adjust the axial force.
The next test of the power turbine is scheduled for summer 2007.
Gas turbine plant GTE-16PA
on the eve of special tests
Today, aviation is almost 100% composed of machines that use a gas turbine type of power plant. In other words, gas turbine engines. However, despite the increasing popularity of air travel now, few people know how that buzzing and whistling container that hangs under the wing of an airliner works.
Principle of operation gas turbine engine.
A gas turbine engine, like a piston engine on any car, refers to internal combustion engines. Both of them convert the chemical energy of the fuel into heat, by burning, and then into useful, mechanical. However, how this happens is somewhat different. In both engines, 4 main processes take place - these are: intake, compression, expansion, exhaust. Those. in any case, air (from the atmosphere) and fuel (from tanks) first enter the engine, then the air is compressed and fuel is injected into it, after which the mixture ignites, due to which it expands significantly, and is eventually released into the atmosphere. Of all these actions, only expansion gives energy, all the rest are necessary to ensure this action.
Now what's the difference. In gas turbine engines, all these processes occur constantly and simultaneously, but in different parts of the engine, and in a piston engine, in one place, but at different times and in turn. In addition, the more compressed the air, the more energy can be obtained during combustion, and today the compression ratio of gas turbine engines has already reached 35-40:1, i.e. in the process of passing through the engine, the air decreases in volume, and accordingly increases its pressure by 35-40 times. For comparison, in piston engines, this figure does not exceed 8-9: 1, in the most modern and advanced models. Accordingly, having equal weight and dimensions, a gas turbine engine is much more powerful, and its efficiency is higher. This is the reason for such a widespread use of gas turbine engines in aviation today.
And now more about the design. The four processes listed above take place in the engine, which is shown in the simplified diagram under the numbers:
- air intake - 1 (air intake)
- compression - 2 (compressor)
- mixing and ignition - 3 (combustion chamber)
- exhaust - 5 (exhaust nozzle)
- The mysterious section at number 4 is called the turbine. This is an integral part of any gas turbine engine, its purpose is to obtain energy from gases that exit the combustion chamber at high speeds, and it is located on the same shaft as the compressor (2), which drives it.
Thus, a closed cycle is obtained. Air enters the engine, is compressed, mixed with fuel, ignited, directed to the turbine blades, which remove up to 80% of the gas power to rotate the compressor, all that is left determines the final engine power, which can be used in many ways.
Depending on the method of further use of this energy, gas turbine engines are divided into:
- turbojet
- turboprop
- turbofan
- turboshaft
The engine shown in the diagram above is turbojet. It can be said to be “clean” gas turbine, because after passing through the turbine, which rotates the compressor, the gases exit the engine through the exhaust nozzle at great speed and thus push the aircraft forward. Such engines are now used mainly in high-speed combat aircraft.
Turboprop engines differ from turbojet engines in that they have an additional turbine section, which is also called a low-pressure turbine, consisting of one or more rows of blades that take the energy left after the compressor turbine from the gases and thus rotate the propeller, which can be located both in front and behind the engine. After the second section of the turbine, the exhaust gases actually exit by gravity, having practically no energy, so just exhaust pipes are used to remove them. Similar engines are used in low-speed, low-altitude aircraft.
Turbofans engines have a similar scheme with turboprops, only the second section of the turbine does not take all the energy from the exhaust gases, so these engines also have an exhaust nozzle. But the main difference is that the low-pressure turbine drives the fan, which is enclosed in a casing. Therefore, such an engine is also called a dual-circuit engine, because the air passes through the internal circuit (the engine itself) and the external one, which is necessary only to direct the air stream that pushes the engine forward. Because they have a rather "chubby" shape. It is these engines that are used on most modern airliners, since they are the most economical at speeds approaching the speed of sound and efficient when flying at altitudes above 7000-8000m and up to 12000-13000m.
Turboshaft the engines are almost identical in design to turboprops, except that the shaft that is connected to the low-pressure turbine comes out of the engine and can power absolutely anything. Such engines are used in helicopters, where two or three engines drive a single main rotor and a compensating tail propeller. Even tanks, the T-80 and the American Abrams, now have similar power plants.
Gas turbine engines are also classified according to other signs:
- by input device type (adjustable, unregulated)
- by compressor type (axial, centrifugal, axial-centrifugal)
- according to the type of air-gas path (straight-through, loop)
- by turbine type (number of stages, number of rotors, etc.)
- by type of jet nozzle (adjustable, unregulated), etc.
Turbojet engine with axial compressor received wide application. With the engine running, the process is continuous. The air passes through the diffuser, slows down and enters the compressor. Then it enters the combustion chamber. Fuel is also supplied to the chamber through the nozzles, the mixture is burned, the combustion products move through the turbine. The products of combustion in the turbine blades expand and cause it to rotate. Further, gases from the turbine with reduced pressure enter the jet nozzle and break out at great speed, creating thrust. The maximum temperature also occurs in the water of the combustion chamber.
The compressor and turbine are located on the same shaft. Cold air is supplied to cool the combustion products. In modern jet engines, the operating temperature can exceed the melting point of rotor blade alloys by about 1000 °C. The cooling system for turbine parts and the choice of heat-resistant and heat-resistant engine parts are one of the main problems in the design of jet engines of all types, including turbojet ones.
A feature of turbojet engines with a centrifugal compressor is the design of the compressors. The principle of operation of such engines is similar to engines with an axial compressor.
Gas turbine engine. Video.
Useful related articles.
TO aircraft engines include all types of heat engines used as propulsion devices for aviation-type aircraft, i.e. devices that use aerodynamic quality to move, maneuver, etc. within the atmosphere (airplanes, helicopters, cruise missiles of classes "B-B", "V-3", "3-V", "3-3", aerospace systems, etc.). This implies a wide variety of used engines - from piston to rocket.
Aircraft engines (Fig. 1) are divided into three broad classes:
- piston (PD);
- air-jet (WFD including GTD);
- missile (RD or RKD).
The last two classes are subject to a more detailed classification, in particular the class WFD.
By principle of air compression WRDs are divided into:
- compressor , i.e., including a compressor for mechanical compression of air;
- compressorless
:
- once-through WFD ( SPVRD) with air compression only from velocity pressure;
- pulsating WFD ( PUVRD) with additional air compression in special intermittent gas-dynamic devices.
Rocket engine class LRE also refers to the compressor type of heat engines, since in these engines the working fluid (fuel) is compressed in a liquid state in turbopump units.
Solid propellant rocket engine (RDTT) does not have a special device for compressing the working fluid. It is carried out at the beginning of fuel combustion in the semi-enclosed space of the combustion chamber, where the fuel charge is located.
By operating principle there is a division: PD And PUVRD work in a cycle periodical actions, while WFD, GTD And RKD cycle is carried out continuous actions. This gives them advantages in terms of relative power, thrust, weight, etc., which determined, in particular, the expediency of their use in aviation.
By principle of jet thrust WRDs are divided into:
- direct reaction engines;
- indirect reaction engines.
Engines of the first type create tractive force (thrust P) directly - that's all rocket engines (RKD), turbojet without afterburner and with afterburner chambers ( TRD And TRDF), turbojet bypass (turbofan And TRDDF), once-through supersonic and hypersonic ( SPVRD And scramjet), pulsating (PUVRD) and numerous combined engines.
Indirect reaction gas turbine engines (GTD) transfer the power generated by them to a special propeller (propeller, propfan, helicopter main rotor, etc.), which creates tractive effort using the same air-jet principle ( turboprop , turbopropfan , turboshaft engines - TVD, TVVD, TVGTD). In this sense, the class WFD combines all engines that create thrust according to the air-jet principle.
Based on the considered types of engines of simple circuits, a number of combined engines , connecting the features and advantages of engines of various types, for example, classes:
- turbo-jet engines - TRDP (TRD or turbofan + SPVRD);
- rocket-ramjet - RPD (LRE or RDTT + SPVRD or scramjet);
- rocket-turbine - RTD (TRD + LRE);
and many other combinations of engines of more complex schemes.
Piston engines (PD)
Two-row radial 14-cylinder air-cooled piston engine. General form.
piston engine (English) piston engine ) -
Classification of piston engines. Aircraft piston engines can be classified according to various criteria:
- Depending on the type of fuel used- for light or heavy fuel engines.
- According to the method of mixing- on engines with external mixture formation (carburetor) and engines with internal mixture formation (direct fuel injection into cylinders).
- Depending on the method of ignition of the mixture- for positive ignition and compression ignition engines.
- Depending on the number of strokes- for two-stroke and four-stroke engines.
- Depending on the cooling method- for liquid and air-cooled engines.
- By number of cylinders- for four-cylinder, five-cylinder, twelve-cylinder engines, etc.
- Depending on the location of the cylinders- in-line (with cylinders arranged in a row) and star-shaped (with cylinders arranged in a circle).
In-line engines, in turn, are divided into single-row, two-row V-shaped, three-row W-shaped, four-row H-shaped or X-shaped engines. Axial engines are also divided into single-row, double-row and multi-row.
- By the nature of the change in power depending on the change in altitude- for high-altitude, i.e. engines that retain power as the aircraft rises to altitude, and low-altitude engines whose power decreases with increasing flight altitude.
- Propeller drive method- for motors with direct transmission to the propeller and gear motors.
Modern aircraft piston engines are four-stroke radial engines that run on gasoline. The cylinders of reciprocating engines are usually cooled by air. Previously, piston engines with water-cooled cylinders were also used in aviation.
The combustion of fuel in a piston engine is carried out in cylinders, while thermal energy is converted into mechanical energy, since under the pressure of the resulting gases, the piston moves forward. The translational movement of the piston, in turn, is converted into rotational movement of the engine crankshaft through the connecting rod, which is the connecting link between the cylinder with the piston and the crankshaft.
Gas turbine engines (GTE)
Gas turbine engine - a heat engine designed to convert the energy of fuel combustion into the kinetic energy of a jet stream and (or) into mechanical work on the engine shaft, the main elements of which are a compressor, a combustion chamber and a gas turbine.
Single-shaft and multi-shaft engines
The simplest gas turbine engine has only one turbine, which drives the compressor and at the same time is a source of useful power. This imposes a restriction on the operating modes of the engine.
Sometimes the engine is multi-shaft. In this case, there are several turbines in series, each of which drives its own shaft. The high-pressure turbine (the first after the combustion chamber) always drives the engine compressor, and the subsequent ones can drive both an external load (helicopter or ship propellers, powerful electric generators, etc.) and additional compressors of the engine itself, located in front of the main one.
The advantage of a multi-shaft engine is that each turbine operates at optimum speed and load. With a load driven from the shaft of a single-shaft engine, the throttle response of the engine, that is, the ability to quickly spin up, would be very poor, since the turbine needs to supply power both to provide the engine with a large amount of air (power is limited by the amount of air) and to accelerate the load. With a two-shaft scheme, a light high-pressure rotor quickly enters the regime, providing the engine with air, and the low-pressure turbine with a large amount of gases for acceleration. It is also possible to use a less powerful starter for acceleration when starting only the high pressure rotor.
Turbojet engine (TRD)
Turbojet engine (English) turbojet engine ) - a heat engine that uses a gas turbine, and jet thrust is formed when combustion products flow out of a jet nozzle. Part of the work of the turbine is spent on compressing and heating the air (in the compressor).
Scheme of a turbojet engine:
1. input device;
2. axial compressor;
3. combustion chamber;
4. turbine blades;
5. nozzle.
In a turbojet engine, the compression of the working fluid at the inlet to the combustion chamber and the high value of air flow through the engine are achieved due to the combined action of the oncoming air flow and the compressor located in the TRD tract immediately after the inlet device, in front of the combustion chamber. The compressor is driven by a turbine mounted on the same shaft with it, and running on the same working fluid, heated in the combustion chamber, from which a jet stream is formed. In the inlet device, the static air pressure increases due to the deceleration of the air flow. In the compressor, the total air pressure increases due to the mechanical work performed by the compressor.
Pressure ratio in the compressor is one of the most important parameters of the turbojet engine, since the effective efficiency of the engine depends on it. If for the first samples of turbojet engines this indicator was 3, then for modern ones it reaches 40. To increase the gas-dynamic stability of compressors, they are made in two stages. Each of the cascades operates at its own speed and is driven by its own turbine. In this case, the shaft of the 1st stage of the compressor (low pressure), rotated by the last (lowest speed) turbine, passes inside the hollow shaft of the compressor of the second stage (high pressure). Engine stages are also called low and high pressure rotors.
The combustion chamber of most turbojet engines has an annular shape and the turbine-compressor shaft passes inside the chamber ring. Upon entering the combustion chamber, the air is divided into 3 streams:
- primary air- enters through the front openings in the combustion chamber, slows down in front of the injectors and takes a direct part in the formation of the fuel-air mixture. Directly involved in the combustion of fuel. The fuel-air mixture in the fuel combustion zone in the WFD is close to stoichiometric in composition.
- secondary air- enters through the side openings in the middle part of the combustion chamber walls and serves to cool them by creating an air flow with a much lower temperature than in the combustion zone.
- tertiary air- enters through special air channels in the outlet part of the combustion chamber walls and serves to equalize the temperature field of the working fluid in front of the turbine.
The gas-air mixture expands and part of its energy is converted in the turbine through the rotor blades into the mechanical energy of the rotation of the main shaft. This energy is spent primarily on the operation of the compressor, and is also used to drive engine units (fuel booster pumps, oil pumps, etc.) and drive electric generators that provide energy to various on-board systems.
The main part of the energy of the expanding gas-air mixture is used to accelerate the gas flow in the nozzle, which flows out of it, creating jet thrust.
The higher the combustion temperature, the higher the efficiency of the engine. To prevent the destruction of engine parts, heat-resistant alloys equipped with cooling systems and thermal barrier coatings are used.
Turbojet engine with afterburner (TRDF)
Turbojet engine with afterburner - modification of the turbojet engine, used mainly on supersonic aircraft. It differs from the turbojet engine by the presence of an afterburner between the turbine and the jet nozzle. An additional amount of fuel is supplied to this chamber through special nozzles, which is burned. The combustion process is organized and stabilized with the help of a front-end device that provides mixing of the evaporated fuel and the main flow. The increase in temperature associated with the heat input in the afterburner increases the available energy of the combustion products and, consequently, the speed of the exhaust from the jet nozzle. Accordingly, jet thrust (afterburner) also increases up to 50%, but fuel consumption increases sharply. Afterburner engines are generally not used in commercial aviation due to their low fuel economy.
Double-circuit turbojet engine (TRDD)
The first to propose the concept of a turbofan engine in the domestic aircraft engine industry was A. M. Lyulka (Based on research conducted since 1937, A. M. Lyulka submitted an application for the invention of a bypass turbojet engine. The copyright certificate was awarded on April 22, 1941.)
It can be said that from the 1960s to this day, in the aircraft engine industry, the era of turbofan engines. Turbofan engines of various types are the most common class of turbofan engines used on aircraft, from high-speed fighter-interceptors with low bypass turbofans to giant commercial and military transport aircraft with high bypass turbofans.
Scheme of a turbojet bypass engine:
1. low pressure compressor;
2. inner contour;
3. the output stream of the internal circuit;
4. output stream of the outer circuit.
The basis bypass turbojet engines the principle of attaching an additional mass of air to the turbojet engine passing through the external circuit of the engine was established, which makes it possible to obtain engines with a higher flight efficiency compared to conventional turbojet engines.
After passing through the inlet, the air enters the low pressure compressor, called the fan. After the fan, the air is divided into 2 streams. Part of the air enters the outer circuit and, bypassing the combustion chamber, forms a jet stream in the nozzle. The other part of the air passes through an internal circuit completely identical to the turbofan engine mentioned above, with the difference that the last stages of the turbine in the turbofan engine are the fan drive.
One of the most important parameters of a turbofan engine is the bypass ratio (m), that is, the ratio of air flow through the external circuit to the air flow through the internal circuit. (m \u003d G 2 / G 1, where G 1 and G 2 are the air flow through the internal and external circuits, respectively.)
When the bypass ratio is less than 4 (m<4) потоки контуров на выходе, как правило, смешиваются и выбрасываются через общее сопло, если m>4 - streams are ejected separately, since mixing is difficult due to a significant difference in pressures and velocities.
The turbofan engine is based on the principle of increasing the flight efficiency of the engine, by reducing the difference between the speed of the expiration of the working fluid from the nozzle and the flight speed. The reduction in thrust, which will cause a decrease in this difference between speeds, is compensated by an increase in air flow through the engine. The consequence of an increase in air flow through the engine is an increase in the area of the front section of the engine inlet, which results in an increase in the diameter of the engine inlet, which leads to an increase in its drag and mass. In other words, the higher the bypass ratio, the larger the diameter of the engine, all other things being equal.
All turbofan engines can be divided into 2 groups:
- with mixing flows behind the turbine;
- without mixing.
In a turbofan engine with a mixture of flows ( TRDDsm) air flows from the external and internal circuits enter a single mixing chamber. In the mixing chamber, these flows are mixed and leave the engine through a single nozzle with a single temperature. TRDSM are more efficient, however, the presence of a mixing chamber leads to an increase in the dimensions and weight of the engine
Turbofan engines, like turbofan engines, can be equipped with adjustable nozzles and afterburners. As a rule, these are turbofan engines with low bypass ratios for supersonic military aircraft.
Military turbofan EJ200 (m=0.4)
Bypass turbojet engine with afterburner (TRDDF)
Dual-circuit turbojet engine with afterburner - modification of the turbofan engine. Differs in the presence of an afterburner chamber. Has found wide application.
The combustion products leaving the turbine are mixed with the air coming from the external circuit, and then heat is supplied to the general flow in the afterburner, which operates on the same principle as in TRDF. The products of combustion in this engine flow from one common jet nozzle. Such an engine is called dual-circuit engine with a common afterburner.
TRDDF with deflectable thrust vector (OVT).
Thrust vector control (VCT) / Thrust vector deviation (VVT)
Special rotary nozzles, on some turbofan engines (F), allow you to deflect the flow of the working fluid flowing from the nozzle relative to the engine axis. OVT leads to additional losses of engine thrust due to the additional work on turning the flow and complicates the control of the aircraft. But these shortcomings are fully compensated by a significant increase in maneuverability and a reduction in the aircraft takeoff run and landing run, up to and including vertical takeoff and landing. OVT is used exclusively in military aviation.
High Bypass Turbofan / Turbofan Engine
Scheme of a turbofan engine:
1. fan;
2. protective fairing;
3. turbocharger;
4. the output stream of the internal circuit;
5. output stream of the outer circuit.
turbofan engine (English) turbofan engine ) is a turbofan engine with a high bypass ratio (m>2). Here, the low-pressure compressor is converted into a fan, which differs from the compressor in a smaller number of steps and a larger diameter, and the hot jet practically does not mix with the cold one.
This type of engine uses a single-stage, large-diameter fan that provides high airflow through the engine at all flight speeds, including low takeoff and landing speeds. Due to the large diameter of the fan, the nozzle of the outer contour of such turbofan engines becomes quite heavy and is often shortened, with straighteners (fixed blades that turn the air flow in the axial direction). Accordingly, most turbofan engines with a high bypass ratio - no mixing.
Device inner contour such engines are similar to the turbojet engine, the last stages of the turbine of which are the fan drive.
Outer loop Such a turbofan engine, as a rule, is a single-stage large-diameter fan, behind which there is a directing vane made of fixed blades, which accelerate the air flow behind the fan and turn it, leading to an axial direction, the outer contour ends with a nozzle.
Due to the fact that the fan of such engines, as a rule, has a large diameter, and the degree of air pressure increase in the fan is not high, the nozzle of the external circuit of such engines is quite short. The distance from the engine inlet to the outer contour nozzle exit can be much less than the distance from the engine inlet to the inner contour nozzle exit. For this reason, quite often the nozzle of the outer contour is mistaken for a fan fairing.
Turbofan engines with a high bypass ratio have a two- or three-shaft design.
Advantages and disadvantages.
The main advantage of such engines is their high efficiency.
Disadvantages - large weight and dimensions. Especially - the large diameter of the fan, which leads to significant air resistance in flight.
The scope of such engines is long- and medium-haul commercial airliners, military transport aviation.
Turbopropfan engine (TVVD)
Turbopropfan engine (English) turbo propfan engine ) -
