Low pressure turbine of a gas turbine engine. Isentropic flow velocity in relative motion

The invention relates to turbines low pressure gas turbine engines for aviation applications. Low pressure turbine gas turbine engine includes rotor, stator with back support, labyrinth seal with inner and outer flanges on back support 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 working surface the inner flange of the labyrinth seal was 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, adopted 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 low reliability. angular contact bearing, perceiving 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 working surfaces of the outer flange of the labyrinth directed towards the axis of the turbine, allows to provide reliable performance labyrinth seal during transient 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 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 the gas forces acting on the rotor 2 at its outlet, between the disk last step 5 of the rotor 2 and the rear support 4, an unloading cavity 6 is made high blood 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 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 combs 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 to the flow path 7 of the turbine 1. The outer flange 11 of the labyrinth seal is made with an external closed annular air cavity 23, limited with outside 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.

3. GAS EXPANSION PROCESS IN TURBINE

The chapter deals with next questions:
- Appointment in the TRD;
- scheme and principle of operation axial;
- circumferential force effective work gas, turbine efficiency and power;
- the main parameters that determine the power of the turbine;
— teamwork turbines and compressors in turbojet engines;
- multistage turbines and features of the operation of twin-shaft turbine engines;
- VRD output devices.

The gas, which has significant potential energy, enters the turbine from the combustion chamber.
is a bladed machine that converts the energy of gas compressed and heated in combustion chambers into mechanical work on the shaft. In a turbojet engine, the turbine is used to rotate the compressor rotor and all service units: fuel, oil, hydraulic pumps and etc.
Compared to other engines that convert gas energy into mechanical work, it has a number of advantages:
- the possibility of obtaining high power in one unit with small dimensions and weight;
— high efficiency, which is due to the good aerodynamics of the flow path and the absence of sharp turns flow;
- simplicity and reliability of the design.
Turbines are classified according to the direction of movement of the gas flow, according to the number of stages and other features.
In the direction of gas flow, turbines can be radial , when the flow moves from the center to the periphery along the radius of the turbine elements, and axial, in which the flow moves along the axis of the turbine.
Axial turbines are used in turbojet engines.
— According to the number of turbine stages, turbojet engines are performed as one, two or many stages, depending on the degree of gas expansion in the turbine.
The classification of turbines according to other criteria is considered in the next paragraph.

3.2. SCHEME AND PRINCIPLE OF OPERATION OF THE AXIAL STAGE

The main elements of the turbine stage are the nozzle apparatus (SA) and Working wheel(RK) fig. 26.
Blades SA and RK form a system of channels in the turbine flow path through which the gas flow flows.
To consider the principle of operation of a turbine stage, we cut it with a cylindrical surface a - a and turn it flat. Let's get a planar turbine lattice, consisting of a section SA and RK (Fig. 27).
In cross section, the blades SA and RK are aerodynamic profiles.
Gas from the combustion chamber absolute speed flow C 3 , pressure P 3 and temperature T 3 enters the channels of the nozzle apparatus. The nozzle apparatus is designed to convert the potential energy of gas flow pressure into kinetic energy. For this purpose, the SA channels are made tapering in the flow (f 3 ΄< f 3 , where f is the cross-sectional area of ​​the channel).

The flow rate in SA increases from C 3 to C 3 ", and the pressure and temperature of the gas fall (P 3 "<Р 3 и Т 3 "<Т 3).
With an absolute speed C3 "gas enters the blades of the impeller rotating at a circumferential speed U. In the interblade channel of the RC, the gas moves with a relative speed W 3" equal to the geometric difference of the absolute C 3 "and the circumferential speed U at the entrance to the RC , i.e. W 3 "= C 3" - U.
The speed plan at the entrance to the RC is shown in fig. 27. To ensure a shockless entry, the leading edges of the RC blades are installed in the direction of the relative velocity W 3 ". Due to the increase in circumferential speed from the blade base to the end and the need to ensure a shockless entry at all radii, the RC blade is subjected to "twist".
In the impeller, the kinetic energy of the gas flow is converted into mechanical work. The absolute flow rate decreases in the channels of the Republic of Kazakhstan from C 3 "to C 4 .
Depending on the type of turbine, the gas in the interblade channels of the RC either continues to expand (the pressure drops from P 3 "to P 4), or only changes the direction of movement, and the pressure remains unchanged.
A turbine in which gas expands in the interblade channels of the RC is called reactive. A turbine in which only a turn of the flow in the RC is carried out is called active.
In a jet turbine, the interblade channels are made tapering (f 4 In turbojet engines, only jet turbines are used. Active turbines are used in turboexpanders, turbopumps. Mechanical work on the turbine shaft is obtained due to the fact that on the blades of the RK, which are under the action of gas-dynamic forces, circumferential forces are created, that is, forces that coincide with the direction of speed. These forces create a torque on the turbine shaft. In a jet turbine, the circumferential force on the RV blades arises for two reasons:

a) an active gas impulse associated with the occurrence of an aerodynamic force P a on a blade in the flow (Fig. 28);

b) due to the reactive force P p , arising during acceleration of the gas jet from the speed W 3 "to W 4> W 3". The forces of Ra and Pp can be decomposed into axial and circumferential components.
The resulting axial components of the active P ao and reactive P ro forces, equal to
ΔР o \u003d Р аo - Р ro, is perceived by the motor rotor bearings.
The resulting circumferential components of the active P ai and reactive pp And forces creates a circumferential force P u= R a u+ Р p u , used to produce torque and net power on the turbine shaft.

3.3. TERMINAL FORCE, GAS EFFICIENCY, TURBINE EFFICIENCY AND POWER

A). Determining the magnitude of the circumferential force P u.
The magnitude of the force R u can be obtained on the basis of the well-known theorem technical mechanics: "The change in the amount of movement of a second mass of gas in the direction of rotation of the impeller (circumferential direction) is equal to the second impulse of the force acting in the same direction."
To compile the equation of momentum, build a combined speed plan for the turbine stage (Fig. 29).

From the combined velocity plan, it can be seen that
W 3 "u \u003d C 3" u - u
W 4 u \u003d u - C 4 u
Δ C u \u003d C 3 "u - C 4 u
When compiling the equation for changing the momentum, we consider the direction of rotation (the direction of peripheral speed u) to be a positive direction.
The final circumferential force is
P u \u003d [kg];
b). Efficient gas operation.
The work of the circumferential force of 1 kg gas Lu is equal to

Where GG — second gas consumption [kg/s].
Substituting the value of the circumferential effort, we get the formula for the work of the circumferential effort

The work of 1 kg of gas transferred to the turbine shaft is called the effective work of gas
Le - This work less work circumferential force by the amount of losses: gas friction, gas overflow in gaps, friction in bearings, vortex formation. The listed losses are small and for powerful turbines amount to 2-3% of total power. Therefore, with sufficient accuracy for practical purposes, it is believed that Le Lu. Then the effective work of the gas is

Thus, the more efficient operation of the gas, the greater the swirling of the gas in the impeller and the circumferential speed or revolutions of the turbine rotor,

V). K p d turbine.

There are losses on the way of converting the adiabatic work of gas expansion in the turbine into mechanical work on its shaft. The amount of losses is taken into account by the effective efficiency of the turbine, which is equal to the ratio of the effective work Le to the adiabatic work of gas expansion in the turbine L hell exp those.

Turbine efficiency η T takes into account both internal (hydraulic) losses and energy losses with output speed. Loss with output speed is relative, since kinetic energy, underused to create power on the turbine shaft, is subsequently used to create jet thrust engine.
For modern single-stage turbojet engines, the efficiency value is equal to η T = 0,7 — 0,86.
G). The power developed by the turbine.
Turbine power is the work done by the gas for one second and transferred to the turbine shaft.
From the definitions, the power of the turbine is;
N T =
Turbine power is determined by the value of the second weight gas flow GG, gas temperature in front of the turbine T 3 *, degree of gas expansion in the turbine π T and turbine efficiency η T . The power of the turbine is the greater, the greater the value of these parameters.
In modern turbojet engines, the power developed by the turbine reaches high values ​​NT = 10,000–50,000 hp. With. and more.
This power is spent mainly on the rotation of the engine compressor and only 2-3% on the drive of service units.

3.4. MAIN PARAMETERS DETERMINING THE POWER OF THE TURBINE

The main parameters that determine the power of the turbine are:
— second weight gas flow GG;
- turbine rotor speed n;
— gas temperature in front of the turbine Tz*;
— degree of reactivity of the turbine ρ .

A). Second weight gas flow GG.
The value of the second gas flow rate can be determined from the continuity equation, given that a critical pressure drop or close to it is usually set in the nozzle apparatus.
This means that in a narrow (critical) section of SA (fcr) critical speed is set Skr, equal to the local speed of sound A. The equation for this case will be written as:

Where γcr is the specific gravity of the gas in the critical section of the SA [kg/m3].
It is known that
, A

Since the pressure and temperature of the gas in the critical section of SA Rkr And Tkr proportional to pressure Rz and gas temperature Tz at the turbine inlet, we can write:
or

.
Thus, at constant temperature gas in front of the turbine Tz gas consumption GG determined by the gas pressure Rz in front of her. Increasing gas pressure Rz leads to an increase in gas consumption and turbine power;

b). Turbine rotor speed n.

At a constant gas temperature in front of the turbine Tz* = Const, increase in turbine rotor speed n leads to an increase in turbine power NT.
This is explained as follows. Increasing the speed of the turbine rotor n(motor rotor) leads to an increase in air consumption GV and the degree of increase in air pressure in the engine compressor πK. Increase πK leads to an increase in pressure at the outlet of the compressor Р2* and at the turbine inlet Р3*= σКСР2*.
An increase in pressure Pz*, on the one hand, increases the gas flow through the turbine G G, on the other hand, the degree of gas expansion in the turbine increases π T. Thus, with an increase in the speed of the turbine rotor, the power of the turbine N t increases due to an increase in gas consumption G G and degree of gas expansion in the turbine πT .
It is known that at Тз*=Const the power of the turbine NT is proportional to the number of revolutions of the turbine n to the power of 2.5, i.e.
NT = f(n2.5)

V). Gas temperature in front of the turbine Tz*
At given and constant speeds of the turbine rotor n= Const an increase in gas temperature in front of the turbine Tz* leads to an increase in turbine power NT , since in this case the adiabatic work of gas expansion in the turbine Ladrash increases, to the first degree, and the gas flow through the turbine GG decreases to the power of 1/2.

The gas temperature in front of the turbine is limited by the strength of the turbine blades. In modern engines, it is equal to Tz* = 1100-1300°K.

G). Degree of reactivity of the turbine ρ .

The degree of reactivity of the turbine characterizes the distribution of the work of gas expansion between the nozzle apparatus and the turbine wheel.
The degree of reactivity of the turbine is the ratio of the adiabatic work of gas expansion in the impeller Laddr to the adiabatic work of gas expansion in the turbine stage Laddress
.
The value of the degree of reactivity of the turbine can vary from 0 to 1, i.e.
0< ρ <1.
At ρ = 0, gas expansion occurs only in the nozzle apparatus, the turbine is purely active, and at p = 1, the turbine is purely reactive.
The degree of reactivity of the turbine affects the efficiency of the turbine, and hence its power. Addiction η T = f(ρ ) is shown in Fig. 30. The nature of the dependence is such that there is an optimal value ρ ≈ 0.5, at which the turbine efficiency takes on a maximum value. This is explained as follows. Degree of gas expansion in the turbine π T= Р3*/Р4 can be considered as the product of the degrees of expansion of the gas in SA π SA\u003d P3 * / Pz "on the degree of gas expansion in the Republic of Kazakhstan π RK = R "3 / P4, i.e. π T = π SA · π RK. For a given degree of gas expansion in the turbine π T increase in the degree of re-activity ρ means an increase in the expansion of gas in the Republic of Kazakhstan, i.e. an increase in π RK. This is possible due to an increase in gas pressure in front of the RC Rz. "The increase in Rz" is accompanied by

a decrease in the absolute C "3 and relative W C" speeds in front of the RK. Reducing the speed W c "leads to a decrease in hydraulic (internal) losses, and consequently, to an increase in turbine efficiency η m. On the other hand, an increase in the expansion of gas in the Republic of Kazakhstan with an increase in the degree of reactivity of the turbine ρ leads to an increase in losses with output speed (kinetic energy increases), which leads to a decrease in turbine efficiency η T.

3.5. JOINT OPERATION OF TURBINE AND COMPRESSOR IN TJD

Since in the TRD system the compressor and turbine are connected by a common shaft, their work is interdependent. The interdependence of their work, in addition to mechanical connection, is due to the total flow of air through the compressor and gas through the turbine, which determine their power.
The power developed by the turbine Nt is the available power. It can be equal to, more or less than the required power for the rotation of the compressor NK;
Depending on this, the following modes of joint operation of the turbine and compressor are distinguished:
1. Equilibrium mode, when Nt = NК;
2. Acceleration mode (increase in engine speed), when NT > NK;
3. Braking mode to reduce engine speed), when Nt< NК.
It is obvious that it is possible to change the mode of operation of the engine (control the engine) by changing the power of the turbine.
The most convenient parameter with which you can change the power of the turbine is the gas temperature in front of the turbine Tz *. The change in Tz* is achieved by changing the amount of fuel Gt supplied to the combustion chamber of the engine.
It was previously shown that the power required to rotate the compressor NK is proportional to the engine speed n to the third degree, i.e.
NK = f (n3),
and the power developed by the turbine Nt, at a given and constant gas temperature in front of it Tz * = Const, is proportional to the number of revolutions n to the power of 2.5, i.e.
NT=f(n2.5).
The combined graphs of dependences NК = f (n) and NT = f (n) are shown in fig. 31. The graph shows that with an increase in the engine speed, the compressor power NK grows faster than the turbine power Nt.

The power of the turbine is proportional to the gas temperature Tz*.
Curve 1 on the graph shows the dependence NT= f (n) at Тз*max = Сonst, and curves 2, 3, 4... at lower but constant temperatures Тз*.
At the points of intersection of curves 1, 2, 3, 4... with the curve NK = f (n), the compressor and turbine powers are equal, i.e. N T \u003d N K. These points determine the equilibrium modes. Minimum nmin and maximum nmax engine revolutions are achieved at Т3*=Т3*max. Turnovers less than nmin or greater than nmax can only be obtained by raising the temperature above the maximum permissible T 3 * ma x , which can lead to turbine failure.
With an increase in speed from nmin to nmax, the gas temperature in front of the turbine T3* first decreases from T 3 *max to T 3 *min at medium speed (Fig. 31), and then increases again to T3*max at n = nmax. This nature of the change in temperature Т3* is explained by the conditions of joint operation of the compressor and turbine in the TRD system and is due to the different law of change in NK and NT with respect to the number of revolutions.
The high value of Tz* at nmax and nmin indicates a high thermal stress of the engine in these modes. Therefore, the operation of the engine at maximum speed nmax is allowed for a limited time (5-10 minutes), and the speed of small gas n mg usually 1000-1500 rpm exceed nmin i.e.
n mg\u003d (1000-1500) rpm + nmin.
When starting the engine in the rev range where NT< NК раскрутка ротора турбокомпрессора производится с по-мощью пусковых двигателей (электростартеров, турбодетандеров и др.). Сначала в раскрутке ротора принимает участие только пусковой двигатель, затем в работу вступает турбина и раскрутка ротора до оборотов nmg continues jointly with the starting engine and turbine. At rpm n mg or several smaller, but larger nmin, the starting motor is automatically switched off.
Time of continuous operation at n mg is also limited, since T3 * is relatively large, and the cooling efficiency of turbine parts in this mode is insufficient.
To increase engine speed above n mg it is necessary to increase the power of the turbine, which is achieved by increasing the supply of fuel to the combustion chamber. At the same time, the gas temperature Tz* increases, an excess of turbine power Nt appears, and the engine rotor spins up to speed at which N T = N K (curves a and b in Fig. 31). Reducing the rotor speed is achieved by reducing the fuel supply to the combustion chamber, reducing Tz * and Nt. Turnovers fall to a value at which again N T \u003d N K (curve in in Fig. 31).

3.6. MULTI-STAGE TURBINES AND FEATURES OF OPERATION OF TURBINES OF TWO-SHAFT ENGINES
1. Multistage turbines

The capabilities of a single-stage turbine are limited by the maximum (critical) pressure drop in the nozzle apparatus, when at the exit from it (the critical section of the oblique cut) the flow velocity reaches the speed of sound. This pressure difference (it is approximately 2) provides the adiabatic work of gas expansion
Lhell exp≤ 25,000–30,000 kg m/kg at a gas temperature at the turbine inlet of 850–960 °C and a circumferential velocity at an average radius equal to U=350—370m/sec.
When a greater pressure drop needs to be generated in the turbine in order to obtain a greater amount of power, two or multi-stage turbines are used.
A multi-stage turbine, in comparison with a single-stage one, has the following advantages:
a) lower gas energy losses in the flow path, which is due to lower flow rates due to lower pressure drops in each stage;
b) using the heat recovery effect. Due to gas friction, heat is released, which in a single-stage turbine is a loss, and in a multi-stage turbine it is partially used in the next stage;
c) better use of the gas output speed from the previous ones in subsequent stages, which reduces losses with the output speed and increases the efficiency of the turbine.
The disadvantages of multistage turbines are:
a) Structural complexity;
b) Increase in length and weight (however, the diameter of a multi-stage turbine is less than a single-stage one);
c) High temperature regime of the blades of the first stage and worse conditions for cooling the blades of the second and subsequent stages.
In modern turbojet engines, two and three-stage turbines are widely used.

2. Features of operation of turbines of twin-shaft engines

The turbine of a two-shaft engine is two-stage, but there is only a gas-dynamic connection between the stages. The impeller of the turbine of the first stage drives the rotor of the high-pressure compressor (HPR), and the impeller of the second stage drives the rotor of the low-pressure compressor (RPR). The diagram of the high and low pressure rotors is shown in fig. 32.
The first stage of the turbine (HPR) and the second stage of the turbine (RND) are designed in such a way that critical (or close to it) pressure drops are established in the nozzle devices at the calculated and close to it modes. The distribution of the work of expanding the gas between stages when changing the operating modes of the engine occurs automatically. This is due to the following main reasons.

A). When the engine speed changes, the degree of gas expansion at the turbine stages in a certain range of modes, when the pressure drop in the engine outlet nozzle is close to critical, remains practically constant, i.e.
π turbojet and π TRND \u003d Const, and therefore,
π TΣ = π TRVD · π TRND = Const;
b). With a constant degree of expansion of the gas in the turbine, the efficiency of the turbine remains unchanged, i.e.
η turbojet and η TRND = Сonst ;
V). Since the efficient operation of the turbine
L THIS = ,
then Letrnd and Letrvd are proportional only to the gas temperature before the turbine stage Tz*rn d and Tz*rvd, respectively. When changing the engine operating mode, there is a proportional change in Tz * rnd and Tz * rvd.
Therefore, the distribution of available effective work between the stages remains unchanged, i.e.
LETRND / LEТ RVD = Const .
It is known that engine throttling leads to an increase in the work required to rotate the low-pressure compressor (stages "heavier") and a decrease in the required work to rotate the high-pressure compressor (stages "facilitate"). With a constant distribution of the available work between the turbine stages, this leads to a more intensive reduction in the speed of the HPP than the HPP;
G). With a significant throttling of the engine, when a subcritical pressure drop is established at the outlet, the overall expansion ratio decreases
gas in the turbine π TΣ, mainly due to the fall π TRND and LETRND, and π TRVD almost does not change. This leads to an even more intense drop in the speed of the RPR in comparison with the HPH, which contributes to the stable operation of the two-stage compressor.

Air compression in turbojet compressors.
1.1. Requirements for turbojet compressors and types of compressors.
1.2. Air compression in centrifugal compressors.
1.3. Unstable operation of a centrifugal compressor and measures to combat it.
1.4. Compression of air in axial compressors.
1.5. Unstable operation of the axial compressor and the fight against it.
2. Organization of the combustion process in the combustion chambers of turbojet engines.
2.1 Purpose of combustion chambers.
2.2 Basic requirements for combustion chambers and assessment of their implementation.
2.3. Types of combustion chambers and their arrangement.
2.4. The principle of operation and the working process of the combustion chamber.
2.5. Dependence of the completeness and stability of combustion on operating conditions.
3. The process of gas expansion in the turbine.
3.2 Scheme and principle of operation of the axial stage.
3.3. Circumferential force, gas efficiency, turbine efficiency and power.
3.4. The main parameters that determine the power of the turbine
3.5 Joint operation of the turbine and compressor in the turbojet engine.
3.6. Multistage turbines and operation features of twin-shaft engine turbines.

The methodical manual was compiled by the master p / o Zabolotny V.A.

Please read before asking a question: FAQ

• Further

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 rather 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 ) -

The invention relates to the field of aircraft gas turbine engines, in particular to a unit located between a high pressure turbine and a low pressure turbine of the internal circuit of a bypass aircraft engine. The continuous annular transition channel between the high pressure turbine and the low pressure turbine with an expansion ratio of more than 1.6 and an equivalent opening angle of a flat diffuser of more than 12° contains perforated outer and inner walls. The flow swirl behind the impeller of the high-pressure turbine is transformed in the direction of its strengthening at the walls and weakening in the center. The swirl is transformed by profiling the high-pressure turbine stage and by a swirler located behind the impeller of the high-pressure turbine with a height of 10% of the channel height, 5% of the height on the inner and outer walls of the channel, or by a full-height twisting-untwisting device. EFFECT: invention allows to reduce losses in the transition channel between high and low pressure turbines. 2 w.p. f-ly, 6 ill.

The field of technology to which the invention belongs

The invention relates to the field of aircraft gas turbine engines, in particular to a unit located between a high pressure turbine and a low pressure turbine of the internal circuit of a bypass aircraft engine.

State of the art

Aircraft gas turbines of bypass engines are designed to drive compressors. The high pressure turbine is designed to drive the high pressure compressor, and the low pressure turbine is designed to drive the low pressure compressor and fan. In aircraft engines of the fifth generation, the mass flow rate of the working fluid through the internal circuit is several times less than the flow rate through the external circuit. Therefore, the low-pressure turbine in its power and radial dimensions is several times higher than the high-pressure turbine, and its rotational speed is several times less than the rotational speed of the high-pressure turbine.

This feature of modern aircraft engines is structurally embodied in the need to make a transition channel between the high pressure turbine and the low pressure turbine, which is an annular diffuser.

Severe restrictions on the overall and mass characteristics of an aircraft engine in relation to the transition channel are expressed in the need to make a channel of a relatively short length, with a high degree of diffuseness and a clearly detachable equivalent opening angle of a flat diffuser. The degree of diffuseness is understood as the ratio of the output cross-sectional area to the input. For modern and advanced engines, the degree of diffuseness has a value close to 2. The equivalent opening angle of a flat diffuser is understood as the opening angle of a flat diffuser having the same length as the annular conical diffuser and the same degree of diffuseness. In modern aviation gas turbine engines, the equivalent opening angle of a flat diffuser exceeds 10°, while a continuous flow in a flat diffuser is observed only at an opening angle of no more than 6°.

Therefore, all the designs of the transition channels are characterized by a high loss factor, due to the separation of the boundary layer from the diffuser wall. The figure 1 shows the evolution of the main parameters of the transition channel of the company General Electric. In figure 1, the degree of diffuserity of the transition channel is plotted along the horizontal axis, and the equivalent opening angle of the flat diffuser is plotted along the vertical axis. Figure 1 shows that the initially high values ​​of the effective opening angle (≈12°) evolve to significantly lower values, which is associated only with a high level of losses. According to the results of studies of an annular diffuser with an opening degree of 1.6 and an effective opening angle of a flat diffuser of 13.5°, the loss coefficient varied from 15% to 24% depending on the distribution law of the swirl along the channel height .

Analogues of the invention

Distant analogues of the invention are diffusers described in patents US 2007/0089422 A1, DAS 1054791. In these designs, to prevent separation of the flow from the wall of the diffuser, the suction of the boundary layer is used from the section located in the middle of the channel with the ejection of the exhausted gas into the nozzle. However, these diffusers are not transition channels between the high pressure turbine and the low pressure turbine.

Brief description of the drawings

Non-limiting embodiments of the present invention, its additional features and advantages will be described in more detail below with reference to the accompanying drawings, in which:

figure 1 depicts the evolution of the flow part of the inter-turbine transition channel in the turbofan engine from General Electric,

figure 2 depicts the dependence of the losses of the kinetic energy of the flow in the channel on the integral parameter of the swirl of the flow F ¯ C T in the form of a linear approximation, where ν=0 is the swirl of the flow uniform in height; ν=-1 - flow swirl increasing in height; ν=1 - flow swirl decreasing in height; y \u003d -1.36F st +0.38 - approximation dependence corresponding to the reliability coefficient R \u003d 0.76,

figure 3 depicts the extrapolation of separation losses in the annular diffuser from the value of wall swirl,

figure 4 depicts a diagram of the transition channel,

figure 5 depicts a scheme of perforation,

Fig.6 depicts a diagram of the power rack device with a supply channel.

Disclosure of invention

The problem to which the present invention is directed is to create a transition channel with an opening degree of more than 1.6 and with an equivalent opening angle of a flat diffuser exceeding 12°, the flow in which would be unseparated, and the level of losses, respectively, is minimally possible. It is proposed to reduce the loss factor from 20-30% to 5-6%.

The task is solved:

1. Based on the transformation of the existing swirl behind the high-pressure turbine at the inlet to the annular diffuser in the direction of its strengthening on the inner and outer walls of the channel and weakening in the middle of the channel.

2. Based on the perforation of the inner and outer walls of the annular diffuser, variable in length, adapted to the local structure of turbulence.

3. Based on the suction of the boundary layer from the zone of possible separation of the flow from the walls of the diffuser.

In this connection, a non-separated annular transition channel is proposed between the high pressure turbine (HPT) and the low pressure turbine (LPT) with an expansion ratio of more than 1.6 and an equivalent opening angle of a flat diffuser of more than 12°, containing an outer wall and an inner wall. The outer and inner walls are perforated, and the swirling behind the impeller of the high-pressure turbine (HPT) is transformed in the direction of its strengthening at the walls and weakening in the center. The twist is transformed by profiling the stage of the high pressure turbine (HPT) and by means of a swirler located behind the impeller of the high pressure turbine (HPT) with a height of 10% of the channel height, 5% of the height on the inner and outer walls of the channel, or by twisting untwisting device full height.

The converted twist is limited by the achievement of the integral parameter of the twist to the level F article =0.3-0.35. The perforation section, located at a distance of 0.6-0.7 of the length of the transition channel from the inlet section, is connected to the cavity in the power racks, which have slots at 80% of the height of the racks symmetrically to the geometric middle of the channel, and the slots are located near the inlet edge.

As is known, the gas moves in the diffuser by inertia in the direction of pressure growth, and the separation (delamation) of the flow from the walls is physically due to the insufficient inertia of the inner near-wall layers of the boundary layer. Points 1, 2 are designed to increase the inertia of the near-wall gas flow by increasing the speed of movement, and, accordingly, its kinetic energy.

The presence of swirling in the near-wall gas flow increases the speed of movement, and hence its kinetic energy. As a result, the resistance of the flow to separation (delamination from the walls) increases, and the losses decrease. Figure 2 shows the results of an experimental study of an annular diffuser with a degree of disclosure of 1.6 and an equivalent opening angle of a flat diffuser of 13.5°. The vertical axis shows the loss factor, defined in the traditional way: the ratio of mechanical energy losses in the diffuser to the kinetic energy of the gas flow at the diffuser inlet. The horizontal axis represents the integral swirl parameter defined as follows:

F with t \u003d F in t + F p e r F.,

where Ф. = 2 π ∫ R R + H ρ w u r 2 d r 2 π ∫ R R + H ρ w 2 r d r (R + H 2)

The integral swirl parameter at the channel inlet, ρ is the density, w is the axial velocity, u is the circumferential velocity, r is the current radius, R is the radius with the inner generatrix of the diffuser, H is the channel height, Ф w is the integral swirl parameter considered in the range heights from 0% to 5% of the sleeve section, i.e.

Ф в t \u003d 2 π ∫ R R + 0.05 H ρ w u r 2 d r 2 π ∫ R R + H ρ w 2 r d r (R + H 2) ;

Ф lane - the same parameter, but in the height range from 95% to 100% of the sleeve section, i.e.

Фper = 2 π ∫ R + 0.95 H R + H ρ w u r 2 d r 2 π ∫ R R + H ρ w 2 r d r (R + H 2) .

As can be seen from figure 2, the losses in the transition channel are reduced as the proportion of near-wall twist increases.

The figure 3 shows a linear extrapolation of the dependence ξ (Ф st) to the level of friction losses in an equivalent channel of constant cross section. In this case, the near-wall swirl (10% of the channel height) should account for approximately 30% of the flow swirl.

As is known, in the case of a turbulent flow regime in channels, a laminar flow regime takes place immediately near the wall due to the impossibility of transverse pulsating motion. The thickness of the laminar sublayer is approximately 10 μ ρ τ s t. In the last expression, μ is the dynamic viscosity, τ st is the friction stress on the wall. As is known, the friction stress decreases rapidly along the diffuser, and at the separation point it is generally equal to zero. Therefore, the thickness of the laminar sublayer in the transition channel with a solid wall increases rapidly along the flow. Correspondingly, the thickness of the near-wall layer of the flow with a low level of kinetic energy increases.

The perforation of the inner and outer walls of the transition channel makes possible the transverse pulsating movement at any distance from the perforated wall. Since in a turbulent flow the longitudinal fluctuating flow is statistically related to the transverse flow, perforation makes it possible to increase the area of ​​the turbulent flow proper. The higher the degree of wall perforation, the thinner the laminar sublayer, the higher the gas velocity in the near-wall layer, the higher the kinetic energy of the near-wall flow and its resistance to separation (delamination from the wall).

Description of the design of the transition channel between the high pressure turbine and the low pressure turbine

The transition channel between the high-pressure turbine (HPT) and the low-pressure turbine (LPT) of the internal circuit of a bypass turbojet engine (Figure 4) is an annular diffuser having an inner wall 1 and an outer wall 2. The inner and outer walls at the junction with the HPT and LPT have defined radii.

Power racks 3 pass through the transition channel, which provide lubrication, ventilation and cooling of the HPT and LPT rotor supports. Racks 3 have an asymmetric aerodynamic profile in cross-section, which provides flow spin-up in the center of the channel and flow twist at the channel walls to the level Ф st =0.3-0.35.

Walls 1 and 2 are perforated (Figure 5). To avoid overflow of the working fluid in the perforations, the parts of the perforation 4 are isolated from each other by transverse walls 5.

From the perforation section 9, located at a distance of 0.6-0.7 from the entrance to the diffuser, suction and removal through the supply channel 6 into the slots 7 of the racks 3 are organized. minimum local static pressure. In the channel connecting the cavity 9 with the cavity of the uprights 3, there are measuring washers 8 that regulate the gas flow.

Behind the HPT impeller 11, a twisting device 12 is installed, which increases the flow swirl near the walls. The height of the blades of the apparatus 12 is 10% of the height of the channel at the inlet. If necessary, the curler 12 can be converted into a untwisting-twisting device located along the entire height of the channel. The central part of the device spins the flow, and the near-wall twists it, so that as a result of the swirl of the flow at the inlet to the diffuser, it is Ф st = 0.3-0.35.

In the event that a continuous flow in the diffuser is achieved only due to the profiling of the nozzle apparatus 10 and the impeller 11 of the HPT and the twisting-unwinding effect of the power racks 3, the swirling device 12 and the slot 7 with the channel 6 are absent.

Implementation of the invention

The non-separated flow regime in the transition channel is achieved by swirling the flow in the near-wall flow zones, spinning the flow in the center, perforating the meridional generatrix of the transition channel, and suction of the boundary layer.

Features of the organization of the working process in modern gas turbine engines are such that behind the high-pressure turbine there is a flow swirl of about 30-40 °. A high level of swirling at the inner and outer walls (at a distance of 5% of the channel height) should be maintained, and if necessary, strengthened by step profiling and, if necessary, by installing a swirling blade apparatus at the transition channel inlet. Flow swirling at heights from 5% of the sleeve section to 95% of the same section should be reduced both by profiling the step and by spinning the flow with power racks structurally passing through the channel. If necessary, to achieve the desired flow promotion, install an additional spinning vane at the inlet to the transition channel. The flow spin-up in the central part of the channel is designed to reduce the radial static pressure gradient and reduce the intensity of secondary flows that thicken the boundary layer and reduce its resistance to separation. The value of the relative near-wall twist should be as close as possible to the value of 0.3-0.35.

Since the installation of an additional blade apparatus is associated with the appearance of losses in this apparatus, it should be installed only if the decrease in the loss factor in the transition channel significantly exceeds the loss in the additional twisting and untwisting device. As an option, it is possible to install an additional twisting apparatus on the sleeve and the periphery, limited by heights from 5% to 10% H (Figure 4).

Perforation of the meridional generatrix of the transition channel changes the flow regime in the laminar sublayer to turbulent. Extrapolation of the logarithmic velocity profile to the region of the laminar sublayer up to a distance from the solid wall equal to 8% of the thickness of the laminar sublayer gives the velocity value τ с r ρ 6.5, which is only 2 times less than the velocity at the boundary of the laminar sublayer, while as the flow velocity in the laminar sublayer itself (at this distance) is 4 times less, and the specific kinetic energy is 16 times less.

Extrapolation of the logarithmic law of velocity distribution, which is characteristic of a purely turbulent flow regime to the area of ​​the laminar sublayer, suggests complete freedom for the movement of turbulent eddies. This possibility exists under two conditions: 1) the degree of perforation of the solid surface is close to 100%;

2) turbulent eddies of all sizes in a given section have complete freedom to move in the transverse direction.

In reality, these conditions are unattainable in full, but in practice it is possible to come close to them. As a result, the speed of movement near the perforated surface will be several times higher than the speed of movement at the same distance from the wall near the solid surface. In this case, the density of the perforation elements and its structure must be consistent with the maximum energy spectrum of turbulent fluctuations in relation to their linear size for a given section of the transition channel.

The perforation density (the ratio of the perforation area to the total area) should be kept as high as possible for structural and rigidity reasons.

The perforation structure is adapted to the linear size of energy-containing local turbulence vortices, which is determined by the height of the transition channel and its average radius in a given section. The following model can be taken as a perforation structure model:

d min \u003d (0.2-0.5) l e (R, II);

d max \u003d (1.5-2) l e (R, II);

d ¯ = (0.6 − 0.8) ;

d min ¯ = (0.2 − 0.3) ;

d max ¯ = (0.1 − 0.2) ;

d min - minimum perforation diameter; d=l e (R, II) - the main diameter of the perforation, equal to the linear size of the energy-containing vortices of the turbulent structure; d max - maximum perforation diameter; d ¯ = S d S - proportion of the main perforation size; S d - perforation area, made according to the size d=(l e (R, II); S - total perforation area; d min ¯ = S d min S - proportion of the minimum perforation size; S dmin - perforation area, made to size d min ; d max ¯ = S d max S - proportion of the maximum perforation size; S dmax - perforation area, made according to the size d max (Fig.5).

The size of energy-containing vortices l e (R, II) is determined by calculation, depending on the adopted turbulence model.

In transition channels with a very high degree of expansion (n>2) and a very large equivalent opening angle of a flat diffuser (α equiv >17°), the maximum achievable near-wall swirl (Ф st ≈0.3) and the maximum achievable and properly structured perforation (S ¯ ≈ 0.8, where S ¯ = S per S, S per - the total area of ​​the perforated surface, S - the total area of ​​the meridional contours) may not be enough to organize a continuous flow along the entire length of the transition channel. In this case, a possible detachment in the last third of the diffuser length should be prevented by suction of the boundary layer through part of the perforation. The removal of the sucked gas should be organized in the central part of the channel through the corresponding holes in the power drains, which are located near the inlet edge of the wall profile, i.e. where local static pressure is minimal. The area of ​​the part of the perforation 9, working for suction, and the area of ​​the flow sections in the racks 7 must be consistent with each other.

The cavity in the power racks has slots located near the input edge, the vertical length of which can reach 0.8 of the height of the racks. The slots are located symmetrically with respect to the middle of the channel. The set of cavities and channels associated with perforations and slots in power racks organizes the suction of the boundary layer in the transition channel.

The organization of boundary layer suction is expedient only if the mixing loss during the blowing of the sucked gas to the inlet to the transition channel is less than the decrease in losses in the diffuser due to suction.

List of used literature

1. Gladkov Yu.I. Investigation of the variable along the radius of the inlet flow swirl on the efficiency of inter-turbine transition channels of the GTE [Text]: abstract of the dissertation for the degree of candidate of technical sciences 05.07.05 / Yu.I.Gladkov - Rybinsk State Aviation Technological Academy named after P.A.Soloviev. - 2009 - 16 p.

2. Schlichting, G. Theory of the boundary layer [Text] / G. Schlichting. - M.: Nauka, 1974. - 724 p.

1. A non-separated annular transition channel between a high-pressure turbine (HPT) and a low-pressure turbine (LPT) with an expansion ratio of more than 1.6 and an equivalent opening angle of a flat diffuser of more than 12°, containing an outer wall and an inner wall, characterized in that the outer and the inner wall is perforated, and the swirling behind the high-pressure turbine (HPT) impeller is transformed in the direction of its strengthening near the walls and weakening in the center due to the profiling of the high-pressure turbine (HPT) stage and due to the swirling device located behind the high-pressure turbine impeller (TVD) with a height of 10% of the height of the channel, 5% of the height on the inner and outer walls of the channel, or due to a twisting-untwisting device of full height.

2. The channel according to claim 1, characterized in that the converted twist is limited to the achievement of the integral parameter of the twist to the level of F article =0.3-0.35.

3. The channel according to claim 1, characterized in that the perforation section, located at a distance of 0.6-0.7 of the length of the transition channel from the inlet section, is connected to the cavity in the power racks having slots at 80% of the height of the racks symmetrically to the geometric middle of the channel , and the slots are located near the leading edge.

Similar patents:

The invention relates to the field of energy, mainly for waste steam systems of thermal power plants, for example, steam emissions during operation of the main safety valves of boilers, blowdowns of superheaters, grinding of boilers and waste heat boilers at discharge steam flow rates of more than 30 t/h and the degree of underexpanded steam jet underexpansion n=pa/pc>1, where pa - atmospheric air pressure, pc - static steam pressure at the outlet of the exhaust pipeline

The exhaust device of the turbomachine comprises a housing with an inlet located around the axis of rotation of the turbine, a diffuser, an outlet located in the outer wall of the housing, and an additional partition. The diffuser includes axial and radial parts formed, respectively, by inner and outer duct walls located inside the housing around the axis of rotation of the turbine. The additional partition is made inside the device housing in a plane perpendicular to the turbine rotation axis, with a perimeter equal to the perimeter of the device housing walls parallel to it. In the additional partition, a hole is made coaxial to the axis of rotation of the turbine, the diameter of which is equal to the maximum diameter of the outer duct wall of the radial part of the diffuser. In the lower part of the additional partition, through grooves are made symmetrically and "mirror" relative to the vertical axis of the said partition. Along the perimeter of the through grooves, hollow boxes made in the form of truncated pyramids with two curvilinear faces are motionlessly and hermetically installed. Smaller bases of said truncated pyramids are directed towards the turbine of the device, the space from the upper edge of the additional partition to the upper edge of the housing wall containing the inlet of the device is closed by a sealed flat wall. The invention improves the efficiency of the device and efficiency. gas turbine plant. 3 ill.

The invention relates to the design of support or mounting devices for the turbine outlet. The turbine outlet device contains hollow aerodynamic profiled racks located behind the impeller of the last turbine stage, as well as aerodynamic profiled contours. The contours are formed by the front and rear blades placed between the racks with offset relative to each other. The middle lines of the input sections of the contours and the input sections of the profiled racks are turned in the direction of rotation of the impeller of the last turbine stage at an angle of 20-40° to its longitudinal axis. The middle lines of the outlet sections of the circuits are directed along the longitudinal axis of the turbine. The blades are installed with an offset relative to each other by a distance equal to 0.03÷0.15 of the front blade chord length. Along the length of the contour chord, the blades are installed in the position of matching the front of the trailing edge of the front blade and the front of the trailing edge of the rear blade or are displaced relative to it. The number of circuits installed between the racks is determined by the relationship protected by the present invention. EFFECT: invention makes it possible to increase the efficiency of the last stage of the turbine, as well as to reduce the swirl of the outgoing flow. 3 ill.

The invention relates to exhaust devices and can be used as part of a gas compressor unit with a gas turbine plant. The exhaust device contains a diffuser, an adapter with flow-dividing ribs and a cassette-type silencer placed at an angle of 30-60° to the adapter axis. Each of the silencer cassettes consists of a power frame sheathed with sheets, the cavity between which is filled with sound-absorbing material. On the side inclined towards the diffuser, the cassettes are sheathed with a perforated sheet, and on the opposite side - solid. EFFECT: invention makes it possible to increase the efficiency of noise reduction in the output device by ensuring a uniform flow movement. 2 ill.

The invention relates to mechanical engineering and can be used in the exhaust tract of a gas compressor unit or a gas turbine power plant. The diffuser of the exhaust path of the gas turbine plant contains a shell with flanges, a casing covering the shell and sound insulation placed between the shell and the casing. The shell is made of movable, telescopically connected parts with movement limiters. The casing is formed by an elastic material, for example fabric "Atom", fixed on the shell. The invention will improve the reliability of the design of the diffuser, as well as reduce its metal consumption. 3 ill.

An outlet for use with a multi-stage turbine is configured to direct steam from the turbine to the condenser and includes a support cone surrounding the turbine rotor, a guide, and a guide cap. The guide is located radially outside the support cone, while the guide and the support cone are configured to direct fluid from the turbine. The guide cap extends from the edge and rear surface of the guide to the turbine and helps to prevent the formation of fluid vortices in the outlet. Another invention of the group relates to a steam turbine, including the above outlet pipe. EFFECT: group of inventions makes it possible to increase the performance of the turbine. 2 n. and 6 z.p. f-ly, 5 ill.

The invention relates to energy. Part of the low pressure of the steam turbine, including the control element at the inlet, a group of stages with intermediate chambers and an exhaust pipe connected to the condenser, divided by a pipe system into inlet and outlet volumes, while the outlet volume of the condenser is connected to the intermediate chamber, for example, before the last stage, through a bypass pipe with a valve. The proposed technical solution is based on the operation of the last low-pressure stage at low steam flow rates, when its impeller does not generate power, but receives it from the rotor, spending it on pumping steam towards the exhaust. With this "compressor" mode of operation, the pressure before the last stage is lower than in the condenser. This allows directing into the chamber before the last stage the steam cooled by the condenser pipe system while flowing from its inlet volume to the outlet volume. The claimed invention makes it possible to increase the reliability and efficiency of a steam turbine at low steam flow rates through a group of stages of the low-pressure part by reducing the ventilation heating of the flow part and eliminating its consequences without the use of cooling moisture injections, which increase erosion, and without increasing the flow rate of working steam, which reduces heat supply and electricity. 1 ill.

The invention relates to the field of aircraft gas turbine engines, in particular to a node located between the high pressure turbine and the low pressure turbine of the internal circuit of a bypass aircraft engine