Aircraft engines. Cooling system of the last stage of the low-pressure axial turbine of a bypass turbojet engine Relative flow velocity at the inlet to the AC

03.03.2020

Transmission

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- for 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 takeoff run of the aircraft 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 ) -

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Ministry of Education and Science of the Russian Federation

Federal Agency for Education

Samara State Aerospace University

named after Academician S.P. Queen

Department of the Theory of Aircraft Engines

Course work

on the course: "Theory and calculation of blade machines"

Axial turbine designaviationengineJT9 D20

Samara 2008

Exercise

Perform a design calculation of the main parameters of the high-pressure turbocharger and construct a meridional section of the high-pressure turbine of the JT9D-70A turbofan engine, perform a thermodynamic calculation of the turbine, a kinematic calculation of the second stage of the turbine, and profile the impeller blade in three sections: sleeve, middle and peripheral sections.

The initial parameters of the turbine are known from the thermodynamic calculation of the engine in takeoff mode (H P =0 and M P =0).

Table 1. Initial data for turbine design

high pressure turbine
Parameter	Numerical value	Dimension















TTND = TT
RTND = RT

Essay

Coursework on thermogasdynamic design of the JT9D20 axial turbine.

Explanatory note: 32 pages, 1 figure, 2 tables, 3 appendices, 4 sources.

TURBINE, COMPRESSOR, FLOW PART, IMPELLER, NOZZLE DEVICE, STAGE, FLOW OUTPUT ANGLE, EFFECTIVE ANGLE, PROFILE SETTING ANGLE, GRID PIT, GRID WIDTH

In this course work, the diametrical dimensions of the high-pressure turbine were calculated, the meridional section of the flow path was constructed, the kinematic calculation of the stage at the average diameter and the calculation of the parameters for the blade height with the swirl law b = const were performed with the construction of velocity triangles at the inlet at the outlet of the RC in three sections (sleeve, peripheral and section on the average diameter). The profile of the blade of the impeller of the second stage was calculated, followed by the construction of the contour of the profile in the lattice in three sections.

Conventions

D - diameter, m;

Relative bushing diameter;

h - blade height, m;

F - cross-sectional area, m 2;

G - mass flow rate of gas (air), kg/s;

H - flight altitude, km; compressor head, kJ/kg;

i - specific enthalpy, kJ/kg;

k is the isentropic index;

l - length, m;

M - Mach number;

n - speed, 1/min;

Р - pressure, kPa;

Reduced speed;

s - flow velocity, m/s;

q(), (), () - gas dynamic functions of;

R - gas constant, kJ/kggrad;

L * k(t) - specific work of the compressor (turbine);

k(t) - efficiency of the compressor (turbine);

S - axial width of the crown, m;

T - temperature, K;

Assigned resource, h;

V - flight speed, m/s;

z - number of steps;

k, t - the degree of increase (decrease) of the total pressure;

The coefficient of restoration of the total pressure of air (gas) in the engine elements; tensile stresses, MPa;

Mass flow change factor;

U - circumferential speed, m/s;

Y t * =U t cf /C * t s - turbine load parameter;

Gap size, m;

U 2 t cf h t out /D cf out - stress parameter in the turbine blades, m 2 /s 2;

K tk, K tv - matching parameters of the gas generator, turbofan.

Indices

a - axial component;

c - air section at the compressor inlet

vent - fan

vzl - takeoff;

w - bushing section;

d - gases section at the outlet of the turbine

k - compressor section at the outlet of the compressor

kr - critical

ks - combustion chamber

n - cross section of the undisturbed flow

on - guide apparatus;

cool - cooling;

n - flight parameter, peripheral diameter;

pr - given parameters;

ps - retaining stage

s - isentropic parameters;

c - second section at the exit of the nozzle

cp - average parameter;

st - step parameter;

t - turbine fuel section at the turbine inlet

h - hourly

* - braking parameters.

Abbreviations

HP - high pressure;

LP - low pressure;

VNA - input guide vane;

GDF - gas dynamic functions

GTE - gas turbine engine

Efficiency - efficiency factor;

ON - guide vane;

RK - impeller;

SA - turbine nozzle apparatus;

SAU - standard atmospheric conditions

Turbofan engine - turbojet bypass engine.

Introduction

1. Design calculation of the main parameters of the high pressure turbine

1.1 Calculation of the geometrical and operating parameters of the HP turbine

1.2 Construction of the meridional section of the HP turbine flow path

2. Gas-dynamic calculation of the HP turbine

2.1 Distribution of heat drop by steps

2.2 Calculation of the step by average diameter

2.3 Calculation of the effective operation of the stage, taking into account friction losses of the disk and in the radial clearance

2.4 Calculation of flow parameters at different radii

Conclusion

List of sources used

Introduction

This work contains a simplified version of the gas-dynamic calculation of an axial turbine, in which the variant search for optimal (compromise) parameters is replaced by reliable statistical recommendations obtained by systematizing materials for the calculation of turbines of modern gas turbine engines. The design is carried out according to the initial parameters obtained in the thermogasdynamic calculation of the engine.

The purpose of designing an axial aircraft turbine is to determine the main geometric, kinematic and thermodynamic parameters as a whole and its individual stages, which provide the calculated values of the specific and general parameters of the engine. In this regard, the design tasks involve: selection of the main geometric parameters of the turbine being designed for given parameters of the working fluid, taking into account the intended purpose of the gas turbine engine; distribution of heat drop over the steps, calculation of flow parameters in the gaps between the steps; calculation of flow parameters in the elements of the flow path of the second stage of the turbine at the average diameter; selection of the swirl law and calculation of changes in flow parameters along the radius (blade height) of the designed stage; performing profiling of working blades of the designed stage.

1. Design calculation of the main parameters of the turbine of high

pressure

1.1 Calculation geometric and regime parameters HP turbines

The geometric parameters of the turbine to be determined are shown in Figure 1.

Figure 1. - Geometric model of an axial turbine

1. The value of the ratio D cf / h 2 (h 2 - the height of the rotor blades at the outlet of the HP turbine) is determined by the formula

where e t is the stress parameter, the value of which is usually within (13 ... 18) 10 3 m 2 / s 2.

We accept e t \u003d 15 10 3 m 2 / s 2. Then:

In order to obtain high efficiency, it is desirable to have. Therefore, a new value is chosen. Then,

2. Given the value of the axial gas velocity at the turbine inlet (C 0 =150 m / s), determine the reduced axial velocity l 0 (l 0 = 0.20 ... 0.25)

Annular area at the inlet to the SA of the HP turbine:

3. Calculate the annular area at the outlet of the turbine. To do this, the magnitude of the axial velocity component at the outlet of the turbine is preliminarily estimated. We accept that /= 1.5; . Then

4. According to the selected value, the height of the working blade at the outlet of the HP turbine is determined:

5. Average diameter at the HP turbine outlet

6. Peripheral diameter at the outlet of the valve:

7. Sleeve diameter at the outlet of the valve:

8. The shape of the flow part looks like: Therefore:

The height of the nozzle vane at the turbine inlet is estimated as follows:

9. Peripheral diameter of the nozzle apparatus at the HP turbine inlet:

10. Sleeve diameter at the HP turbine inlet:

11. HP turbine rotor speed:

1.2 Construction of the meridional section of the flowparts

HP turbines

The presence of the meridional shape of the flow path is necessary to determine the characteristic diameters Di in any control section of the step, and not only in sections "0" and "2". These diameters serve as the basis for performing, for example, the calculation of flow parameters at various radii of the flow path, as well as the design of control sections of the blade airfoil.

1. The width of the crown of the nozzle apparatus of the first stage:

accept kSA = 0.06

2. First stage impeller ring width:

accept kRK = 0.045

3. Width of the crown of the nozzle apparatus of the second stage:

4. Second stage impeller ring width:

5. The axial clearance between the nozzle apparatus and the impeller is usually determined from the ratio:

Axial clearance between the nozzle apparatus and the impeller of the first stage:

6. Axial clearance between the impeller of the first stage and the nozzle apparatus of the second stage:

7. Axial clearance between the nozzle apparatus and the impeller of the second stage:

8. The radial clearance between the ends of the blade feathers and the body is usually taken in the range of 0.8 ... 1.5 mm. In our case, we take:

2 . G azodynamic calculation of the turbine VD

2.1 Distributionheat drop reduction by steps

Thermodynamic parameters of the working fluid at the inlet andexiting the stairs.

1. Find the average value of the heat drop per step

.

The heat drop of the last stage is taken equal to:

Accept:

kJ/kg

Then: kJ/kg

2. Determine the degree of reactivity (for the second stage)

m

; ; .

3. Let us determine the parameters of the thermodynamic state of the gas at the inlet to the second stage

; ;

; ; .

4. Calculate the value of isentropic work in the stage when the gas expands to pressure.

Accept:

.

5. Let us determine the parameters of the thermodynamic state of the gas at the outlet of the stage under the condition of isentropic expansion from pressure to:

; .

6. Calculate the degree of gas reduction in the stage:

.

7. Determine the total pressure at the stage inlet:

,

8. We accept the angle of flow exit from the RC.

9. Gas-dynamic functions at the exit from the stage

; .

10. Static pressure downstream

.

11. Thermodynamic parameters of the flow at the outlet of the stage under the condition of isentropic expansion from pressure to

; .

12. The value of isentropic work in the stage when the gas expands from pressure to

.

2.2 Step calculation according to average at diameter at

Flow parameters behind the nozzle

1. Let us determine the isentropic velocity of gas outflow from the SA:

.

2. Determine the reduced isentropic flow velocity at the outlet of the SA:

;

3. Speed coefficient CA is accepted:

.

4. Gas-dynamic functions of the flow at the outlet of the SA:

; .

5. Determine the total pressure recovery coefficient from the table:

.

6. The angle of the flow exit from the nozzle blades:

;

Where.

7. Angle of flow deflection in an oblique section of SA:

.

8. Effective angle at the outlet of the nozzle array

.

9. The installation angle of the profile in the lattice is found according to the graph, depending on.

Accept: ;

;

.

10. Blade profile chord SA

.

11. The value of the optimal relative step is determined from the graph depending on and:

12. Optimal SA lattice spacing in the first approximation

.

13. Optimum number of SA blades

.

We accept.

14. The final value of the optimal pitch of the SA blades

.

15. The size of the throat of the SA channel

.

16. Parameters of the thermodynamic state of the gas at the outlet of the SA under the condition of isentropic expansion in the nozzle array

; .

17. Static pressure in the gap between SA and RK

.

18. Actual gas velocity at the outlet of the SA

.

19. Thermodynamic parameters of the flow at the outlet of the SA

;

; .

20. Density of the gas at the outlet of the SA

.

21. Axial and circumferential components of the absolute flow velocity at the outlet of the SA

;

.

22. Circumferential component of the relative flow velocity at the entrance to the AC

.

23. The angle of entry of the flow into the RC in relative motion

.

24. Relative flow velocity at the inlet to the AC

.

25. Thermodynamic parameters of the gas at the entrance to the AC

;

; .

26. Reduced flow velocity in relative motion

.

27. Total pressure in relative air movement

.

Flow parameters at the outlet of the RC

28. Thermodynamic flow parameters

;

;.

29. Isentropic flow velocity in relative motion

.

30. Reduced isentropic flow velocity in relative motion:

.

We accept, because relative motion is energy-isolated motion.

31. Reduced flow velocity in relative motion

Let's accept:

,

Then:

; .

32. Using the graph, we determine the total pressure recovery factor:

.

33. The angle of the flow exit from the RC in relative motion (15º<в 2 <45є)

Let's calculate:

;

.

34. Let's determine from the table the angle of flow deviation in the oblique section of the rotor blades:

.

35. Effective angle at the outlet of the DC

.

36. Let's determine from the table the angle of installation of the profile in the working blade:

Let's calculate:;

.

37. Blade profile chord RK

.

38. The value of the optimal relative lattice spacing of the Republic of Kazakhstan is determined from the tables:

.

39. Relative pitch of the RK lattice in the first approximation

.

40. Optimum number of blades RK

.

We accept.

41. The final value of the optimal pitch of the blades of the Republic of Kazakhstan

.

42. The size of the throat of the channel of the working blades

.

43. Relative speed at the exit from the Republic of Kazakhstan

44. Enthalpy and temperature of the gas at the outlet of the RC

; .

45. Density of gas at the outlet of the RC

46. Axial and circumferential components of the relative velocity at the exit from the RC

;

.

47. Circumferential component of the absolute flow velocity behind the RC

48. Absolute gas velocity behind the RK

.

49. The angle of the flow exit from the RC in absolute motion

50. Total enthalpy of gas behind the RC

.

2.3 Calculation of the effective operation of the stage, taking into account friction losses

disk and in the radial clearance

To determine the effective operation of the stage, it is necessary to take into account the energy losses associated with leakage of the working fluid into the radial clearance and friction of the stage disk against the gas. For this we define:

51. Specific work of gas on the blades of the Republic of Kazakhstan

52. Leakage losses, which depend on the design features of the stage.

In the designs of modern GTE turbines, bandages with labyrinth seals are usually used on impellers to reduce leakage. Leakage through such seals is calculated by the formula:

We accept the flow coefficient of the labyrinth seal:

The gap area is determined from the expression:

To determine the pressure first, the isentropic reduced flow velocity at the outlet to the RC at the peripheral diameter and the corresponding gas-dynamic function are found:

; .

Peripheral pressure

Seal pressure ratio

We accept the number of scallops:

Leakage loss

53. Energy loss due to friction of the stage disc on the gas

where D 1w is taken according to the drawing of the flow part

54. Total energy loss due to leakage and disk friction

55. The total enthalpy of the gas at the outlet of the RC, taking into account losses due to leakage and friction of the disk

;

56. Gas enthalpy according to static parameters at the outlet of the RC, taking into account losses due to leakage and friction of the disk

57. Total gas pressure at the outlet of the RC, taking into account losses due to leakage and disk friction

58. Actual effective operation of a stage

59. Actual efficiency steps

60. The difference between the actual effective work and the given one

which is 0.78%.

2.4 Calculation of parameters flow at different radii

turbine pressure blade wheel

At values D cf / h l< 12 по высоте лопатки возникает переменность параметров потока, определяемая влиянием центробежных сил и изменением окружной скорости. В этом случае для снижения потерь энергии лопатки необходимо выполнять закрученными. Применение закона закрутки dб/dr = 0 позволяет повысить технологическое качество лопаток. Применение закона б 1 =const позволяет выполнять сопловые венцы с б 1л =const, а закон б 2 =const позволяет улучшить технологичность лопаток соплового венца последующей ступени.

Determination of parameters for the spigot section of the blade

1. Relative bushing diameter

2. Flow exit angle in absolute motion

3. Speed ratio

4. Absolute flow rate at the outlet of the SA

5. Circumferential component of absolute speed

6. Axial component of absolute velocity

7. Isentropic velocity of gas outflow from SA

8. Thermodynamic parameters at the outlet of the SA

; ;

;

; .

9. Static pressure

.

10. Gas density

11. Circumferential speed in the sleeve section at the entrance to the RC

12. Circumferential component of the relative velocity at the entrance to the DC

13. The angle of entry of the flow into the RC in relative motion

.

14. Relative speed at the hub

15. Thermodynamic parameters at the entrance to the RC in relative motion

,

16. Total pressure at the inlet to the valve in relative motion

17. Reduced relative speed at the entrance to the RC

Parameters in peripheral section

18. Relates. peripheral section diameter

19. Angle of flow exit from SA in absolute motion

20. Speed ratio

21. Absolute speed at the exit from the SA

22. Circumferential and axial components of absolute speed

23. Isentropic velocity of gas outflow from SA

24. Thermodynamic parameters of the flow at the outlet of the SA

;

, ; .

25. Static pressure

26. Gas Density

27. The circumferential speed of rotation of the wheel on the periphery

28. Circumferential component of the relative velocity at the entrance to the RC

29. The angle of entry of the flow into the RC in relative motion

.

30. Relative flow velocity at the periphery

31. Thermodynamic parameters of the flow in relative motion at the entrance to the AC

,

32. Total pressure at the inlet to the CV in relative motion

.

33. Reduced relative velocity at the entrance to the RC

Calculation of flow parameters at the outlet of the RC

34. Relative bushing diameter

35. Flow angle in absolute motion

36. Peripheral speed in the sleeve section at the outlet of the valve

37. Static pressure at the outlet of the valve

38. Thermodynamic parameters in RK

,

39. Isentropic flow velocity at the outlet of the RC

40. Reduced isentropic speed

41. Flow velocity behind the RK in relative motion.

, Where

speed factor.

42. Thermodynamic parameters of the flow at the outlet of the RC

;

43. Gas density behind the working crown

44. Flow exit angle in relative motion

45. Circumferential and axial components of the relative flow velocity

46. Absolute speed at the output of the working crown

47. Circumferential component of absolute speed

48. Total enthalpy and temperature of the flow at the outlet of the AC

49. Gas-dynamic functions at the outlet of the RC

;

50. Total flow pressure in absolute motion at the outlet of the valve

Calculation of parameters in the peripheral section at the outlet of the RC

51. Relative diameter of the peripheral section

52. Flow angle in absolute motion

53. Peripheral speed in the peripheral section at the outlet of the RC

54. Static pressure at the outlet of the valve

55. Thermodynamic parameters during isentropic expansion in the Republic of Kazakhstan

;

56. Isentropic flow velocity at the outlet of the RC

57. Reduced isentropic speed

58. Flow velocity behind RK in relative motion

Speed ratio;

59. Thermodynamic parameters of the flow at the outlet of the RC

;

60. Gas density behind the working crown

61. Flow outlet angle in relative motion

62. Circumferential and axial components of the relative flow velocity

63. Absolute exit speed from RK

64. Circumferential component of absolute speed

65. Total enthalpy and temperature of the flow at the outlet of the AC

66. Gas-dynamic functions at the outlet of the RC

;

67. Total flow pressure in absolute motion at the outlet of the valve

3. Profiling of the impeller blade

Table 2. - Initial data for profiling of RV blades

Initial parameter and calculation formula	Dimension	Control sections

D (according to the drawing of the flow part of the stage)

Table 3. - Calculated values for the profiling of the blades RK

Value	Average diameter	Periphery

Conclusion

In the course work, the flow path of the high-pressure turbine was calculated and built, a kinematic calculation of the second stage of the high-pressure turbine at an average diameter was made, the calculation of effective operation, taking into account friction losses of the disk and in the radial clearance, the calculation of the parameters for the height of the blade with the swirl law b = const with the construction of triangles of speeds. Profiling of the impeller blade in three sections was performed.

List of sources used

1. Thermogasdynamic design of axial turbines for aircraft gas turbine engines using p-i-T functions: Proc. allowance / N.T. Tikhonov, N.F. Musatkin, V.N. Matveev, V.S. Kuzmichev; Samar. state aerospace un-t. - Samara, 2000. - 92. p.

2. Mamaev B.I., Musatkin N.F., Aronov B.M. Gas-dynamic design of axial turbines for aircraft gas turbine engines: Textbook. - Kuibyshev: KuAI, 1984 - 70 p.

3. Design calculation of the main parameters of aircraft GTE turbochargers: Proc. allowance / V.S. Kuzmichev, A.A. Trofimov; KuAI. - Kuibyshev, 1990. - 72 p.

4. Thermogasdynamic calculation of gas turbine power plants. / Dorofeev V.M., Maslov V.G., Pervyshin N.V., Svatenko S.A., Fishbein B.D. - M., "Engineering", 1973 - 144 p.

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Device

The first circuit contains high and low pressure compressors, a combustion chamber, high and low pressure turbines and a nozzle. The second circuit consists of a guide vane and a nozzle. This design is basic, but some deviations are possible, for example, the flows of the inner and outer circuits can mix and exit through a common nozzle, or the engine can be equipped with an afterburner.

Now briefly about each constituent element of the turbofan engine. A high pressure compressor (HPC) is a shaft on which movable and fixed blades are fixed, forming a stage. Movable blades during rotation capture the air flow, compress it and direct it into the housing. The air enters the fixed blades, slows down and is additionally compressed, which increases its pressure and gives it an axial motion vector. There are several such stages in the compressor, and the compression ratio of the engine directly depends on their number. The same design is for the low pressure compressor (LPC), which is located in front of the HPC. The difference between them is only in size: LPC blades have a larger diameter, covering the cross section of both the primary and secondary circuits, and a smaller number of steps (from 1 to 5).

In the combustion chamber, compressed and heated air is mixed with fuel, which is injected by injectors, and the resulting fuel charge ignites and burns, forming gases with a large amount of energy. The combustion chamber can be one, annular, or made of several pipes.

The turbine in its design resembles an axial compressor: the same fixed and movable blades on the shaft, only their sequence is changed. First, the expanded gases fall on the fixed blades, which equalize their movement, and then on the movable ones, which rotate the turbine shaft. There are two turbines in the turbofan engine: one drives the high-pressure compressor, and the second drives the low-pressure compressor. They work independently and are not mechanically connected to each other. The LPC drive shaft is usually located inside the HPC drive shaft.

A nozzle is a converging pipe through which exhaust gases exit in the form of a jet stream. Usually, each circuit has its own nozzle, but it also happens that the jet streams at the outlet enter a common mixing chamber.

The outer, or second, circuit is a hollow annular structure with a guide vane, through which air passes, pre-compressed by a low-pressure compressor, bypassing the combustion chamber and turbines. This air flow, falling on the fixed blades of the guide vane, is leveled and moves to the nozzle, creating additional thrust due to compression of the LPC alone without burning fuel.

The afterburner is a pipe placed between the low pressure turbine and the nozzle. Inside it has swirlers and fuel injectors with igniters. The afterburner makes it possible to create additional thrust by burning fuel not in the combustion chamber, but at the turbine outlet. Exhaust gases after passing LPT and HPT have a high temperature and pressure, as well as a significant amount of unburned oxygen coming from the secondary circuit. Through the nozzles installed in the chamber, fuel is supplied, which mixes with gases and ignites. As a result, the output thrust sometimes doubles, however, the fuel consumption also increases. Turbofan engines equipped with an afterburner are easily recognizable by the flame that escapes from their nozzle during flight or upon launch.

cross section of the afterburner, swirlers are visible in the figure.

The most important parameter of a turbofan engine is the bypass ratio (k) - the ratio of the amount of air that has passed through the second circuit to the amount of air that has passed through the first. The higher this figure, the more economical the engine will be. Depending on the degree of bypass, the main types of bypass turbojet engines can be distinguished. If its value is<2, это обычный ТРДД, если же к>2, then such engines are called turbofan engines (TVRD). There are also turboprop-fan engines, in which the value reaches 50 or even more.

Depending on the type of exhaust gas discharge, turbofan engines are distinguished without mixing flows and with it. In the first case, each circuit has its own nozzle, in the second, the gases at the outlet enter the common mixing chamber and only then go outside, forming a jet thrust. Mixed-flow engines, which are installed on supersonic aircraft, can be equipped with an afterburner, which allows you to increase thrust even at supersonic speeds, when secondary thrust plays little role.

Principle of operation

The principle of operation of the TVRD is as follows. The air flow is captured by the fan and, partially compressed, is directed in two directions: to the first circuit to the compressor and to the second circuit to the fixed blades. In this case, the fan does not play the role of a screw that creates thrust, but a low-pressure compressor that increases the amount of air passing through the engine. In the primary circuit, the flow is compressed and heated as it passes through the high pressure compressor and enters the combustion chamber. Here it mixes with the injected fuel and ignites, resulting in the formation of gases with a large supply of energy. The flow of expanding hot gases is directed to the high-pressure turbine and rotates its blades. This turbine rotates the high-pressure compressor, which is mounted with it on the same shaft. Next, the gases rotate the low-pressure turbine, which drives the fan, after which they enter the nozzle and break out, creating jet thrust.

At the same time, in the second circuit, the air flow captured and compressed by the fan hits the fixed blades, which straighten the direction of its movement so that it moves in the axial direction. In this case, the air is additionally compressed in the second circuit and goes outside, creating additional traction. Also, the thrust is affected by the combustion of oxygen in the secondary air in the afterburner.

Application

The scope of application of bypass turbojet engines is very wide. They were able to cover almost the entire aviation, displacing the turbojet and theater engines. The main disadvantage of jet engines - their inefficiency - was partially overcome, so that now most civilian and almost all military aircraft are equipped with turbofan engines. For military aviation, where compactness, power and lightness of engines are important, turbofan engines with a low bypass ratio (to<1) и форсажными камерами. На пассажирских и грузовых самолетах устанавливаются ТРДД со степенью двухконтурности к>2, which saves a lot of fuel at subsonic speeds and reduces the cost of flights.

Low bypass bypass turbojet engines in military aircraft.

SU-35 with 2 AL-41F1S engines installed on it

Advantages and disadvantages

Bypass turbojet engines have a huge advantage over turbojet engines in the form of a significant reduction in fuel consumption without loss of power. But at the same time, their design is more complex, and the weight is much greater. It is clear that the greater the bypass ratio, the more economical the motor, but this value can be increased in only one way - by increasing the diameter of the second circuit, which will make it possible to pass more air through it. This is the main disadvantage of the turbofan. It is enough to look at some turbojet engines installed on large civil aircraft to understand how they make the overall structure heavier. The diameter of their second circuit can reach several meters, and in order to save materials and reduce their weight, it is shorter than the first circuit. Another disadvantage of large structures is the high drag during flight, which to some extent reduces the flight speed. The use of turbofan engines in order to save fuel is justified at subsonic speeds, when the sound barrier is overcome, the secondary jet thrust becomes ineffective.

Various designs and the use of additional structural elements in each individual case makes it possible to obtain the desired version of the turbofan engine. If economy is important, turbofan engines with a large diameter and a high bypass ratio are installed. If you need a compact and powerful engine, conventional turbofan engines with or without an afterburner are used. The main thing here is to find a compromise and understand what priorities a particular model should have. Military fighters and bombers cannot be equipped with engines with a three-meter diameter, and they don’t need it, because in their case the priority is not so much economy as speed and maneuverability. Here, turbofan engines with afterburners (TRDDF) are also more often used to increase traction at supersonic speeds or during launch. And for civil aviation, where the aircraft themselves are large, large and heavy engines with a high bypass ratio are quite acceptable.

For the first time an aircraft with a turbojet engine ( TRD) took to the air in 1939. Since then, the design of aircraft engines has been improved, various types have appeared, but the principle of operation for all of them is approximately the same. To understand why an aircraft with such a large mass can take to the air so easily, you need to understand how an aircraft engine works. A turbojet engine propels an aircraft using jet propulsion. In turn, jet thrust is the recoil force of the gas jet that flies out of the nozzle. That is, it turns out that the turbojet installation pushes the plane and all the people in the cabin with the help of a gas jet. The jet stream, flying out of the nozzle, is repelled from the air and thus sets the aircraft in motion.

Turbofan engine device

Design

The device of the aircraft engine is quite complicated. The operating temperature in such installations reaches 1000 degrees or more. Accordingly, all the parts that make up the engine are made of materials that are resistant to high temperatures and fire. Due to the complexity of the device, there is a whole field of science about turbojet engines.

TRD consists of several main elements:

fan;
compressor;
the combustion chamber;
turbine;
nozzle.

A fan is installed in front of the turbine. With its help, air is drawn into the unit from the outside. In such installations, fans with a large number of blades of a certain shape are used. The size and shape of the blades provide the most efficient and fast air supply to the turbine. They are made from titanium. In addition to the main function (drawing in air), the fan solves another important task: it is used to pump air between the elements of the turbojet engine and its shell. Due to this pumping, the system is cooled and the destruction of the combustion chamber is prevented.

A high power compressor is located near the fan. With its help, air enters the combustion chamber under high pressure. In the chamber, air is mixed with fuel. The resulting mixture is ignited. After ignition, the mixture and all adjacent elements of the installation are heated. The combustion chamber is most often made of ceramic. This is due to the fact that the temperature inside the chamber reaches 2000 degrees or more. And ceramics is characterized by resistance to high temperatures. After ignition, the mixture enters the turbine.

View of the aircraft engine from the outside

A turbine is a device consisting of a large number of blades. The flow of the mixture exerts pressure on the blades, thereby setting the turbine in motion. The turbine, due to this rotation, causes the shaft on which the fan is mounted to rotate. It turns out a closed system, which for the operation of the engine requires only the supply of air and the presence of fuel.

Next, the mixture enters the nozzle. This is the final stage of the 1st engine cycle. This is where the jet stream is formed. This is how an airplane engine works. The fan forces cold air into the nozzle, preventing it from being destroyed by an excessively hot mixture. The cold air flow prevents the nozzle collar from melting.

Various nozzles can be installed in aircraft engines. The most perfect are considered mobile. The movable nozzle is able to expand and contract, as well as adjust the angle, setting the correct direction of the jet stream. Aircraft with such engines are characterized by excellent maneuverability.

Types of engines

Aircraft engines are of various types:

classic;
turboprop;
turbofan;
straight-through.

Classic installations work according to the principle described above. Such engines are installed on aircraft of various modifications. Turboprop function somewhat differently. In them, the gas turbine has no mechanical connection with the transmission. These installations drive the aircraft with the help of jet thrust only partially. This type of installation uses the main part of the energy of the hot mixture to drive the propeller through the gearbox. In such an installation, instead of one, there are 2 turbines. One of them drives the compressor, and the second - the screw. Unlike classic turbojet, screw installations are more economical. But they do not allow aircraft to develop high speeds. They are installed on low-speed aircraft. TRDs allow you to develop much greater speed during the flight.

Turbofans engines are combined units that combine elements of turbojet and turboprop engines. They differ from the classic ones in the large size of the fan blades. Both the fan and propeller operate at subsonic speeds. The speed of air movement is reduced due to the presence of a special fairing in which the fan is placed. Such engines consume fuel more economically than classic ones. In addition, they are characterized by higher efficiency. Most often they are installed on liners and large-capacity aircraft.

Aircraft engine size relative to human height

Direct-flow air-jet installations do not involve the use of moving elements. Air is drawn in naturally thanks to a fairing mounted on the inlet. After the intake of air, the engine works similarly to the classic one.

Some aircraft fly on turboprop engines, which are much simpler than turbojet engines. Therefore, many people have a question: why use more complex installations, if you can limit yourself to a screw one? The answer is simple: turbojet engines are superior in power to screw engines. They are ten times more powerful. Accordingly, the turbojet engine produces much more thrust. This makes it possible to lift large aircraft into the air and fly at high speed.

In contact with

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-unwinding 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

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 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 itself. 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 swirling 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.

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Introduction

1. Design calculation of the main parameters of the turbine of high

pressure

1.1 Calculation geometric and regime parameters HP turbines

1. The width of the crown of the nozzle apparatus of the first stage:

accept kSA = 0.06

2. First stage impeller ring width:

accept kRK = 0.045

3. Width of the crown of the nozzle apparatus of the second stage:

4. Second stage impeller ring width:

5. The axial clearance between the nozzle apparatus and the impeller is usually determined from the ratio:

Axial clearance between the nozzle apparatus and the impeller of the first stage:

6. Axial clearance between the impeller of the first stage and the nozzle apparatus of the second stage:

7. Axial clearance between the nozzle apparatus and the impeller of the second stage:

8. The radial clearance between the ends of the blade feathers and the body is usually taken in the range of 0.8 ... 1.5 mm. In our case, we take:

2 . G azodynamic calculation of the turbine VD

2.1 Distributionheat drop reduction by steps

Thermodynamic parameters of the working fluid at the inlet andexiting the stairs.

1. Find the average value of the heat drop per step

.

The heat drop of the last stage is taken equal to:

Accept:

kJ/kg

Then: kJ/kg

2. Determine the degree of reactivity (for the second stage)

m

; ; .

3. Let us determine the parameters of the thermodynamic state of the gas at the inlet to the second stage

; ;

; ; .

4. Calculate the value of isentropic work in the stage when the gas expands to pressure.

Accept:

.

5. Let us determine the parameters of the thermodynamic state of the gas at the outlet of the stage under the condition of isentropic expansion from pressure to:

; .

6. Calculate the degree of gas reduction in the stage:

.

7. Determine the total pressure at the stage inlet:

,

8. We accept the angle of flow exit from the RC.

9. Gas-dynamic functions at the exit from the stage

; .

10. Static pressure downstream

.

11. Thermodynamic parameters of the flow at the outlet of the stage under the condition of isentropic expansion from pressure to

; .

12. The value of isentropic work in the stage when the gas expands from pressure to

.

2.2 Step calculation according to average at diameter at

Flow parameters behind the nozzle

1. Let us determine the isentropic velocity of gas outflow from the SA:

.

2. Determine the reduced isentropic flow velocity at the outlet of the SA:

;

3. Speed ​​coefficient CA is accepted:

.

4. Gas-dynamic functions of the flow at the outlet of the SA:

; .

5. Determine the total pressure recovery coefficient from the table:

.

6. The angle of the flow exit from the nozzle blades:

;

Where.

7. Angle of flow deflection in an oblique section of SA:

.

8. Effective angle at the outlet of the nozzle array

.

9. The installation angle of the profile in the lattice is found according to the graph, depending on.

3. Speed coefficient CA is accepted: