RU2525385C1 - Gas turbine engine - Google Patents

Abstract

FIELD: machine building.

SUBSTANCE: turbine is fitted by a cooled nozzle diaphragm with its blades along the airfoil from the leading edge having the first, second, third and fourth inner cavities connected to a flow passage through the holes in the blade airfoil, and a bypass device. The combustion chamber is fitted by annular space between the inner, outer case and the annular flame tube with heads. The input of the annular flame tube head is connected to the flow passage of the compressor sequentially from the compressor via a circular segment of the vaned diffuser with the output of the latter being connected to the input of the air line - channel pipe branch with its output being connected to the input of the third inner cavity of the cooled nozzle diaphragm's blade with one of its outputs being connected to the input of the flame tube head. The third inner cavity is fitted by two more outputs. One of the outputs is connected to the input of the first inner cavity of the blade via the annular space of the combustion chamber and the annular cavity-receiver. The second input is connected to the fourth inner cavity of the nozzle diaphragm blade via the hole in a partition wall. The nozzle diaphragm has at least one or more blades whose third inner cavity is fitted by the fourth output which connects it to the second inner cavity via the hole in the partition wall. The bypass device is installed in the said blades and is kinematically coupled with a valve set at the input of the fuel nozzle of the head connected to the said blade. The second cavity of the said blades is connected to the second cavity of the blade which is not equipped with the bypass device.

EFFECT: invention provides for efficient operation of the combustion chamber of a gas turbine engine and cooling system of a high temperature gas turbine in different operation modes.

6 cl, 11 dwg

Description

The invention relates to the field of gas turbine engineering and can be used for gas turbine plants for various purposes.

A technical solution is known according to (1), where the gas turbine engine contains a diffuser 5 (Fig. 1), a countercurrent-type vortex combustion chamber with an annular flame tube 3, in the annular channel of which, between the annular visor 13, the outer annular sleeve 4 and the casing 18, is front-mounted the device is a swirl 6 with rotary blades 21. Fuel is supplied to the combustion chamber through jet fuel nozzles 28 located between the blades 21 of the swirl 6. The flow and swirl of the air-fuel flow from the front-end device changes as The orbit of the blades 21 of the swirler 6.

The disadvantage of this technical solution is the low efficiency of the gas turbine engine due to large losses of air pressure in the combustion chamber, determined by the pressure drop in front of and behind the swirl of the front device. Losses increase significantly with the rotation of the blades of the swirl to increase flow swirl.

The operation of a diffuser with a large opening angle, in the absence of special devices for organizing air flows, is also economically inefficient (2). In figure 1, this angle is 50 °, which is significantly greater than the optimal value of this angle for the annular diffuser.

It is also irrational to use the air heated in the combustion chamber, from the openings of the inner casing 2, in the turbine cooling system.

Also known is a gas turbine engine (3) comprising a compressor 28 (FIGS. 2, 3), a multi-sector diffuser 36 with a frontal device including a fuel supply channel 34 with jet openings 50 and a flame stabilizer 90. There is an annular flame tube 12, a turbine with a cooled nozzle blade 42, the cooling air to which is supplied via a pneumatic pipe 110, 106 through a heat exchanger 100 from a cavity 102 connected to a peripheral channel of the diffuser flow part 32. A variant is presented (Fig. 4) when part of the air due to the compressor enters the annulus 24, 26 of the combustion chamber through peripheral 31 and root 33 sections of the flow part of the multi-sector diffuser, and the cooling system of the nozzle blade 42 is connected to the annulus of the combustion chamber.

The disadvantage of this technical solution is the non-optimal operation of the combustion chamber at various operating modes of the gas turbine engine due to the lack of structural elements that provide a change in the passage area for air to the front-end device.

The use of low-pressure air from the heat exchanger does not allow for efficient cooling of the inlet edge of the nozzle blade by its perforation.

Also known is a gas turbine engine with a variable geometry of the combustion chamber (4) containing a diffuser 58 (Fig. 5), an annular combustion chamber with a countercurrent flame tube having a front device with swirls 84, 86, and bypass windows 64 in the area of the front device and 92 in diffuser area. There are two annular bands 104, 96 with springs 132 and 134 connected to a control mechanism including a drive, a system of levers, springs, rings and hinges, so that the bands 104, 96 can synchronously move along the flame tube - the axis of the engine. In the tape 96, covering the pipe 60 of the front-end device, there are holes 98. At the maximum mode of the gas turbine engine, the tape 96 is in a position where its holes 98 coincide with the windows 64 of the entrance to the pipe 60. The tape 104 is located so that it covers the bypass windows 92 In this case, the air from the diffuser 58 enters the combustion zone of the flame tube through the swirls 84, 86, passing holes 98, windows 64, distributed in the pipe 60. In the combustion zone, the optimum air-fuel ratio is provided. When reducing the engine operating mode by reducing the amount of fuel supplied to the engine, the optimal ratio of air to fuel in the combustion zone is ensured by blocking the passage section of windows 64. To do this, with the help of the control mechanism and spring 134, the annular tape is shifted to close the windows 64. At the same time, with the help of spring 132, the annular strip 104 is shifted to the position when the bypass ports 92 are opened for the passage of air into the flame tube. Thus, in the well-known gas turbine engine, an attempt is made to provide air bypass in partial modes of its operation, which should ensure optimal combustion and reduce air pressure losses in the combustion chamber.

However, this technical solution does not fully meet the stated goal. The location of the windows 92 and the end face 106 of the diffuser 58 does not allow for unhindered flow of air into the bypass windows 92 when they are opened. The air leaving the diffuser at high speed reduces the static pressure behind the end face 106 in front of the window 92, which does not allow to reduce the resistance of the combustion chamber. In addition, the partial closure of the bypass holes 64 leads to a decrease in pressure in front of the swirls 84, 86, which reduces their efficiency.

In the prototype under consideration, there is also no effective system for the cooling of the inlet edge of the nozzle blade of the turbine.

The claimed invention is aimed at solving a number of problems: increasing the efficiency of a gas turbine engine by reducing air pressure losses; improving the environmental performance of the engine while ensuring the stability of starting and operating characteristics; the implementation of the optimal and cost-effective use of the coolant of the working fluid in the cooling system of high-temperature turbines and the combustion chamber.

These problems are solved in a gas turbine engine containing a compressor, scoop diffusers, duct pipes, an annular cavity cavity, a looped combustion chamber with an annulus between the inner, outer casings and a vortex-type annular flame tube with front-end devices, a turbine with a cooled nozzle apparatus, the blades of which along the profile of the pen from the inlet edge there are first, second, third and fourth internal cavities connected to the flowing part through openings in the shoulder blades, and a bypass device , and in accordance with the invention, the input of the frontal device of the annular flame tube is connected to the compressor flow path in series from the compressor, through the annular segment of the blade diffuser, the output of which is connected to the input of the channel pipe, the output of which is connected to the entrance to the third internal cavity of the cooled blade of the nozzle apparatus, one of the exits from which is connected to the entrance to the frontal device of the flame tube.

In addition, there are two more exits from the third internal cavity of the blade of the nozzle apparatus. One outlet is located in the peripheral shelf of the scapula. It connects the third internal cavity of the blade with the annular space of the combustion chamber, which is connected through the windows in the internal housing of the combustion chamber, the annular cavity cavity and the hole in the internal shelf of the blade to the entrance to its first internal cavity. An annular cavity-receiver is located between the turbine and the compressor and is limited by the internal housing of the combustion chamber with a nozzle device mounted on it and above the rotor casing.

The second exit from the third internal cavity of the blade is a window in the dividing wall with the fourth internal cavity of the blade of the nozzle apparatus.

In addition, there are at least one or more blades in the nozzle apparatus, in which the third internal cavity has a fourth outlet connecting it through a window in the dividing wall with the second internal cavity. Moreover, in these blades there is also a bypass device, and their second cavity through holes in the peripheral shelf is connected to the second cavity of the blade, which does not have a bypass device. Such blades are located next to the blades that do not have a bypass device, in which the wall between the second and third cavity is continuous - does not have a window. Bypass devices have a kinematic connection with valves at the entrance to the fuel nozzles. Moreover, the bypass device is connected with the nozzle valve located in the front device connected to the blade of the corresponding bypass device. The bypass device is also connected to the control mechanism. One control mechanism may be associated with one or more bypass devices and their corresponding nozzle valves.

Still, the internal section of the channel pipe — the pneumatic line connecting the blade diffuser with the third internal cavity of the cooled blade of the nozzle apparatus, is made by a diffuser with an equivalent opening angle of less than 12 °.

The bypass device itself, located inside the cooled nozzle blade, has a rotary valve with an axis of rotation, resting on the nozzle connecting the third internal cavity of the blade with the entrance to the front device of the flame tube, from the back and trough of the blade. The rotary valve is made profiled in the form of a flap. The seat for this valve is in the “open” position - the wall of the window between the second and third internal cavities of the scapula, and in the “closed” position - the end of the pipe connecting the third internal cavity of the scapula with the entrance to the frontal device of the flame tube.

And even at the height of the feather of the nozzle blade, holes are made connecting the first, second and fourth internal cavities of the blade with the turbine flow part, the holes of the first cavity being made at the input edge of the blade, the second at the trough and back, and the fourth at the output edge.

The implementation of the gas turbine engine in the claimed manner, including the connection of the front-end device with the compressor through the nozzle blade, the blade and channel pipe, provides minimal pressure loss in the gas generator and effective convective cooling of the nozzle blade. Special profiling of the channel nozzle provides the maximum degree of conversion of the pressure head to pressure. The optimal and cost-effective use of the coolant of the working fluid in the cooling system of the combustion chamber and the nozzle apparatus of high-temperature turbines can be achieved by connecting the holes in the inlet edge of the blade with the compressor through the convective cooling system of the blade — the third internal cavity and the annular space of the combustion chamber. And finally, the presence of bypass devices and associated fuel valves allows the environmentally friendly operation of the combustion chamber in its entire range of operating modes, without additional loss of air pressure.

The invention is illustrated by drawings, which depict:

6 is a longitudinal section of a gas generator of a gas turbine engine;

Fig.7 - element A in Fig.6, the bypass device in the "open" position;

Fig.8 is a section B-B in Fig.7;

Fig.9 - element a in Fig.6, the bypass device in the "closed" position;

figure 10 - element B in figure 7, the bypass device in the "open" position;

11 - element A in Fig.6 cross section of Fig.7, rotated in the circumferential direction by one blade (360 ° / N, where N is the number of blades in the nozzle apparatus).

The device - a gas turbine engine contains a compressor 1 (Fig.6) with working 2 and straightening blades 3, a diffuser 4 (Fig.7) with a straightening apparatus 5, scapular diffuser 6, duct pipes 7, loop combustion chamber 8 (Fig.6) with the annular space 9 between the inner 10 and outer 11 bodies and the vortex-type annular flame tube 12 with front devices 13 having an input 14 (Fig. 8) and an output 15. The rotor 16 (Fig. 7) connects the compressor to the turbine 17, which has a cooled nozzle apparatus 18 (Fig. 9) with a uniformly alternating hollow blade mi 19, 20 (Figs. 8, 9, 11). In addition, the nozzle vanes has a first 21 (FIG. 10), a second 22, a third 23 and a fourth 24 inner cavities separated by walls 25, 26 (FIG. 11), and 27 (FIG. 10), in which there are windows 28 (FIG. .9) and 29 (Fig. 10). The blade feather has an input 30 (figure 10), output 31 edges and holes 32 (figure 10), 33 (figure 9) and 34. In the inner 35 (figure 10) and peripheral 36 shelf of the cooled blade, holes 37 are made , 38, 39, 40 and pipe 41 (Fig. 9). There is a segmented cavity 42 (Fig. 10), an annular cavity-receiver 43 (Fig. 9), a casing 44, bypass devices 45 (Fig. 7) with a valve 46 on the axis 47, with springs 48 and with control mechanisms 49, fuel nozzles 50. The air flow directions 51 (FIG. 8), 52, 53 (FIG. 7), 54 (FIG. 9), 55, 56, 57, 58, 59 (FIG. 11), 60 and 61 are indicated. The end face is shown. 62 (Fig. 10), windows 63, 64, 65, inlet 66 (Fig. 8) and outlet 67 of the duct diffuser pipe, valve 68 (Fig. 8), fuel manifold 69 and the position of the bypass device “open” in Fig. 7 and "Closed" in Fig.9.

The input 14 of the front device 13 of the annular flame tube 12 is connected to the flow part of the compressor 1, in series from the compressor 1: through a straightening device 5, a diffuser 4; annular segment of the blade diffuser 6; the outlet from which is further connected to the inlet 66 of the pneumatic conduit - channel pipe 7, the outlet 67 of which, through a window 65 in the inner housing 10 is connected through an opening 40 in the inner shelf 35 to the third inner cavity 23 of the cooled blade 19, 20 of the nozzle apparatus 18. Finally, the pipe 41 on the peripheral shelf 36 of the blade 19, 20 connects the third internal cavity 23 with the input 14 of the front device 13. In the pipe 41 of the blade 20 of the nozzle apparatus 18 there is a valve 46 of the bypass device 45. The blades 20 are installed next to the blades 19, which do not have transfer devices 45. The second internal cavities 22 of pairwise adjacent vanes 19 and 20 are interconnected via openings 38 in their peripheral shelves 36 and a segment cavity 42. The first internal cavity 21, in the region of the inlet edge 30 of the vanes 19, 20 is connected through openings 37 in the inner shelf 35 and windows 63 in the inner housing 10 of the combustion chamber 8 with the annular cavity receiver 43. The annular cavity receiver 43 is located between the compressor 1 and the turbine 17. At the periphery, the annular cavity 43 is limited to the annular inner housing 10 of the combustion chamber 8 s a nozzle device 18 fixed on it, inside - rings: a scapular diffuser 6 and above a rotary casing 16. An annular cavity receiver 43 is connected through windows 61 in the inner case 10 of the loopback combustion chamber 8 to the annular space 9. The third is also connected to the annular space 9 the inner cavity 23 of the blade 19, 20 through the hole 39 in the peripheral shelf 36.

The internal passage section of the pipe 7, from the input 66 to the output 67 is made by a diffuser with an equivalent opening angle of less than 12 °.

The valve 46 of the bypass device 45, located inside the cooled nozzle vane 20, has a transverse axis of rotation 47, resting on the nozzle 41 from the back and trough of the vane 20. The rotary valve 46 itself is shaped in the form of a flap and through a mechanism consisting of levers and springs 48, connected to a control mechanism 49. Behind the rotary valve 46 in the front device 13, a fuel nozzle 50 is installed with a valve 68 between the nozzle 50 and the fuel manifold 69. At a similar location along the length of the front device 13 tsya fuel injector 50 and the scapula 19. These nozzles 50 are connected to the fuel manifold 69 directly. The number of control mechanisms 49 is less than or equal to the number of blades 20 with a rotary valve 46. If the number of control mechanisms 49 does not correspond to the number of blades 20, then the rotary valves 46 of several blades 20 with the help of additional levers and springs 48 are combined into groups, connecting to a common one the control mechanism 49. In this case, the control mechanism 49 is simultaneously connected via levers and springs 48 to a valve 68 mounted between the fuel manifold 69 and the fuel injector 50 of the corresponding front device 13, behind the blade 20 or a group of blades 20, a rotary valve 46 which is connected with the same control mechanism 49.

In the wall 26 of the blades 20 there is a window 28 connecting the second internal cavity 22 with the third 23. In the wall 27 of the blades 19, 20 there is a window 29 connecting the third 23 internal cavity with the fourth 24. The wall of the window 28 is also a seat for the bypass rotary valve 46 open position. In the closed position, the seat for the valve 46 is the end face 62 of the pipe 41.

The blade blade 19, 20 has openings. So, at the inlet edge 30, a series of openings 32 are provided for defensive cooling. Holes 32 connect the first internal cavity 21 to the flow part of the turbine 17. On the back and trough of the blade 19, 20, rows of holes 33 are made connecting the second internal cavity 22 with the flow part of the turbine 17. And a number of holes are made in the region of the outlet edge 31 of the blade 19, 20 34, connecting the fourth internal cavity 24 with the flow part of the turbine 17.

During the operation of the gas turbine engine in the compressor 1 on the working blades 2 and the straightening blades 3, air is compressed - the working fluid of the gas turbine engine. The turbine 17 rotates the compressor 1 through the rotor 16. The high-speed, swirling air stream 51 from the compressor 1 enters the diffuser 4, where it rotates in the axial direction in the straightening device 5 and in the blade diffuser 6 and reduces the speed, converting the pressure head to pressure. Then the air through the inlet 66 enters the channel pipe 7, where it continues to slow down, increasing the pressure. Further, the air through the outlet 67 of the pipe 7 and the hole 40 in the inner shelf 35 passes into the third internal cavity 23 of the cooled blade 19, 20 of the nozzle apparatus 18 of the turbine 17. Here, the main air stream is divided into several streams. Part of the stream 56 spreads through the streams 60 and 61. The air stream 60 through the opening 39 in the peripheral shelf 36 flows into the annular space 9 of the combustion chamber 8. The stream 61 through the window 29 in the wall 27 enters the fourth internal cavity 24 of the blades 19, 20, from where through the holes 34 in the outlet edge 31 goes into the flow part of the turbine 17. Air 60 passing into the annular space 9, flows from it through the windows 64 by a stream 55 into the annular cavity receiver 43, from where by the stream 58 through the windows 63 and openings 37 into the first internal cavity 21 vanes 19, 20, in the region of the entrance chrome ki 30. From this cavity 21, air exits through openings 32 at the inlet edge 30 to the flow part of the nozzle apparatus 18 of the turbine 17. This air forms a defensive cooling of the blade 19, 20, including its inlet edge 30. When the gas turbine engine operates in the region of the nominal mode, the valve 46 bypass device 45 inside the blades 20 is in the "open" position (Fig.7). In this position, the valve 46 closes the window 28 in the wall 26. In this case, part 52 of the air flow entering the third cavity 23 passes through the pipe 41 to the inlet 14 of the front device 13. Here, the air is mixed with the fuel entering the front device 13 through the nozzle 50. The prepared air-fuel mixture exits through outlet 15 into the annular flame tube 12, where it ignites and burns, forming hot gases that enter the turbine duct 17 through the nozzle apparatus 18. Air 52 also flows from the third cavity 23 of the blade 19 to the front device 13. At the same time, the amount of air 52 from the blade 19, 20 and fuel from the nozzle 50 entering the front device 13, after the blade 19, 20, provides air-fuel mixture at the outlet 15 in a concentration realizing combustion in the combustion chamber 8 with a low content harmful substances.

When reducing the power of the gas turbine engine by reducing fuel consumption in the fuel manifold, the output 15 changes the concentration of the air-fuel mixture. To ensure the preservation of the optimal concentration of the air-fuel mixture in the combustion zone of the combustion chamber 8, the following actions are performed. One of the control mechanisms 49 through levers and springs 48 simultaneously turns the valve 46 and acts on the corresponding fuel valve 68, cutting off the fuel supply to the nozzle 50 of the front device 13 behind the rotary valve 46. At the same time, the rotary valve 46 sits on the end face 62 of the pipe 41, blocking the air 52 exit from this pipe 41 to the front device 13 (Fig.9). Another window 28 opens in the wall 26 of the blade 20. Air through this window 28 passes into the second internal cavity 22 of the blade 20, which is divided here into two streams 57, 58. The stream 57 exits through the opening 38 of the blade 20 and the segment cavity 42, from where the hole 38 of the adjacent blade 19 by the flow 59 enters its second internal cavity 22, from where it goes through the holes 33 into the flow part of the turbine 17. The air stream 58 also through the holes 33 of the blade 20 enters the flow part of the turbine 17.

These actions by redistributing air and fuel in the combustion chamber 8 provide at the outlet 15 behind the blades 19 and the blades 20 with valves in the "open" position the air-fuel mixture in a concentration that implements combustion in the combustion chamber 8 with a low content of harmful substances. In this case, the total resistance of the air path of the gas generator does not change.

With a further decrease in the power of the gas turbine engine by reducing the fuel consumption in the fuel manifold, the following control mechanism 49 is actuated beyond the boundaries of the optimal concentration of the air-fuel mixture at the outlet 15. The above-described algorithm for the redistribution of air and fuel associated with the blocked nozzle 50 and the blade 20 or groups of nozzles 50 and the blades 20 are repeated.

Thus, in all operating modes of the gas turbine engine, the combustion chamber 8 is ensured that the combustion chamber 8 is operating in an optimal mode, with low emissions of harmful substances associated with underburning or overheating. Switching of all control mechanisms 49 does not lead to a noticeable change in pressure losses in the gas generator, providing high efficiency of the gas turbine engine in all operating modes.

High engine efficiency is also ensured by low pressure losses behind the compressor, in channels made by a diffuser with an equivalent opening angle of less than 12 °.

The inventive device provides an economical use of the cooling resource of cooling air. In the inventive gas turbine engine, air 60, due to the compressor with a low temperature, first convectively cools the blade 19, 20 and the heat pipe 12. Partially using the coolant, air 60 is further used in the cooling cooling of the blade 1, where its elevated temperature does not significantly affect the efficiency cooling.

Special profiling of the rotary valve 46 provides in the "open" and "closed" position flow around it with an air stream with low pressure loss.

Information sources

1. Copyright certificate of the USSR 1726917, F23R 3/04, publ. 04/15/1992.

2. RF patent No. 2469214, F04D 29/44, publ. 12/10/2012

3. US patent 5791148, F01C 1/00, publ. 08/11/1998.

4. US patent 4497170, F02C 3/14, publ. 02/05/1985.

5. US patent 4532762, F23R 3/26, publ. 08/06/1985.

6. US patent 4918926, F01C 1/00, publ. 04/24/1990.

7. US patent 5077967, F01C 3/00, publ. 06/07/1992.

8. US patent 5335501, F01C 1/00, publ. 08/09/1994.

9. US patent 5619855, F01C 7/08, publ. 04/15/1997.

10. US patent 3691761, F01C 9/14, publ. 09/19/1972.

11. US patent 4527386, F01C 3/14, publ. 07/09/1985.

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17. RF patent 2117874, F23R 3/38, publ. 08/20/1998.

Claims (6)

1. A gas turbine engine containing a compressor, blade diffusers, a duct pipe, an annular receiver cavity, a combustion chamber with an annulus between the inner, outer casing and an annular flame tube with front-end devices, a turbine with a cooled nozzle apparatus, the blades of which along the profile of the pen from the inlet the edges have first, second, third and fourth internal cavities connected to the flowing part through openings in the shoulder blades, and a bypass device, characterized in that the front-end device input The properties of the annular flame tube are connected to the compressor flow path in series from the compressor through the annular segment of the blade diffuser, the output of which is connected to the inlet of the pneumatic pipe — the channel pipe, the output of which is connected to the entrance to the third internal cavity of the cooled blade of the nozzle apparatus, one of the outputs of which is connected to the input in the front device of the flame tube, in addition, there are two more exits from the third internal cavity, one of which is through the annulus of the combustion chamber and the ring the receiver cavity is connected to the entrance to the first internal cavity of the blade, and the second through the window in the separation wall to the fourth internal cavity of the blade of the nozzle apparatus, and there are at least one or more blades in the nozzle apparatus, in which the third internal the cavity has a fourth outlet connecting it through a window in the dividing wall with the second internal cavity, and in these blades there is a bypass device having a kinematic connection with a valve located at the entrance to the fuel a trunk of a frontal device connected to this blade, and the second cavity of these blades is connected to the second cavity of the blade, which does not have a bypass device. 2. The gas turbine engine according to claim 1, characterized in that the internal section of the channel pipe — a pneumatic pipe connecting the blade diffuser with the third internal cavity of the cooled blade of the nozzle apparatus, is made with a diffuser with an equivalent opening angle of less than 12 °. 3. The gas turbine engine according to claim 1, characterized in that the bypass device located inside the cooled nozzle blade has a valve with a transverse axis of rotation resting on the side of the back and trough of the blade on a pipe connecting the third internal cavity of the blade with the input of the front of the flame tube . 4. The gas turbine engine according to claim 1, characterized in that in the cooled blade with a bypass device, a bypass window is made in the partition between the second and third internal cavities, the walls of which are a seat for the rotary valve. 5. The gas turbine engine according to claim 1, characterized in that the rotary valve is profiled in the form of a flap. 6. The gas turbine engine according to claim 1, characterized in that, according to the height of the feather of the nozzle blade, holes are made connecting the first, second and fourth internal cavities of the blade with the flow part of the turbine, the holes of the first cavity being made at the input edge of the blade, the second at the trough and the back, and the fourth in the exit edge.

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