EP0781967B1 - Annular combustion chamber for gas turbine - Google Patents

Description

Technical field

The invention relates to the field of combustion technology. It relates to a gas turbine ring combustion chamber, which is operated with premix burners, and a method for Operation of this device.

State of the art

Gas turbines essentially consist of the components compressor, Combustion chamber and turbine. For environmental reasons is increased instead of diffusion combustion worked with a low-pollutant premix combustion.

It is known prior art (see H. Neuhoff and K. Thoren: "The new gas turbines GT 24 and GT 26 - high efficiency thanks to sequential combustion ", ABB Technik 2 (1994), pp. 4-7 and D. Viereck: "The gas turbine GT13E2 - a trend-setting concept for the future ", ABB Technik 6 (1993), pp. 11-16), between the compressor and the one with several Ring combustion chamber equipped with premix burners Gas turbine to arrange a plenum in which very low air speeds to rule. Through the plenum, there should be an equal distribution the air on the burners can be reached. additionally this creates a way of cooling air for the Combustion chamber and the turbine can be removed at a high pressure level.

The air coming out of the compressor is very high Speed (approx. 200 m / s) and is to the contained in it Recover kinetic energy, as lossless as possible delayed in a deflection diffuser.

In order to achieve low-pollution combustion, Fuel and combustion air premixed in the burner. For the operationally reliable design of the premixing process at the point of interference, in the vicinity of which there is a zone with stoichiometric mixture, but the speed be very high so that the flame does not kick back can be. The air in the plenum is very low Speeds (approx. 10 m / s) must therefore back to high speeds (approx. 80 to 100 m / s) in the Premix zone can be accelerated.

Around the flame downstream of the premix burner on a fixed Stabilize place, the speed in the combustion chamber strong at least locally downstream of the burner lowered. Usually a local recirculation zone with negative Speeds generated. In the combustion chamber the speed then about 50 m / s to a sufficient Preservation time and the heat transfer between hot gas and to keep the combustion chamber wall small. At the exit of the The combustion chamber accelerates again, so that on Entry of the turbine speeds up to close to the gas the speed of sound can be reached.

The multiple accelerations and decelerations of the flowing Media (air, fuel / air mixture, hot gases) between Compressor outlet and turbine inlet have the disadvantage that they are each subject to losses. You require in addition, multiple redirections of the entire air mass flow, because the distance between the compressor outlet and Turbine inlet kept small for rotor dynamic reasons must be, so that the size of the combustion chamber quite large and complicated according to the known prior art is.

Presentation of the invention

The invention tries to avoid all these disadvantages. you is based on the task of a gas turbine ring combustion chamber, which is equipped with special premix burners develop, which is characterized by a small size and simplified compared to the known prior art is, with improved premixing of fuel and Air occurs with a lower total pressure drop.

According to the invention, this is done in a gas turbine ring combustion chamber, which is arranged downstream of a compressor and on it Front plate with at least one arranged in a ring Premix burner row is equipped, achieved by direct downstream of the compressor outlet from the guide vanes of the last compressor series for each burner one as Diffuser-trained burner air duct leads to the latter downstream end there is at least one longitudinal vortex generator is located, at least in or downstream of the longitudinal vortex generator a fuel injection is provided and downstream the fuel injection ends in the combustion chamber Mixing channel of constant channel height and with a length that is approximately corresponds to twice the hydraulic duct height, is arranged.

The combustion air is released immediately after it leaves the compressor into individual air flows for the burners and for the Cooling of the combustion chamber and turbine split, after that the speed of the air for the burners to about that delayed half the value of the compressor outlet speed, then at least one longitudinal vortex per combustion air duct generated in the air, during or downstream of the Longitudinal vortex generation fuel is added to the mixture now flows along in a mixing channel and with a total swirl contaminated flows into the combustion chamber and there finally burns.

The advantages of the invention include that the combustion chamber compared to the prior art has smaller dimensions and the area to be cooled in the Combustion chamber is reduced. The pressure loss between the compressor outlet and turbine entry is smaller. Furthermore there is a very good and robust uniform distribution of the air on the burners and the premixing of fuel and combustion air will be improved.

It is particularly useful if the ratio of the number the blades of the last row of compressors to the number of premix burners is an integer, especially 1 or 2, because then a combustion air duct directly to one or two blade ducts the last row of compressors can be coupled.

It is advantageous if the mixing channel is approximately rounded Cross-section, because then a good mixing of air and fuel is achieved. But also mixed channels with a rectangular cross section are conceivable. Likewise can if there is only one burner row, the mixing channel be designed as a segmented annular gap.

It is also advantageous if the combustion air channels are spiral are arranged around the axis of the gas turbine. In this way axial length can be saved.

Finally, the axes of the mixing channels are advantageous (i.e. the direction of flow of the entering the combustion chamber Mixture), arranged so that it coincides with the axis of the Gas turbine an angle, preferably an angle of 45 °, form. This will allow the mixture and flame stabilization further improved.

Furthermore, it is useful if in the presence of more than one ring-shaped premix burner row from row to row in opposite directions in the circumferential direction are. As a result, the total swirl in the combustion chamber becomes too Zero.

It is also an advantage if there is additional air in the Boundary layer of the mixing channel is injected because of this a flashback in the mixing zone is further prevented becomes.

It is advantageous if when using fuel with average calorific value (MBtu) of this in an area of high air speed (> 100 m / s) is mixed in. This will even with these fuels, which have a very high flame speed have backfire to the fuel injector safely avoided.

Brief description of the drawing

In the drawing are several embodiments of the invention shown.

Fig. 1

einen Teillängsschnitt einer Gasturbinenanlage mit einer mit Vormischbrennern bestückten Ringbrennkammer nach dem Stand der Technik;

Fig. 2

einen Teillängsschnitt einer Gasturbinenanlage mit einer erfindungsgemässen vierreihigen Ringbrennkammer;

Fig. 3

einen Teilquerschnitt einer zweireihigen Brennkammer entsprechend einem Schnitt in der Ebene III-III der in Fig. 2 dargestellten vierreihigen Brennkammer;

Fig. 4

eine Abwicklung der Vormischstrecke (entlang IV-IV in Fig. 3) zwischen Verdichteraustritt und Brennkammerfrontplatte .

Show it:

Fig. 1

a partial longitudinal section of a gas turbine system with an annular combustion chamber equipped with premix burners according to the prior art;

Fig. 2

a partial longitudinal section of a gas turbine plant with a four-row annular combustion chamber according to the invention;

Fig. 3

a partial cross section of a two-row combustion chamber corresponding to a section in the plane III-III of the four-row combustion chamber shown in Fig. 2;

Fig. 4

a settlement of the premixing section (along IV-IV in Fig. 3) between the compressor outlet and the combustion chamber front plate.

It is only essential for understanding the invention Elements shown. The system is not shown for example the exhaust gas casing of the gas turbine with an exhaust pipe and chimney and the inlet sections of the compressor section and the low pressure compressor stages. The flow direction of the Work equipment is indicated by arrows.

Way of carrying out the invention

The invention is described below using exemplary embodiments and FIGS. 1 to 4 explained in more detail.

1 shows a partial longitudinal section of a gas turbine system with an annular combustion chamber according to the prior art. Between a compressor 1 and a turbine 2, of which only one guide vane 3 of the first row of guide vanes is shown is an annular combustion chamber 4, which with premix burners 5 of the double cone design is equipped, arranged. The supply of fuel 6 to each premix burner 5 is realized over fuel lances 7. The annular combustion chamber 4 is cooled convectively or by means of impingement cooling. The compressor 1 essentially consists of the blade carrier 8, in which the guide vanes 9 are hooked in and out of the rotor 10, which receives the blades 11. In Fig. 1 are each only the last compressor stages are shown. At the exit of the Compressor 1, a deflection diffuser 12 is arranged. It ends in a arranged between the compressor 1 and the annular combustion chamber 4 Plenary 13.

The air 14 emerging from the compressor 1 has a very high high speed. It is delayed in the deflection diffuser 12, to recover the kinetic energy it contains so that in the adjoining the deflection diffuser 12 Plenary 13 only very low air speeds to rule. This can result in a uniform distribution of the air 14 the burner 5 can be reached and there can be cooling air without any problems for the combustion chamber 4 and the turbine 2 are removed. There but on the other hand for the reliable design of the premixing process of air 14 and fuel 6 at the mixing point the fuel 6 the speed in order to avoid must be high from flashback, the air 14 in the premixing zone be accelerated again strongly before again downstream of the burner 5 in the combustion chamber 4 for reasons of flame stability the speed is reduced. Then at the downstream end of the combustion chamber 4 the gas in turn accelerates so that at the entrance to the Turbine 2 speeds close to the speed of sound can be achieved. The multiple accelerations and decelerations between compressor outlet and turbine inlet with losses and the required multiple redirections of the air mass flow lead to a very large one Height. For example, for a gas turbine the 170 MWel class according to the prior art (see FIG. 1) the outer diameter in the combustion chamber area approx. 4.5 m.

2 is an embodiment of the invention based on a four-row gas turbine ring combustion chamber. in the Difference from the prior art described above Air 14 is no longer delayed to plenary conditions, but instead the delay in the air 14 is only limited to that Speed level of the premix section. This allows the multiple redirection of the total air mass flow is eliminated and the size of the combustion chamber is significantly reduced become.

In the embodiment variant of the invention shown in FIG. 2 is immediately downstream of the compressor outlet the guide vanes 9 of the last row of compressor blades Burner air distribution system arranged to each burner 5 of the annular combustion chamber 4, each designed as a diffuser Burner air duct 15 leads. At the downstream end of the combustion air duct 15 there is at least one longitudinal vortex generator 16. In or downstream of the longitudinal vortex generator 16 at least one fuel injection 17 is provided and downstream of the fuel injection 17 is in the combustion chamber 4 ending mixing channel 19 of constant height H and with a length L, which is about twice the value of the hydraulic channel diameter D corresponds to arranged. The hydraulic channel diameter is defined as a ratio of four times Transverse area of the channel to the channel circumference. With a circular Channel therefore applies: H = D.

According to the invention, the deflection diffuser 12 and 12 is therefore omitted plenary session 13.

The air from the compressor 1 is immediately after the outlet from the compressor 1 into a large number of individual channels divided, namely into the combustion air channels 15 and in annular Channels 20 arranged on the hub side or housing side for the cooling air 21 of the combustion chamber 4 and the turbine 2, the is provided here at a high pressure level. Furthermore can air 22 from the channels 20 for flushing out the Mixing channel 19 forming boundary layer can be removed. This is only an example for the innermost mixing channel 19 shown.

The combustion air channels 15 are designed as diffusers and delay the air speed to about half the value the compressor outlet speed, with a maximum of 75% of the dynamic energy can be converted into pressure gain.

After the combustion air 14 to an appropriate speed level was delayed at the longitudinal vortex generator 16 generates one or more longitudinal vortices per combustion air duct 15. In the longitudinal vortex generator 16 is an integrated Fuel injection 17 fuel 6, which for example is supplied by fuel lances 7, mixed with the air 14. Of course, in another embodiment the fuel injection 17 also downstream of the longitudinal vortex generator 16 may be arranged. The longitudinal vortices generated guarantee a good mixture of fuel 6 and combustion air 14 in the subsequent mixing channels 19. These have a constant height H and are approximately double as long as two hydraulic channel diameters D. In the present In this case, the mixing channels 19 have a circular shape Cross section, are therefore a mixing tube. The mixing tube axes 24 are arranged parallel to the axis 25 of the gas turbine. In other exemplary embodiments, not shown here in the drawing can the mixing channels 19 a right or have polygonal cross-section or they can also be a segmented annular gap.

It is advantageous if those caused by the longitudinal vortex generator 16 Longitudinal vortex in the mixing channel 19 a total swirl generate the after the fuel / air mixture 23 into the combustion chamber 4 to a highly turbulent flame stabilization zone leads by the vertebrae bursting and on the Axis a zone with very low or negative axial speed is produced. A flashback in the Mixing zone can be achieved by a balanced axial speed profile with a cant on the axis and by a additional injection of air 22 into the boundary layer of the Mixing channel 19 can be safely prevented.

It is expedient if the number of guide vanes 9 last Compressor series and the number of premix burners 5 in have an integer relationship to each other. This can a burner air duct 15 directly to, for example, one or two blade channels of the last row of compressors are coupled become.

Comparing FIGS. 1 and 2, the reduction in the area of the combustion chamber wall to be cooled can be clearly seen according to the invention. A gas turbine from the 170 MWel class, eg GT13E2, should serve as an example. While according to the prior art (FIG. 1) the outer diameter in the area of the combustion chamber is approximately 4.5 m, this value is only 3.5 m when using the invention, so that the size is reduced by approx. 20% is reached. Due to the greatly reduced area to be cooled in the new combustion chamber and the extremely low NOx emissions that can be achieved with good premix burner technology at relatively high flame temperatures (theoretically approx. 5 ppm NOx at 15% O ₂ and 1850 K flame temperature), the combustion chamber can be cooled via film or effusion cooling.

3 and 4 show a further exemplary embodiment. In Fig. 3 is a partial cross section of a two-row annular combustion chamber corresponding to a section in the plane III-III of the in Fig. 2 shown four-row combustion chamber. The annular combustion chamber 4 according to FIG. 3 is thus with two rows Premix burners 5 equipped. The arrows in Fig. 3 are intended an opposite angle of attack of the burner 5 in the side by side Clarify rows. By this opposite Angle of attack is achieved in the combustion chamber 4 no total swirl is generated. The cross section of the mixing channels 19 is not round in this embodiment, but elliptical.

4 is a settlement of the premix section between along the compressor outlet and the combustor faceplate 18 IV-IV shown. The mixing tube axes 24 are opposite the shaft in the circumferential direction, i.e. the mixing tube axis 24 forms an angle of α with the machine axis 25 approx. 45 °. This will allow the mixture and flame stabilization improved in the combustion chamber 4.

In a further embodiment, not shown are the combustion air channels 15 spiral about the axis 25 of the Gas turbine arranged to the axial length of the machine to keep it as small as possible.

The invention is particularly suitable for the use of MBtu as fuel, i.e. fuel with a medium calorific value, for example in the gasification of heavy oil, coal and Tar arises. The fuel admixture can be used in this case very easily into a higher speed range (> 100 m / s) to be relocated to these fuels, too are characterized by a high flame speed, to avoid backfire to the fuel injector. The high-frequency generated by the last row of compressor runs (> 1000 Hz) pressure pulsations (wake of the blades) particularly support the fuel-air mixing process, because between the end of the compressor 1 and the fuel injection 17 only a short delay section, i.e. a short burner air duct 15 designed as a diffuser, is required

LIST OF REFERENCE NUMBERS

1

compressor

2

turbine

3

Guide vane from item 2

4

annular combustion chamber

5

premix

5a

outer row of burners

5b

inner row of burners

6

fuel

7

fuel lance

8th

blade carrier

9

Guide vane from item 1

10

rotor

11

Blade from pos. 1

12

deflection diffusor

13

plenum

14

air

15

Combustion air duct designed as a diffuser

16

Longitudinal vortex generators

17

fuel injection

18

front panel

19

mixing channel

20

Channel for item 21

21

cooling air

22

Air to flush out the boundary layer in pos. 19

23

Fuel / air mixture

24

Axis from item 19

25

machine axis

H

Height of item 19

L

Length of item 19

D

hydraulic channel diameter

α

Angle between items 24 and 25

Claims (15)

1. Gas-turbine annular combustion chamber (4) which is arranged downstream of a compressor (1) and is equipped on its front plate (18) with at least one row of premix burners (5) arranged in an annular form, characterized in that in each case a combustion-air duct (15) designed as a diffuser leads directly downstream of the compressor outlet from the guide vanes (9) of the last compressor row to each burner (5), at the downstream end of which combustion-air duct (15) at least one longitudinal-vortex generator (16) is located, at least one fuel injection means (17) being provided in or downstream of the longitudinal-vortex generator (16), and a mixing duct (19) which ends in the combustion chamber (4) and has a constant height (H) and a length (L) which corresponds approximately to twice the value of the hydraulic duct diameter (D) being arranged downstream of the fuel injection means (17).
2. Gas-turbine annular combustion chamber according to Claim 1, characterized in that the ratio of the number of blades (9) of the last compressor row to the number of premix burners (5) is integral.
3. Gas-turbine annular combustion chamber according to Claim 2, characterized in that the ratio of the number of blades (9) of the last compressor row to the number of premix burners (5) is one.
4. Gas-turbine annular combustion chamber according to Claim 2, characterized in that the ratio of the number of blades (9) of the last compressor row to the number of premix burners (5) is two.
5. Gas-turbine annular combustion chamber according to Claim 1, characterized in that the combustion-air ducts (15) are arranged spirally around the axis (25) of the gas turbine.
6. Gas-turbine annular combustion chamber according to Claim 1, characterized in that the mixing duct (19) has a round cross section.
7. Gas-turbine annular combustion chamber according to Claim 1, characterized in that the mixing duct (19) has a rectangular cross section.
8. Gas-turbine annular combustion chamber according to Claim 1, characterized in that the mixing duct (19) is a segmented annular gap.
9. Gas-turbine annular combustion chamber according to Claim 1, characterized in that the axes (24) of the mixing ducts (19) and the axis (25) of the gas turbine are parallel.
10. Gas-turbine annular combustion chamber according to Claim 1, characterized in that the axes (24) of the mixing ducts (19) form an angle (α) with the axis (25) of the gas turbine.
11. Gas-turbine annular combustion chamber according to Claim 10, characterized in that the angle (α) is about 45°.
12. Gas-turbine annular combustion chamber according to one of claims 1 to 11, characterized in that, in the case of more than one annular premix-burner row, the burners (5) are set in an opposed manner from row (5a) to row (5b) in the peripheral direction.
13. Method of operating a gas-turbine annular combustion chamber according to one of claims 1 to 12, characterized in that the combustion air (15), directly after discharge from the compressor (1), is split up into individual air flows for the burners and for the cooling of the combustion chamber and the turbine, in that the velocity of the air (14) for the burners (5) is then decelerated in the combustion-air ducts (15) to about half the value of the compressor outlet velocity, and in that at least one longitudinal vortex is then generated in the air (14) per combustion-air duct (15), fuel (6) being admixed during or downstream of the longitudinal-vortex generation, the mixture flowing along in a mixing duct (19) and flowing with an overall swirl imposed on it into the combustion chamber (4) and being burnt there.
14. Method according to Claim 13, characterized in that air (22) is additionally injected into the boundary layer of the mixing duct (19).
15. Method according to Claim 13, characterized in that, when fuel (6) having an average calorific value (MBtu) is used, this fuel (6) is intermixed in a region of high air velocity of greater than 100 m/s.

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