EP3221642A1 - Lance injector for injecting fuel into a combustion chamber of a gas turbine - Google Patents

Lance injector for injecting fuel into a combustion chamber of a gas turbine

Info

Publication number
EP3221642A1
EP3221642A1 EP15823002.9A EP15823002A EP3221642A1 EP 3221642 A1 EP3221642 A1 EP 3221642A1 EP 15823002 A EP15823002 A EP 15823002A EP 3221642 A1 EP3221642 A1 EP 3221642A1
Authority
EP
European Patent Office
Prior art keywords
tubular body
sealing rings
injector according
foregoing
injector
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP15823002.9A
Other languages
German (de)
French (fr)
Other versions
EP3221642B1 (en
Inventor
Marco Alecci
Pierpaolo Pastorino
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Ansaldo Energia SpA
Original Assignee
Ansaldo Energia SpA
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Ansaldo Energia SpA filed Critical Ansaldo Energia SpA
Publication of EP3221642A1 publication Critical patent/EP3221642A1/en
Application granted granted Critical
Publication of EP3221642B1 publication Critical patent/EP3221642B1/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply

Definitions

  • the present invention relates to a lance injector for injecting fuel into a combustion chamber of a gas turbine.
  • gas turbines especially if used in electric power production plants, may be fed with different types of fuel.
  • gaseous fuels of different nature and characteristics natural gas, syngas
  • fuel oils like diesel fuel can be injected into gas turbines.
  • gas turbines are equipped with burners comprising injectors, normally of the lance type, specifically designed for injecting a controlled flow of fuel.
  • Lance injectors generally comprise a plurality of coaxial tubular bodies, at one end of which a terminal equipped with a nozzle is fitted.
  • the tubular bodies define between them at least one delivery line between an inlet and the nozzle, and a return line allowing the recovery of the excess fuel supplied to the nozzle.
  • a line for feeding cooling water or air can also be provided.
  • a common problem in lance injectors is due to the different thermal expansion of the components in use, with particular reference to the tubular bodies.
  • the thermal stress is in fact remarkable because of the high temperatures in the combustion chamber (i.e. close to the terminal and the nozzle) .
  • the consequent thermal expansion is not only important, but also remarkably different from one component to another, due to the differences in geometry and in the exposure to high temperatures in the combustion chamber and to the cooling air introduced from outside or to the process fluid.
  • the object of the present invention is therefore to provide a lance injector for injecting fuel into a combustion chamber of a gas turbine that can overcome the described limitations .
  • the present invention provides a lance injector for injecting fuel into a combustion chamber of a gas turbine according to claim 1.
  • FIG. 1 is a side view, sectioned according to an axial longitudinal plane, with parts removed for clarity's sake, of a burner of a gas turbine incorporating a lance injector according to an embodiment of the present invention
  • FIG. 2 shows an enlarged detail of the lance injector of Figure 1;
  • FIG. 3 shows a further enlarged detail of the lance injector of Figure 1 ;
  • FIG. 4 shows a variant of the detail of Figure 2, according to a different embodiment of the present invention .
  • the reference number 1 indicates in its entirety a burner unit for injecting fuel, in particular gas, into a combustion chamber 2 of a gas turbine, only partially shown.
  • the burner unit 1 extends along an axis A and comprises a peripheral gas main burner 3, a central gas pilot burner 4 and a lance injector 5 for injecting fuel oil into the combustion chamber 2.
  • the main burner 3 is of the premixing type, is arranged around the pilot burner 4 and is provided with a swirler 7, which comprises a plurality of blades 10, defining respective flow channels for conveying, oblique with respect to the axis A, a flow of combustion air and fuel gas towards the combustion chamber 2.
  • the fuel gas is supplied through nozzles 11 arranged on the blades 10.
  • the pilot burner 4 is coaxial to the main burner 3 and is provided with an axial swirler 8, which comprises a plurality of blades 12, defining respective flow channels for conveying, substantially along the axis A, a further flow of combustion air towards the combustion chamber 2.
  • the lance injector 5 extends along the axis A and has one end inserted in the pilot burner 4.
  • the lance injector 5 comprises an inner tubular body 15, an intermediate tubular body 16, an outer tubular body 17, a terminal element 18 and a nozzle 20. Furthermore, the lance injector 5 is provided with flanges 21, 22.
  • the inner tubular body 15, the intermediate tubular body 16 and the outer tubular body 17 concentrically extend along the axis A and define a first fluid line 23, a second fluid line 24 and a third fluid line 25.
  • the first fluid line 23 is defined in a gap between the inner tubular body 15 and the intermediate tubular body 16 and allows the alternating supply of a flow of cooling air along the lance injector 5 or of a flow of fuel oil to the nozzle 20, thus working as a delivery line.
  • the second fluid line 24 is defined in the inner tubular body 15 and allows to recover the excess fuel oil fed to the nozzle 20 and to reintroduce it in a manifold (not shown) , thus working as a return line.
  • the third fluid line 25 is defined in a gap between the intermediate tubular body 16 and the outer tubular body 17 and may supply a different type of fuel, for example gas, or an inert fluid like water .
  • the nozzle 20 is coupled to the inner tubular body 15, to the intermediate tubular body 16 and to the outer tubular body 17 with the aid of the terminal element 18.
  • the nozzle 20 is mounted on the ends 15a, 16a of the inner tubular body 15 and of the intermediate tubular body 16 by a threaded or hot-keyed coupling.
  • the terminal element 18 is fitted on one end 17a of the outer tubular body 17 and coupled to it by a threaded coupling or, where appropriate, by a hot-keyed coupling.
  • the nozzle 20 and the terminal element 18 are coupled by interference, and the nozzle 20 axially protrudes outwards with respect to the terminal element 18.
  • the flange 21 extends around the axis A and has a through axial cavity 29 and radial fluid passages 27, 28 (Figure 1), respectively communicating with the first fluid line 23 and the third fluid line 25.
  • the axial cavity 29 houses one end 15b of the inner tubular body 15, which extends beyond the edge of the intermediate tubular body 16 and is inserted in a guiding bushing 30.
  • the guiding bushing 30 is tightened between the flange 21 and the flange 22 and is fitted onto the inner tubular body 15.
  • the guiding bushing 30 therefore supports the inner tubular body 15 in a position centred on the axis A and allows its axial sliding, in particular as a result of the thermal expansion and of the subsequent contraction of the portion nearest to the combustion chamber 2.
  • the guiding bushing 30 ensures the fluid-tight seal against the side wall of the inner tubular body 15 to prevent a leakage of fuel oil, as explained in more detail below.
  • the flange 22 is coupled to the flange 21 and has an axial cavity 31, which defines a contrast seat and houses a contrast elastic device 32 co-operating with the inner tubular body 15 to absorb and recover the effects of thermal expansion and contraction.
  • the device 32 comprises an elastic contrast pushing member 33 coupled to the end 15b of the inner tubular body 15 and a stack of belleville washers 35 co-operating with the pushing member 33 and the contrast seat defined in the axial cavity 31 of the flange 22.
  • the pushing member 33 is axially hollow and, in one embodiment, comprises a tubular body 33a, abutting the end 15b of the inner tubular body 15 and provided with a plate 33b for its coupling to the belleville washers 35, which are housed in the contrast seat of the axial cavity 31.
  • the pushing member 33 puts in fluid communication the inner tubular body 15 with a supply duct (not shown) .
  • the sealing members 37 include at least two sealing rings 38 that, in one embodiment, are integral with the inner tubular body 15.
  • the sealing rings 38 define the only points of contact between the inner tubular body 15 and the guiding bushing 30.
  • the inner tubular body 15 and the guiding bushing 30 are then separated by an annular gap 40 except at the sealing rings 38.
  • the sealing rings 38 have a radially rounded outer profile, for example having a circular section with a radius Rl comprised between 0.5 mm and 4 mm, as shown in Figure 3. In this way, the contact between the inner tubular body 15 and the guiding bushing 30 is limited to the crests of the sealing rings 38, but is sufficient to prevent the leakage of fluid.
  • the sealing rings 38 radially protrude from the side surface of the inner tubular body 15 and their radial thickness T is comprised between 10% and 25% of a radius R2 of the side surface of the inner tubular body 15.
  • the radial thickness T of the sealing rings 38 is comprised between 0.5 mm and 1.5 mm.
  • the sealing rings 38 are mutually spaced by a distance D comprised between 2 mm and 5 mm.
  • sealing members 137 in particular comprising at least two sealing rings 138, are present between the inner tubular body and the guiding bushing, here respectively indicated by 115 and 130.
  • the inner tubular body 115 and the guiding bushing 130 are then separated by an annular gap 140 except at the sealing rings 138.
  • the sealing rings 138 are integral with the guiding bushing 130 and radially project towards the inner tubular body 115.
  • the sealing rings 138 have a radially rounded inner profile, so that the contact between the inner tubular body 115 and the guiding bushing 130 is limited to the crests of the sealing rings 138, but is sufficient to prevent the leakage of fluid.
  • the thermal stresses occurring during normal operation may cause serious problems in the lance injectors of known type.
  • the result is a loss of efficiency of the machine, an imperfect control of combustion conditions and emissions and, in serious cases, the machine stop.
  • one of the phenomena observed in repeated thermal cycles during the tests is the bending of the tubular bodies, which is not compatible with a sliding coupling with a minimum clearance between the inner tubular body and the guiding bushing. It was therefore devised the described tight coupling between the inner tubular body and the guiding bushing. In particular, the described coupling limits the points of contact between the inner tubular body and the guiding bushing to the crests of the sealing rings, while for the rest, the inner tubular body and the guiding bushing are separated by an annular gap.
  • the contact limited to the crests of the sealing rings and the greater available clearance allow some misalignment of the end of the inner tubular body with respect to the guiding bushing without compromising the sealing function, and therefore avoiding any leakage.
  • the described solution thus favours the relative axial sliding between the inner tubular body and the intermediate tubular body and avoids any jamming. Consequently, the stress on the element terminal and on the nozzle due to the thermal expansion is reduced and the possible damage is significantly limited, especially with regard to the fitting loss.
  • the fluid lines can otherwise be used for injecting one or more liquid or gaseous fuels or a cooling fluid, for example air or water.
  • one of the tubular bodies in particular one between the intermediate tubular body and the outer tubular body, may be missing.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Nozzles For Spraying Of Liquid Fuel (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A lance injector for injecting fuel into a combustion chamber (2) of a gas turbine comprises: a coaxial first tubular body (15) and second tubular body (16); a nozzle (20) connected to respective first ends (15a, 16a) of the first tubular body (15) and of the second tubular body (16); a guiding bushing (30) fitted on the first tubular body (15) so as to allow the relative axial sliding of the first tubular body (15) with respect to the second tubular body (16) due to expansion and contraction; and sealing members (27) between the first tubular body (15) and the guiding bushing (30). The sealing members are defined by at least two sealing rings (38) and the first tubular body (15) and the guiding bushing (30) are separated by an annular gap, except at the sealing rings (38).

Description

"LANCE INJECTOR FOR INJECTING FUEL INTO A COMBUSTION CHAMBER OF A GAS TURBINE"
TECHNICAL FIELD
The present invention relates to a lance injector for injecting fuel into a combustion chamber of a gas turbine. BACKGROUND ART
As already known, gas turbines, especially if used in electric power production plants, may be fed with different types of fuel. In particular, it is known that gaseous fuels of different nature and characteristics (natural gas, syngas) or fuel oils like diesel fuel can be injected into gas turbines. For this reason, gas turbines are equipped with burners comprising injectors, normally of the lance type, specifically designed for injecting a controlled flow of fuel.
Lance injectors generally comprise a plurality of coaxial tubular bodies, at one end of which a terminal equipped with a nozzle is fitted. The tubular bodies define between them at least one delivery line between an inlet and the nozzle, and a return line allowing the recovery of the excess fuel supplied to the nozzle. Possibly, a line for feeding cooling water or air can also be provided.
A common problem in lance injectors is due to the different thermal expansion of the components in use, with particular reference to the tubular bodies. The thermal stress is in fact remarkable because of the high temperatures in the combustion chamber (i.e. close to the terminal and the nozzle) . The consequent thermal expansion is not only important, but also remarkably different from one component to another, due to the differences in geometry and in the exposure to high temperatures in the combustion chamber and to the cooling air introduced from outside or to the process fluid.
In turn, different expansions produce intense mechanical stress and may cause damage to components, fitting loss and, in general, malfunctioning. A frequent failure caused by a fitting loss is the partial or total obstruction of the fuel fluid passage between the outlet of the flow line, the nozzle and the inlet of the return line. In these cases, the scheduled fuel amount cannot be correctly injected, with negative effects, e.g. on the machine efficiency and on pollutant emissions. In some cases, the plant must be stopped to replace the damaged injectors. DISCLOSURE OF INVENTION
The object of the present invention is therefore to provide a lance injector for injecting fuel into a combustion chamber of a gas turbine that can overcome the described limitations .
The present invention provides a lance injector for injecting fuel into a combustion chamber of a gas turbine according to claim 1.
BRIEF DESCRIPTION OF THE DRAWINGS
The present invention will now be described with reference to the accompanying drawings, which illustrate some examples of non-limiting embodiments, in which:
- Figure 1 is a side view, sectioned according to an axial longitudinal plane, with parts removed for clarity's sake, of a burner of a gas turbine incorporating a lance injector according to an embodiment of the present invention;
- Figure 2 shows an enlarged detail of the lance injector of Figure 1;
- Figure 3 shows a further enlarged detail of the lance injector of Figure 1 ;
- Figure 4 shows a variant of the detail of Figure 2, according to a different embodiment of the present invention .
BEST MODE FOR CARRYING OUT THE INVENTION
In Figure 1, the reference number 1 indicates in its entirety a burner unit for injecting fuel, in particular gas, into a combustion chamber 2 of a gas turbine, only partially shown. The burner unit 1 extends along an axis A and comprises a peripheral gas main burner 3, a central gas pilot burner 4 and a lance injector 5 for injecting fuel oil into the combustion chamber 2. The main burner 3 is of the premixing type, is arranged around the pilot burner 4 and is provided with a swirler 7, which comprises a plurality of blades 10, defining respective flow channels for conveying, oblique with respect to the axis A, a flow of combustion air and fuel gas towards the combustion chamber 2. In one embodiment, the fuel gas is supplied through nozzles 11 arranged on the blades 10.
The pilot burner 4 is coaxial to the main burner 3 and is provided with an axial swirler 8, which comprises a plurality of blades 12, defining respective flow channels for conveying, substantially along the axis A, a further flow of combustion air towards the combustion chamber 2. The lance injector 5 extends along the axis A and has one end inserted in the pilot burner 4.
The lance injector 5 comprises an inner tubular body 15, an intermediate tubular body 16, an outer tubular body 17, a terminal element 18 and a nozzle 20. Furthermore, the lance injector 5 is provided with flanges 21, 22.
The inner tubular body 15, the intermediate tubular body 16 and the outer tubular body 17 concentrically extend along the axis A and define a first fluid line 23, a second fluid line 24 and a third fluid line 25. More in detail, the first fluid line 23 is defined in a gap between the inner tubular body 15 and the intermediate tubular body 16 and allows the alternating supply of a flow of cooling air along the lance injector 5 or of a flow of fuel oil to the nozzle 20, thus working as a delivery line. The second fluid line 24 is defined in the inner tubular body 15 and allows to recover the excess fuel oil fed to the nozzle 20 and to reintroduce it in a manifold (not shown) , thus working as a return line. The third fluid line 25 is defined in a gap between the intermediate tubular body 16 and the outer tubular body 17 and may supply a different type of fuel, for example gas, or an inert fluid like water .
The nozzle 20 is coupled to the inner tubular body 15, to the intermediate tubular body 16 and to the outer tubular body 17 with the aid of the terminal element 18. For example, the nozzle 20 is mounted on the ends 15a, 16a of the inner tubular body 15 and of the intermediate tubular body 16 by a threaded or hot-keyed coupling. The terminal element 18 is fitted on one end 17a of the outer tubular body 17 and coupled to it by a threaded coupling or, where appropriate, by a hot-keyed coupling. The nozzle 20 and the terminal element 18 are coupled by interference, and the nozzle 20 axially protrudes outwards with respect to the terminal element 18.
With reference also to Figure 2, the flange 21 extends around the axis A and has a through axial cavity 29 and radial fluid passages 27, 28 (Figure 1), respectively communicating with the first fluid line 23 and the third fluid line 25.
The axial cavity 29 houses one end 15b of the inner tubular body 15, which extends beyond the edge of the intermediate tubular body 16 and is inserted in a guiding bushing 30. The guiding bushing 30 is tightened between the flange 21 and the flange 22 and is fitted onto the inner tubular body 15. The guiding bushing 30 therefore supports the inner tubular body 15 in a position centred on the axis A and allows its axial sliding, in particular as a result of the thermal expansion and of the subsequent contraction of the portion nearest to the combustion chamber 2. Furthermore, the guiding bushing 30 ensures the fluid-tight seal against the side wall of the inner tubular body 15 to prevent a leakage of fuel oil, as explained in more detail below. The flange 22 is coupled to the flange 21 and has an axial cavity 31, which defines a contrast seat and houses a contrast elastic device 32 co-operating with the inner tubular body 15 to absorb and recover the effects of thermal expansion and contraction.
The device 32 comprises an elastic contrast pushing member 33 coupled to the end 15b of the inner tubular body 15 and a stack of belleville washers 35 co-operating with the pushing member 33 and the contrast seat defined in the axial cavity 31 of the flange 22. The pushing member 33 is axially hollow and, in one embodiment, comprises a tubular body 33a, abutting the end 15b of the inner tubular body 15 and provided with a plate 33b for its coupling to the belleville washers 35, which are housed in the contrast seat of the axial cavity 31. The pushing member 33 puts in fluid communication the inner tubular body 15 with a supply duct (not shown) .
With reference to Figures 2 and 3, the fluid-tight seal between the surface of the inner tubular body 15 and the guiding bushing 30 is ensured by sealing members 37. More in detail, the sealing members 37 include at least two sealing rings 38 that, in one embodiment, are integral with the inner tubular body 15. The sealing rings 38 define the only points of contact between the inner tubular body 15 and the guiding bushing 30. The inner tubular body 15 and the guiding bushing 30 are then separated by an annular gap 40 except at the sealing rings 38.
The sealing rings 38 have a radially rounded outer profile, for example having a circular section with a radius Rl comprised between 0.5 mm and 4 mm, as shown in Figure 3. In this way, the contact between the inner tubular body 15 and the guiding bushing 30 is limited to the crests of the sealing rings 38, but is sufficient to prevent the leakage of fluid. The sealing rings 38 radially protrude from the side surface of the inner tubular body 15 and their radial thickness T is comprised between 10% and 25% of a radius R2 of the side surface of the inner tubular body 15. For example, the radial thickness T of the sealing rings 38 is comprised between 0.5 mm and 1.5 mm.
In the axial direction, the sealing rings 38 are mutually spaced by a distance D comprised between 2 mm and 5 mm.
In a different embodiment, shown in Figure 4, sealing members 137, in particular comprising at least two sealing rings 138, are present between the inner tubular body and the guiding bushing, here respectively indicated by 115 and 130. The inner tubular body 115 and the guiding bushing 130 are then separated by an annular gap 140 except at the sealing rings 138.
The sealing rings 138 are integral with the guiding bushing 130 and radially project towards the inner tubular body 115.
The sealing rings 138 have a radially rounded inner profile, so that the contact between the inner tubular body 115 and the guiding bushing 130 is limited to the crests of the sealing rings 138, but is sufficient to prevent the leakage of fluid.
As already stated, the thermal stresses occurring during normal operation may cause serious problems in the lance injectors of known type. In particular, it is not uncommon to see a fitting loss of the terminal element and of the nozzle, which in turn may cause a partial or total obstruction of the inner fluid passages of the injectors. The result is a loss of efficiency of the machine, an imperfect control of combustion conditions and emissions and, in serious cases, the machine stop.
However, tests performed by the Applicant have shown that the causes of fitting loss do not have to be sought mainly in the stress produced by the thermal expansion of the tubular bodies in the area of the terminal element and of the nozzle showing the visible damage. The results of the tests rather demonstrate that the fitting losses are generally caused by the imperfect relative sliding between the tubular bodies, in particular the inner tubular body and the intermediate tubular body, when the variations in the operating conditions determine a thermal expansion and a subsequent contraction. More specifically, the Applicant has identified a major cause of fitting loss in the coupling between the inner tubular body and the guiding bushing, therefore at the opposite end of the tubular bodies with respect to the terminal element and to the nozzle showing the damage. For example, one of the phenomena observed in repeated thermal cycles during the tests is the bending of the tubular bodies, which is not compatible with a sliding coupling with a minimum clearance between the inner tubular body and the guiding bushing. It was therefore devised the described tight coupling between the inner tubular body and the guiding bushing. In particular, the described coupling limits the points of contact between the inner tubular body and the guiding bushing to the crests of the sealing rings, while for the rest, the inner tubular body and the guiding bushing are separated by an annular gap.
The contact limited to the crests of the sealing rings and the greater available clearance allow some misalignment of the end of the inner tubular body with respect to the guiding bushing without compromising the sealing function, and therefore avoiding any leakage. The described solution thus favours the relative axial sliding between the inner tubular body and the intermediate tubular body and avoids any jamming. Consequently, the stress on the element terminal and on the nozzle due to the thermal expansion is reduced and the possible damage is significantly limited, especially with regard to the fitting loss.
Finally, it is evident that the described lance injector can have modifications and variations without departing from the scope of the present invention, as defined in the appended claims.
In particular, the fluid lines can otherwise be used for injecting one or more liquid or gaseous fuels or a cooling fluid, for example air or water.
In addition, one of the tubular bodies, in particular one between the intermediate tubular body and the outer tubular body, may be missing.

Claims

1. Lance injector for injecting fuel into a combustion chamber of a gas turbine, the lance injector a comprising: a first tubular body (15; 115) and a second tubular body (16; 116), the first tubular body (15; 115) being arranged coaxially inside the second tubular body (16; 116) ;
a nozzle (20) connected to respective first ends (15a, 16a) of the first tubular body (15; 115) and of the second tubular body (16; 116);
a guiding bushing (30; 130) fitted onto the first tubular body (15; 115) so as to allow the first tubular body (15; 115) to axially slide with respect to the second tubular body (16; 116) on account of thermal expansion and contraction; and
sealing members (27; 127) between the first tubular body (15; 115) and the guiding bushing (30; 130);
wherein the sealing members are defined by at least two sealing rings (38; 138) and the first tubular body (15; 115) and the guiding bushing (30; 130) are separated by a annular gap except at the sealing rings (38; 138) .
2. Injector according to claim 1, wherein the sealing rings (38) are integral with the first tubular body (15) .
3. Injector according to claim 1, wherein the sealing rings (138) are integral with the guiding bushing (130) .
4. Injector according to any one of the foregoing claims, wherein sealing portions of the sealing rings (38; 138) have a round profile.
5. Injector according to claim 4, wherein the sealing portions of the sealing rings (38; 138) have circular cross-section with radius (Rl) comprised between 0.5 mm and 4 mm.
6. Injector according to any one of the foregoing claims, wherein the sealing rings (38; 138) have a radial thickness (T) comprised between 10% and 25% of a radius (R2) of a lateral surface of the first tubular body (15; 115) .
7. Injector according to any one of the foregoing claims, wherein the sealing rings (38; 138) have a radial thickness (T) comprised between 0.5 mm and 1.5 mm.
8. Injector according to any one of the foregoing claims, wherein the sealing rings (38; 138) are mutually spaced by a distance (D) comprised between 2 mm and 5 mm.
9. Injector according to any one of the foregoing claims, comprising a contrast elastic device (32), that is loaded in response to the thermal expansion of the first tubular body (15) and releases energy to the first tubular body (15) in response to the thermal contraction of the first tubular body (15) .
10. Injector according to claim 9, wherein the contrast elastic device (32) comprises a pushing member (33) coupled to a second end of the first tubular body (15) and a stack of belleville washers co-operating with the pushing member and a contrast seat.
11. Injector according to any one of the foregoing claims, comprising a first fluid line (23), defined between the first tubular body (15) and the second tubular body (16), and a second fluid line (24), defined inside del first tubular body (15) .
12. Gas turbine burner, comprising a lance injector
(5) according to any one of the foregoing claims.
EP15823002.9A 2014-11-21 2015-11-20 Lance injector for injecting fuel into a combustion chamber of a gas turbine Active EP3221642B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
ITMI20142019 2014-11-21
PCT/IB2015/059018 WO2016079720A1 (en) 2014-11-21 2015-11-20 Lance injector for injecting fuel into a combustion chamber of a gas turbine

Publications (2)

Publication Number Publication Date
EP3221642A1 true EP3221642A1 (en) 2017-09-27
EP3221642B1 EP3221642B1 (en) 2018-09-05

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP15823002.9A Active EP3221642B1 (en) 2014-11-21 2015-11-20 Lance injector for injecting fuel into a combustion chamber of a gas turbine

Country Status (4)

Country Link
EP (1) EP3221642B1 (en)
CN (1) CN107110504B (en)
RU (1) RU2693202C2 (en)
WO (1) WO2016079720A1 (en)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114543118B (en) * 2022-02-21 2023-04-25 中国航发贵阳发动机设计研究所 Centrifugal starting nozzle

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB648699A (en) * 1947-10-17 1951-01-10 Arthur Holmes Fletcher Improvements in or relating to gas-turbine engine fuel systems and liquid fuel injectors therefor
US6357222B1 (en) * 2000-04-07 2002-03-19 General Electric Company Method and apparatus for reducing thermal stresses within turbine engines
US7703287B2 (en) * 2006-10-31 2010-04-27 Delavan Inc Dynamic sealing assembly to accommodate differential thermal growth of fuel injector components
RU2375597C2 (en) * 2008-02-12 2009-12-10 Открытое акционерное общество "Научно-производственное объединение "Сатурн" (ОАО "НПО "Сатурн") Gas turbine engine combustion chamber fuel manifold
US8443608B2 (en) * 2008-02-26 2013-05-21 Delavan Inc Feed arm for a multiple circuit fuel injector
EP2270398A1 (en) * 2009-06-30 2011-01-05 Siemens Aktiengesellschaft Burner, especially for gas turbines
US20130180261A1 (en) * 2012-01-13 2013-07-18 General Electric Company Combustor and method for reducing thermal stresses in a combustor

Also Published As

Publication number Publication date
RU2017121565A3 (en) 2019-05-21
RU2017121565A (en) 2018-12-21
RU2693202C2 (en) 2019-07-01
WO2016079720A1 (en) 2016-05-26
CN107110504B (en) 2019-11-26
EP3221642B1 (en) 2018-09-05
CN107110504A (en) 2017-08-29

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