Low emissions combustor assembly for gas turbine engine

Patent metadata

Field	Value
Publication number	US-10954859-B2
Application number	US-201715658464-A
Country	US
Kind code	B2
Filing date	Jul 25, 2017
Priority date	Jul 25, 2017
Publication date	Mar 23, 2021
Grant date	Mar 23, 2021

Abstract

A combustor assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a combustion chamber, and a fuel injector assembly in communication with the combustion chamber that has a swirler body situated about a nozzle to define an injector passage that converges to a throat. The throat is defined at a distance from the combustion chamber. The nozzle includes a primary fuel injector along a first fuel injector axis and at least one secondary plain jet fuel injector axially forward of the primary fuel injector.

First claim

What is claimed is:

1. A combustor assembly for a gas turbine engine comprising:
a combustion chamber;
a fuel injector assembly in communication with the combustion chamber, comprising:
a swirler body situated about a nozzle to define an injector passage that converges to a throat, the throat defined at a distance from the combustion chamber; and
wherein the nozzle includes a nozzle body extending along a nozzle longitudinal axis between a first end portion and a second end portion having a dome shaped profile, a primary fuel injector along a first fuel injector axis and an array of secondary plain jet fuel injectors axially forward of the primary fuel injector, the first fuel injector axis extending through a primary outlet of the primary fuel injector along the second end portion;
wherein the injector passage includes a first passage section merging into a second passage section, the first passage section defined between the swirler body and the nozzle, and the second passage section defining the throat;
wherein outlets of the secondary plain jet fuel injectors are circumferentially distributed about a periphery of the nozzle body defining the first passage section such that the outlets are spaced apart from a terminal end of the second end portion, and each secondary plain jet defines a second fuel injector axis extending through the respective outlet such that a projection of the second fuel injector axis intersects a radially outer wall of the injector passage along the first passage section; and
wherein the radially outer wall of the first passage section slopes toward the first fuel injector axis such that a first projection of the radially outer wall intersects a projection of the first fuel injector axis along the injector passage to define a first passage angle, the first passage angle being between about 30 degrees and about 50 degrees, and a radially inner wall of the first passage section slopes toward the first fuel injector axis such that a second projection of the radially inner wall intersects the projection of the first fuel injector axis along the injector passage to define a second passage angle, the second passage angle being within about 5 degrees of the first passage angle.

2. The combustor assembly as recited in claim 1 , wherein the injector passage defines a first passage area with respect to a first reference plane that is perpendicular to the first fuel injector axis and intersects the outlet of a forwardmost one of the secondary plain-jet fuel injectors, the injector passage defines a second passage area with respect to a second reference plane along the throat that is perpendicular to the first fuel injector axis, and a ratio of the first passage area to the second passage area being about 1.2 to about 1.5.

3. The combustor assembly as recited in claim 2 , wherein the injector passage diverges aft of the throat with respect to the first fuel injector axis.

4. The combustor assembly as recited in claim 1 , wherein the outlet of a forwardmost one of the secondary plain-jet fuel injectors and the throat define a first distance with respect to the first fuel injector axis, the outlet of the primary fuel injector and the throat defines a second distance with respect to the first fuel injector axis, and a ratio of the first distance to the second distance being between about 2.0 to about 4.0.

5. The combustor assembly as recited in claim 1 , wherein the primary fuel injector is arranged to generate a non-premixed fuel stream, and the secondary plain-jet fuel injectors are arranged to generate a substantially premixed fuel stream.

6. The combustor assembly as recited in claim 5 , comprising a control that meters flow of fuel to the primary fuel injector and flow of fuel to the at secondary plain-jet fuel injectors in a first mode, and meters flow of fuel to the primary fuel injector and to the secondary plain-jet fuel injectors in a second, different mode such that a rate of the flow of fuel to the primary fuel injector and the secondary plain-jet fuel injectors differs from the first mode.

7. A gas turbine engine comprising:
a fan section including a plurality of fan blades rotatable about an engine axis;
a compressor section in fluid communication with the fan section;
a turbine section driving the fan section; and
a combustor section in fluid communication with the compressor section and the turbine section, the combustor section comprising:
a combustion chamber extending from a bulkhead;
a fuel injector assembly along the bulkhead, the fuel injector assembly including a nozzle and a swirler body, the nozzle including a nozzle body extending along a nozzle longitudinal axis between a first end portion and a second end portion having a dome shaped profile, the nozzle including a primary fuel injector along the nozzle longitudinal axis and an array of secondary plain-jet fuel injectors arranged about a periphery of the nozzle, the nozzle longitudinal axis extending through a primary outlet of the primary fuel injector along the second end portion, and the swirler body situated about the nozzle to define an injector passage that converges to a throat such that flow through the injector passage accelerates towards the throat;
wherein the injector passage includes a first passage section merging into a second passage section, the first passage section defined between the swirler body and the nozzle, and the second passage section defining the throat;
wherein outlets of the secondary plain jet fuel injectors are circumferentially distributed about a periphery of the nozzle body defining the first passage section such that the outlets are spaced apart from a terminal end of the second end portion, and each secondary plain jet defines a second fuel injector axis extending through the respective outlet such that a projection of the second fuel injector axis intersects a radially outer wall of the injector passage along the first passage section; and
wherein the radially outer wall of the first passage section slopes toward the nozzle longitudinal axis such that a first projection of the radially outer wall intersects the nozzle longitudinal axis along the injector passage to define a first passage angle, the first passage angle being between 30 degrees and about 50 degrees, and a radially inner wall of the first passage section along the nozzle body slopes toward the nozzle longitudinal axis such that a second projection of the radially inner wall intersects the nozzle longitudinal axis along the injector passage to define a second passage angle, the second passage angle being within about 5 degrees of the first passage angle.

8. The gas turbine engine as recited in claim 7 , wherein the throat is defined at a position axially forward of the combustion chamber relative to the nozzle longitudinal axis.

9. The gas turbine engine as recited in claim 8 , wherein the first passage section being an annulus defined between the swirler body and the nozzle, the first passage section extending from an array of vanes that provide airflow to the first passage section, and the second passage section defining the throat at a position axially aft of the primary fuel injector relative to the nozzle longitudinal axis.

10. The gas turbine engine as recited in claim 7 , comprising a controller that meters flow of fuel to the primary fuel injector and flow of fuel to the array of secondary plain-jet fuel injector in a first operating condition of the engine, and meters flow of fuel to the primary fuel injector and to the array of secondary plain-jet fuel injectors in a second, different operating condition of the engine such that a rate of the flow of fuel to the primary fuel injector and the array of secondary plain-

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Dai Zhongtao
• Smith Lance L
• Cohen Jeffrey M

Classifications

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F23C7/004

  using vanes · CPC title

• F23R3/26

  Controlling the air flow · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F23D2900/00008

  Burner assemblies with diffusion and premix modes, i.e. dual mode burners · CPC title