Aerodynamic injection system for aircraft turbine engine, having improved air/fuel mixing
To improve air/fuel mixing at low speed, the invention concerns an aerodynamic injection system ( 30 ) for a combustion chamber ( 8 ) of an aircraft turbine engine, the system comprising a central body ( 70 ) comprising a first air circulation space ( 72 ) and a swirler ( 74 ) allowing air to penetrate into the first air circulation space ( 72 ), the hollow body ( 70 ) also comprising a fuel circuit ( 92, 93, 82 ) arranged at least partially around the first space ( 72 ), the circuit comprising a second annular space ( 82 ) for a fuel film to circulate and fuel supply structures ( 92, 93 ) communicating with the second space ( 82 ), the spaces ( 72, 82 ) opening into a third air and fuel mixing space ( 86 ). According to the invention, the system comprises fuel injection openings ( 90 ) that bring the fuel circuit into communication with the first space ( 72 ).
1. An aerodynamic injection system ( 30 ) for a combustion chamber ( 8 ) of an aircraft turbomachine, the system ( 30 ) comprising:
a hollow central body ( 70 ) including:
an annular wall ( 80 ),
an external element ( 84 ) located radially outward of the annular wall ( 80 ),
a first space ( 72 ) for air circulation, the first space ( 72 ) located radially inward of the annular wall ( 80 ),
a second space ( 82 ) located radially between the annular wall ( 80 ) and the external element ( 84 ),
a first air introduction swirler ( 74 ) enabling air located outside the hollow central body ( 70 ) to penetrate inside the first space ( 72 ),
a fuel circuit ( 92 , 93 , 82 ) arranged at least partly about the first space ( 72 ), the fuel circuit ( 92 , 93 , 82 ) comprising the second space ( 82 ) for circulating a fuel film and fuel supply means ( 92 , 93 ) communicating with said second space ( 82 ), the second space ( 82 ) being annular, and
fuel injection ports ( 90 ) communicating the fuel circuit ( 92 , 93 , 82 ) with said first space ( 72 ), wherein the fuel injection ports ( 90 ) open into the first space ( 72 ) downstream from the first air introduction swirler ( 74 ); and
an external body ( 62 ) including a second air introduction swirler ( 64 ) that opens into an annular channel ( 66 ) externally delimited by an aerodynamic bowl ( 40 ), the aerodynamic bowl ( 40 ) having a downstream flared end ( 42 ),
wherein the second air introduction swirler ( 64 ) is an axial swirler arranged at an upstream end of the external body ( 62 ),
wherein the first and second spaces ( 72 , 82 ) of the hollow central body ( 70 ) open into a third space ( 86 ) for an air-fuel mixture, the third space ( 86 ) being arranged downstream of the second space ( 82 ) in order to inject the fuel film into the third space ( 86 ),
wherein the third space ( 86 ) extends to a downstream end of the hollow central body ( 70 ) and then is flared to extend up to the downstream flared end ( 42 ) of the aerodynamic bowl ( 40 ), and
wherein the fuel injection ports ( 90 ) are configured such that a portion of fuel in the second space ( 82 ) is diverted from the second space ( 82 ) through the fuel injection ports ( 90 ) into the first space ( 72 ), while another portion of the fuel travels through the second space ( 82 ) and is discharged from a downstream end of the second space ( 82 ) into the third space ( 86 ).
2. The aerodynamic injection system according to claim 1 , wherein the fuel injection ports ( 90 ) pass through said annular wall ( 80 ).
3. The aerodynamic injection system according to claim 1 , wherein the fuel injection ports ( 90 ) are distributed in circumferential rows about a central axis ( 22 ) of the aerodynamic injection system, said rows being axially aligned or staggeredly arranged.
4. The aerodynamic injection system according to claim 1 , wherein the fuel injection ports ( 90 ) each have an outlet axis ( 90 ) included in an axial plane of the injection system.
5. The aerodynamic injection system according to claim 1 , wherein the fuel injection ports ( 90 ) each have an outlet axis ( 96 ) having a non-zero circumferential component ( 98 ), so as to generate a spinning of a fuel flow rate penetrating said first space ( 72 ).
6. The aerodynamic injection system according to claim 1 , wherein an outlet axis ( 96 ) of each of the fuel injection ports ( 90 ) is tilted so as to inject fuel downstream inside said first space ( 72 ).
7. The aerodynamic injection system according to claim 1 , wherein the first air introduction swirler ( 74 ) includes vanes ( 76 ).
8. The aerodynamic injection system according to claim 7 , wherein said fuel supply means ( 92 , 93 ) comprise a plurality of ducts ( 93 ).
9. The aerodynamic injection system according to claim 8 , wherein the plurality of ducts ( 93 ) are arranged alternately with said vanes ( 76 ) of the first air introduction swirler ( 74 ), along a circumferential direction (Dc).
10. The aerodynamic injection system according to claim 7 , wherein the vanes ( 76 ) are tilted so as to generate a spinning of an air flow penetrating said first space ( 72 ).
11. An aircraft turbomachine ( 1 ) comprising the aerodynamic injection system ( 30 ) according to claim 1 .
12. The aerodynamic injection system according to claim 1 , wherein the first air introduction swirler ( 74 ) opens into a first portion of the first space ( 72 ), and the fuel injection ports ( 90 ) open into a second portion of the first space ( 72 ) that is downstream from the first portion.
13. The aerodynamic injection system according to claim 1 , wherein the aerodynamic bowl ( 40 ) is flared in a downstream direction such that the aerodynamic bowl ( 40 ) progressively expands radially in the downstream direction to the downstream flared end ( 42 ).