TURBOMACHINE COMBUSTION CHAMBER SHELL RING

• US 20150059344A1
• Filed: 05/23/2013
• Published: 03/05/2015
• Est. Priority Date: 05/25/2012
• Status: Active Grant

• Current Assignee: Safran Aircraft Engine Company
• Original Assignee: Snecma (Safran SA)
• Inventors: Sandelis, Denis Jean Maurice
• Granted Patent: US 9,377,200 B2
• Time in Patent Office: Days
• US Class Current: 60/722
• CPC Class Codes: F23R 2900/03041 Effusion cooled combustion chamber walls or domes
• CPC Class Codes: F23R 2900/03042 Film cooled combustion chamber walls or domes
• CPC Class Codes: F23R 2900/03044 Impingement cooled combustion chamber walls or subassemblies
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube

Abstract

A turbomachine combustion chamber shell ring in which dilution holes in the turbomachine combustion chamber shell ring are covered with inserts defining chambers around same on an inner face of the shell ring. Ventilation holes, through the insert, induce ventilation of portions of the shell ring surrounding the dilution holes, cool the portions, and prevent crack formation.

Claims

1-10. -10. (canceled)

11. A turbomachine combustion chamber shell ring, comprising:

dilution holes;

ventilation holes surrounding the dilution holes and being finer and more numerous than the dilution holes;

inserts extending over and around the dilution holes on an outer face of the shell ring;

wherein the shell ring is devoid of ventilation holes at portions situated above the inserts;

wherein the inserts each comprise an edge for attaching to the shell ring and an orifice extending over one of the respective dilution holes, and the inserts are traversed by holes directed towards the portions of the shell ring.

12. A shell ring according to claim 11, wherein main portions of the inserts are parallel with the shell ring.

13. A shell ring according to claim 11, wherein main portions of the inserts are inclined in relation to the shell ring in an axial direction of the shell ring.

14. A shell ring according to claim 11, wherein the holes of the inserts are perpendicular to the shell ring.

15. A shell ring according to claim 11, wherein the inserts extend more in a downstream direction of the shell ring than in other directions from centers of the dilution holes.

16. A shell ring according to claim 15, wherein the inserts retract in the downstream direction of the shell ring.

17. A shell ring according to claim 11, wherein the inserts each comprise an inner edge surrounding a respective of the orifices and extending towards a respective of the dilution holes.

18. A shell ring according to claim 17, wherein the inner edge is enclosed between attachment sectors projecting in the respective of the dilution holes, flow sectors being defined in the respective of the dilution holes by the inner edge and the attachment sectors.

19. A shell ring according to claim 18, wherein the dilution holes and the inner edges have centers offset in an axial direction of the shell ring, such that flow sectors have a main surface area downstream from the inner edges.

20. A turbomachine combustion chamber comprising:

a shell ring;

the shell ring comprising dilution holes and ventilation holes surrounding the dilution holes and being finer and more numerous than the dilution holes;

inserts extending over and around the dilution holes on an outer face of the shell ring;

wherein the shell ring is devoid of ventilation holes at portions situated above the inserts;

wherein the inserts each comprise an edge for attaching to the shell ring and an orifice extending over one of the respective dilution holes, and the inserts are traversed by holes directed towards the portions of the shell ring.