Combustion chamber for a turbine engine with homogeneous air intake through fuel injection system

Patent metadata

Field	Value
Publication number	US-10180256-B2
Application number	US-201415023954-A
Country	US
Kind code	B2
Filing date	Sep 29, 2014
Priority date	Oct 1, 2013
Publication date	Jan 15, 2019
Grant date	Jan 15, 2019

Abstract

A combustion chamber for a turbine engine, including an annular end wall provided with injection systems each centered on a respective axis and each having an upstream end forming a bushing for receiving a head of a fuel injector, and an annular shroud covering the end wall and including injector ports respectively arranged facing the injection systems, wherein the annular shroud includes air intake ports separate from the injector ports, and the bushing of each injection system crosses the corresponding injector port and includes at its upstream end a collar having a free end remote from the axis of the injection system by a first distance greater than a second distance separating a rim of the injector port from the axis.

First claim

The invention claimed is:

1. An annular combustion chamber for a turbine engine, comprising:
an annular end wall provided with a plurality of injection systems each centered on a respective axis and each having an upstream end forming a bushing for receiving a head of a fuel injector, a downstream end opening into said combustion chamber, and an annular air inlet arranged between said upstream and downstream ends so that the air admitted through said annular air inlet mixes, within the injection system, with the fuel coming from the fuel injector, and
an annular shroud covering an upstream side of said end wall and comprising a plurality of injector ports respectively arranged facing said injection systems, said annular shroud and said end wall delimiting together an annular space into which the annular air inlet of each injection system opens,
wherein said annular shroud includes a plurality of air intake ports separate from said injector ports, and
wherein said bushing of each of said injection systems crosses the corresponding injector port of said annular shroud and comprises, at an upstream end of the bushing, an annular collar having a free end remote from said axis of the injection system by a first distance greater than or equal to a second distance separating a rim of said corresponding injector port from said axis of the injection system.

2. The annular combustion chamber according to claim 1 , wherein said injector ports are distributed according to a plurality of pairs of consecutive injector ports so that at least one of the air intake ports is circumferentially arranged between each pair of consecutive injector ports along a circumference of said annular shroud.

3. The annular combustion chamber according to claim 2 , wherein said air intake ports are alternately distributed with said injector ports along the circumference of said annular shroud.

4. A combustion chamber module for a turbine engine, comprising:
the annular combustion chamber according to claim 1 , and
an annular row of fuel injectors comprising respective injector heads mounted respectively fitted in said bushings of the injection systems of said combustion chamber.

5. The combustion chamber module according to claim 4 , wherein each injector head includes a central nose for injecting fuel, an axial air intake device arranged around said central nose, and a peripheral fuel injection device arranged around said axial air intake device.

6. The combustion chamber module according to claim 4 , wherein said injector ports of said annular shroud have respective geometric center points inscribed on a first circle centered on an axis of said combustion chamber and having a first diameter.

7. The combustion chamber module according to claim 6 , wherein said air intake ports of said annular shroud have respective geometric center points inscribed on a second circle centered on the axis of said combustion chamber and having a second diameter strictly greater than said first diameter of said first circle.

8. The combustion chamber module according to claim 6 , wherein said air intake ports of said annular shroud have respective geometric center points inscribed on said first circle.

9. A turbine engine for an aircraft, comprising the combustion chamber module according to claim 4 .

Assignees

• Safran Aircraft Engines

Inventors

• Rullaud Matthieu Francois
• Lunel Romain Nicolas
• Noel Thomas Olivier Marie

Classifications

• F23R3/04

  Air inlet arrangements · CPC title

• F02C7/222

  Fuel flow conduits, e.g. manifolds · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F05D2220/30

  in turbines · CPC title

• F23R3/283

  Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances · CPC title