OPTIMIZING THE ANGULAR POSITIONING OF A TURBINE NOZZLE AT THE OUTLET FROM A TURBOMACHINE COMBUSTION CHAMBER

• US 20090266080A1
• Filed: 04/21/2009
• Published: 10/29/2009
• Est. Priority Date: 04/24/2008
• Status: Active Grant

• Current Assignee: Safran Aircraft Engine Company
• Original Assignee: Snecma (Safran SA)
• Inventors: SANDELIS, Denis Jean Maurice, PIEUSSERGUES, Christophe
• Granted Patent: US 8,215,118 B2
• Time in Patent Office: Days
• US Class Current: 60/746
• CPC Class Codes: F01D 9/023 Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
• CPC Class Codes: F05D 2250/311 the axes being in line
• CPC Class Codes: F23R 3/06 Arrangement of apertures along the flame tube
• CPC Class Codes: F23R 3/50 Combustion chambers comprising an annular flame tube within an annular casing toroidal combustion chambers F23R3/52
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A turbomachine including an annular combustion chamber fitted with fuel injectors and nozzle vanes arranged at the outlet from the chamber, the number of nozzle vanes being an integer multiple of the number of fuel injectors, and the head of each injector being situated angularly half-way between the leading edges of two consecutive nozzle vanes, these leading edges being in alignment with primary air holes and/or with dilution air holes.

Claims

1. A turbomachine including an annular combustion chamber comprising inner and outer coaxial walls connected together at their upstream ends by an annular chamber end wall, a nozzle comprising an annular row of stationary vanes situated at the outlet from the chamber, and an annular row of fuel injectors having heads in alignment with orifices in the chamber end wall, the number of nozzle vanes being equal to k times the number of injectors, where k is an integer, wherein the head of each injector lies in a plane that contains the axis of the turbomachine and that is half-way between the leading edges of two consecutive nozzle vanes, and that intersects at least one dilution air inlet orifice passing through the inner wall, and/or through at least one dilution air inlet orifice passing through the outer wall, primary air inlet orifices being formed through the chamber walls upstream from the dilution orifices and on either side of the planes passing via the injector heads.

2. A turbomachine according to claim 1, wherein the leading edge of each nozzle vane lies in a plane that contains the axis of the turbomachine and that intersects at least one primary air inlet orifice or at least one dilution air inlet orifice passing through a coaxial wall of the combustion chamber.

3. A turbomachine according to claim 2, wherein each plane containing a leading edge of a vane intersects a dilution air inlet orifice in the outer wall and/or via a dilution air inlet orifice through the inner wall, and is situated at one injector half-pitch from the head of an injector.

4. A turbomachine according to claim 2, wherein each plane containing a leading edge of a vane intersects a dilution air inlet orifice and a primary air inlet orifice through a coaxial wall of the combustion chamber, and is situated at one injector quarter-pitch from the head of an injector.