Combustion chamber for a turbine engine

1. A combustion chamber for a turbine engine such as an airplane turboprop or turbojet, the combustion chamber comprising:

inner and outer annular walls forming bodies of revolution that are connected together by an annular chamber end wall, wherein

the annular chamber end wall includes openings through which air enters the combustion chamber with a fuel injector head centered in each of the openings,

the inner and outer annular walls include primary air holes and dilution air holes,

the inner wall of the combustion chamber includes a single thickness of material that presents a thickness that varies along a longitudinal axis and in a circumferential direction of said inner wall,

the outer wall of the combustion chamber presents thickness that is substantially constant,

the thickness of the inner wall varies locally according to a group of zones, the zones being defined by relative temperature gradients subjected to the zones from combustion in the combustion chamber,

the inner wall includes a first zone of the group of zones that is a relatively hot zone of combustion at least partially surrounding a first portion of a perimeter of a first primary air hole of the primary air holes, wherein the first zone is subjected to a first temperature gradient and has a first thickness, and

the inner wall includes a second zone of the group of zones that is a relatively cool zone of combustion subjected to a second temperature gradient smaller than the first temperature gradient and that has a second thickness smaller than the first thickness, the second zone being continuous and at least partially surrounding the first zone such that the second zone is at least partially upstream and downstream from the first zone with regard to the openings, wherein the second zone at least partially surrounds a remaining portion of the perimeter of the first primary air hole.

2. The combustion chamber according to claim 1 , wherein the inner wall of the single thickness of material includes at least two adjacent zones of the group of zones made of different materials.

3. The combustion chamber according to claim 1 , wherein the inner wall of varying thickness is made by machining.

4. The combustion chamber according to claim 1 , wherein the inner wall of varying thickness is made by stretching and forming sheet metal.

5. A turbine engine such as an airplane turboprop or turbojet, the engine comprising a combustion chamber according to claim 1 .

6. The combustion chamber according to claim 1 , wherein the thickness of the inner wall increases locally with respect to a temperature increase of the respective zone of the group of zones.

7. The combustion chamber according to claim 1 , wherein the zones are situated between injectors, the zones including annular fastening flanges, and the zones including multiple perforations.