COMBUSTION CHAMBER LINER WITH SPIRAL COOLING CHANNELS
A combustion chamber liner comprising a plurality of spiral cooling channels is formed from a material component. A combustion chamber liner body extends from a first end and a second end. The combustion chamber liner body comprises a combustion chamber liner internal wall that defines a combustion area cavity extending from the first and second end. The combustion chamber liner body also comprises a combustion chamber liner external wall that is opposite the internal wall. The combustion chamber liner body also defines an inlet port, a nozzle exit port opposite the inlet port, and a throat portion. Along the combustion chamber liner external wall, spiral cooling channels are cut into the external wall such that the spiral cooling channels extend between the first and second end.
1 . A combustion chamber liner for a rocket propulsion system, comprising:
a combustion chamber liner body extending between a first and second end;
a combustion chamber liner internal wall defined by the combustion chamber liner body, the combustion chamber liner internal wall defining a combustion area cavity extending between the first and second end;
a combustion chamber liner external wall defined by the combustion chamber liner body, the combustion chamber liner external wall opposite the internal wall;
an inlet port defined by the combustion chamber liner body at the first end;
a nozzle exit port defined by the combustion chamber liner body at the second end;
a throat portion defined by a portion of the combustion chamber liner body between the first and second end; and
a plurality of spiral cooling channels defined by the combustion chamber liner external wall, the plurality of spiral cooling channels extending between the first and second end.
2 . The combustion chamber liner of claim 1 , wherein each spiral cooling channel of the plurality of spiral cooling channels comprises a first chamfer at a first cooling channel end and a second chamfer at a second cooling channel end.
3 . The combustion chamber liner of claim 1 , wherein each spiral cooling channel of the plurality of spiral cooling channels is defined by a uniform width.
4 . The combustion chamber liner of claim 3 , wherein a distance along the circumference of the combustion chamber liner body between each spiral cooling channel of the plurality of spiral cooling channels is at least equal to the uniform width.
5 . The combustion chamber liner of claim 1 , wherein the combustion chamber liner body comprises a material chosen from a group consisting of copper, a copper alloy, titanium, carbon fiber, and Inconel.
6 . The combustion chamber liner of claim 1 , wherein a first portion of the combustion area cavity at the throat portion has a first area that is less than a second area of a second portion of the combustion area cavity at the first end.
7 . A method for fabricating a combustion chamber liner for a rocket propulsion system, comprising:
forming a material component into a combustion chamber liner body extending between a first and second end, wherein the combustion chamber body comprises:
a combustion chamber liner internal wall defined by the combustion chamber liner body, the combustion chamber liner internal wall defining a combustion area cavity extending between the first and second end;
a combustion chamber liner external wall defined by the combustion chamber liner body, the combustion chamber liner external wall opposite the internal wall;
an inlet port defined by the combustion chamber liner body at the first end;
a nozzle exit port defined by the combustion chamber liner body at the second end; and
a throat portion defined by a portion of the combustion chamber liner body between the first and second end; and
cutting, into the combustion chamber liner external wall, a plurality of spiral cooling channels extending between the first and second end.
8 . The method of claim 7 , wherein the combustion chamber body is formed using a computer numerical control machine programmed based at least in part on a digital representation of the combustion chamber body.
9 . The method of claim 7 , wherein the plurality of spiral cooling channels are cut into the combustion chamber liner external wall such that the plurality of spiral cooling channels are defined such that each spiral cooling channel has a uniform width.
10 . The method of claim 9 , wherein a distance along a circumference of the combustion chamber liner body between each spiral cooling channel of the plurality of spiral cooling channels is at least equal to the uniform width.
11 . The method of claim 7 , wherein cutting the plurality of spiral cooling channels results in a first chamfer at a first cooling channel end of the plurality of spiral cooling channels and a second chamfer at a second cooling channel end of the plurality of spiral cooling channels.
12 . The method of claim 7 , wherein the material component is selected from a group consisting of copper, a copper alloy, titanium, carbon fiber, and Inconel.
13 . The method of claim 7 , wherein the combustion chamber body is formed such that a first portion of the combustion area cavity at the throat portion has a first area that is less than a second area of a second portion of the combustion area cavity at the first end.