阅读说明：本技术 具有螺旋冷却通道的燃烧室衬套 (Combustor liner with spiral cooling channels ) 是由 J·索恩伯格 于 2019-12-11 设计创作，主要内容包括：一种包括多个螺旋冷却通道的燃烧室衬套由材料部件形成。燃烧室衬套主体从第一端和第二端延伸。燃烧室衬套主体包括燃烧室衬套内壁,燃烧室衬套内壁限定从第一端和第二端延伸的燃烧区域腔体。燃烧室衬套主体还包括与内壁相对的燃烧室衬套外壁。燃烧室衬套主体还限定入口端口、与入口端口相对的喷嘴出口端口以及喉部。沿着燃烧室衬套外壁,将螺旋冷却通道切入外壁中,使得螺旋冷却通道在第一端和第二端之间延伸。(A combustor liner including a plurality of spiral cooling passages is formed from a material component. The combustor liner body extends from a first end and a second end. The combustor liner body includes a combustor liner inner wall defining a combustion zone cavity extending from a first end and a second end. The combustor liner body also includes a combustor liner outer wall opposite the inner wall. The combustor liner body also defines an inlet port, a nozzle outlet port opposite the inlet port, and a throat. A helical cooling channel is cut into the outer wall along the combustor liner outer wall such that the helical cooling channel extends between a first end and a second end.)

1. A combustor liner for a rocket propulsion system, comprising:

a combustor liner body extending between a first end and a second end;

a combustion liner inner wall defined by the combustion liner body, the combustion liner inner wall defining a combustion zone cavity extending between the first end and the second end;

a combustor liner outer wall defined by the combustor liner body, the combustor liner outer wall being opposite the combustor liner inner wall;

an inlet port defined by the combustor liner body at the first end;

a nozzle outlet port defined by the combustor liner body at the second end;

a throat defined by a portion of the combustor liner body between the first end and the second end; and

a plurality of spiral cooling passages defined by the combustor liner outer wall, the plurality of spiral cooling passages extending between the first end and the second end.

2. The combustion liner of claim 1, wherein each of the plurality of helical cooling passages includes a first chamfer at a first cooling passage end and a second chamfer at a second cooling passage end.

3. The combustion liner of claim 1, wherein each of the plurality of helical cooling channels is defined by a uniform width.

4. The combustion liner of claim 3, wherein a distance between each of the plurality of helical cooling channels along the circumference of the combustion liner body is at least equal to the uniform width.

5. The combustor liner of claim 1, wherein said combustor liner body comprises a material selected from the group consisting of copper, copper alloys, titanium, carbon fiber, and inconel.

6. The combustion liner of claim 1, wherein a first area of a first portion of the combustion zone cavity at the throat is less than a second area of a second portion of the combustion zone cavity at the first end.

7. A method of manufacturing a combustor liner for a rocket propulsion system, comprising:

forming a material component into a combustion liner body extending between a first end and a second end, wherein the combustion liner body comprises:

a combustion liner inner wall defined by the combustion liner body, the combustion liner inner wall defining a combustion zone cavity extending between the first end and the second end;

a combustor liner outer wall defined by the combustor liner body, the combustor liner outer wall being opposite the combustor liner inner wall;

an inlet port defined by the combustor liner body at the first end;

a nozzle outlet port defined by the combustor liner body at the second end; and

a throat defined by a portion of the combustor liner body between the first end and the second end; and

cutting a plurality of helical cooling channels extending between the first end and the second end into the combustor liner outer wall.

8. The method of claim 7, wherein the combustor body is formed using a computer numerically controlled machine programmed based at least in part on a digital representation of the combustor body.

9. The method of claim 7, wherein the plurality of helical cooling channels are cut into the combustor liner outer wall such that the plurality of helical cooling channels are defined such that each helical cooling channel has a uniform width.

10. The method of claim 9, wherein a distance between each of the plurality of helical cooling channels along the circumference of the combustor liner body is at least equal to the uniform width.

11. The method of claim 7, wherein cutting the plurality of helical cooling channels results in a first chamfer at a first cooling channel end of the plurality of helical cooling channels and a second chamfer at a second cooling channel end of the plurality of helical cooling channels.

12. The method of claim 7, wherein the material component is selected from the group consisting of copper, copper alloys, titanium, carbon fiber, and inconel.

13. The method of claim 7, wherein the combustion chamber body is formed such that a first area of a first portion of the combustion region cavity at the throat is less than a second area of a second portion of the combustion region cavity at the first end.

Background

Combustion chambers are commonly used in rocket engines to combust a propellant to generate thrust for a rocket propulsion system. The combustion of the propellant produces hot exhaust gases which, when passed through the nozzle of the combustion chamber, produce the thrust force. However, the temperatures and pressures generated by the combustion process may be extreme, requiring cooling of the combustion chamber to prevent damage or complete thermal destruction of the combustion chamber. Exacerbating this is the general requirement to optimize rocket propulsion systems to reduce weight. It is therefore difficult and expensive to employ complex rocket propulsion systems that provide optimal cooling of the combustion chamber and are also optimized based on weight considerations.

Drawings

Various techniques will be described with reference to the accompanying drawings, in which:

FIG. 1 is a perspective view of a combustor liner including a plurality of spiral cooling channels along a combustor liner outer wall of the combustor liner;

FIG. 2 is a perspective view of a combustor liner showing a combustor liner outer wall of the combustor liner and a plurality of spiral cooling channels extending from a nozzle of the combustor liner to an inlet port of the combustor liner;

FIG. 3 is a longitudinal cross-sectional view of the combustor liner of FIG. 2, illustrating the spiral cooling channels along the combustor liner outer wall of the combustor liner and the combustor liner inner wall of the combustor liner;

FIG. 4 is a perspective view of the combustor liner showing the combustor liner inner wall of the combustor liner from the inlet port to the throat of the combustor liner;

FIG. 5 is a transverse cross-sectional view of the combustor liner from the throat of the combustor liner, illustrating a plurality of spiral cooling channels along the combustor liner outer wall of the combustor liner.

Detailed Description

The techniques and systems described below relate to a combustor liner that, in operation, regeneratively cools a combustor using a helical cooling chamber. In one example, a combustion liner for a rocket propulsion system is made using a thermally conductive material such as copper. The combustion liner may include inlet ports through which a propellant may be introduced into the combustion chamber to combust and generate hot exhaust gases that may be used to generate thrust. Further, the combustion liner may include a nozzle through which hot gases generated by combustion of the propellant are accelerated to generate thrust for the rocket propulsion system. The nozzle may include a converging portion followed by a diverging portion toward an end opposite the inlet port of the combustor liner. At the converging portion of the combustor liner, between the inlet port and the end of the nozzle may be a throat where the flow of hot gases is choked such that as the hot gases exit the throat, the velocity of the hot gases increases as the nozzle area diverges toward the end opposite the inlet port (for supersonic flow). This may be described as isentropic expansion of the hot gas flow through the nozzle from the throat to the nozzle outlet port opposite the inlet port of the combustion liner.

The pressure and temperature within the combustion chamber liner may be abnormally high due to the combustion of the propellant within the combustion chamber. As a result, the combustor liner may experience significant thermal expansion at the inner wall, while the outer wall of the combustor liner may limit the thermal expansion due to the temperature gradient between the inner and outer walls. This temperature gradient may result in thermal stresses that may affect the life of the combustor liner. To prevent structural degradation of the combustor liner due to stresses caused by high temperatures and pressures within the combustor, the combustor liner may include a series of cooling channels along an outer wall of the combustor liner that extend along a circumference of the combustor liner from an inlet port to a nozzle outlet port. Through these cooling chambers, a propellant or other low temperature material may be introduced to provide regenerative cooling to the combustion liner and thereby reduce thermal stresses on the combustion liner.

In one example, each cooling channel along the outer wall of the combustor liner is cut into the outer wall in a spiral manner such that the depth, width, and length of the cooling channel are uniform along the outer wall. The cooling channels may be spaced a uniform distance from each other across the outer wall of the combustor liner. This may improve the efficiency of the regenerative cooling mechanism as the propellant or other cryogenic material flows more evenly through the cooling channel across the span of the outer wall from the inlet port to the nozzle outlet port. Furthermore, because the size of the cooling passage is consistent across the span of the outer wall of the combustion liner, the pressure encountered within the cooling passage is less likely to drop suddenly, which may result in an increase in the efficiency of the fuel pump that passes propellant or other low temperature material through the cooling passage.

In one example, the cooling passages each include a first chamfer near an inlet port of the combustor liner and a second chamfer near an outlet port of the nozzle. These chamfers may provide a gradient in wall thickness of the combustor liner that begins near the inlet port of the combustor liner and near the nozzle outlet port, and the wall thickness decreases until a minimum wall thickness is reached at the throat of the combustor liner. This may result in a smaller temperature gradient at the throat of the combustor liner, where temperature and pressure may be at their maximum during operation of the rocket propulsion system. In addition, the gradient may reduce the pressure drop along the cooling channel as the propellant or other cryogenic material passes through the cooling channel.

In the foregoing and following description, various techniques are described. For purposes of explanation, specific configurations and details are set forth in order to provide a thorough understanding of possible ways in which the techniques may be implemented. However, it will also be apparent that the techniques described below may be practiced in different configurations without specific details. In addition, well-known features may be omitted or simplified in order not to obscure the described technology.

The technology described and suggested in this disclosure improves the field of rocket propulsion systems, particularly the field of combustor liner implementations, by thermal stresses on the combustor liner during operation of the rocket propulsion system. Additionally, the techniques described and suggested in this disclosure improve the efficiency of rocket propulsion systems by reducing thermal gradients along the wall of the combustion chamber liner, thereby improving the performance of fuel pumps and other components of the rocket propulsion systems. Furthermore, the techniques described and suggested in this disclosure must be rooted in rocket propulsion technology in order to overcome problems that arise particularly in the combustion of propellants and other chemicals.

FIG. 1 is a perspective view of a combustor liner 10 including a plurality of spiral cooling channels 116 along a combustor liner outer wall 104 of the combustor liner 10. The combustor liner 10 may be fabricated using any structurally sound material, preferably copper, copper alloys, titanium, carbon fiber, Inconel (Inconel), or any other material capable of withstanding hoop stresses caused by high pressures during rocket propulsion system operation and also capable of withstanding thermal stresses caused by temperature gradients along the wall 104 of the combustor liner 10. The combustion liner 10 may be manufactured using a Computer Numerical Control (CNC) tool that is capable of creating an overall design of the combustion liner 10 and forming the spiral cut cooling channels 116 along the combustion liner outer wall 104 of the combustion liner 10 as described herein.

In one embodiment, the combustion liner 10 is introduced into a CNC machine to cut the spiral cut cooling channels 116 into the combustion liner outer wall 104. The CNC machine may have several axes of motion and use a laser or water jet to create spiral cut cooling channels 116 along the periphery of the combustor liner outer wall 104. For example, a CNC machine may be capable of creating spiral cut cooling channels 116 along three linear axes and two rotational axes. A digitized model of the combustion liner 10 may be created, which may include a graphical representation of the spiral cut cooling channels 116 along the graphical representation of the combustion liner 10 and the combustion liner outer wall 104. The digitized model may be converted into program code that may be executed by a CNC machine to engage a laser or water knife to cut the spiral cut cooling channels 116 onto the combustor liner 10 introduced into the machine. Alternatively, a solid material may be introduced that may be used by a CNC machine to create the combustor liner 10 and spiral cut cooling channels 116 in one process. In some embodiments, the spiral cut cooling channels 116 are defined by encoding a mathematical representation of the spiral pattern, which the CNC machine can use to define and create the spiral cut cooling channels 116.

In one embodiment, the combustion liner 10 includes an inlet port 110, a throat 114, and a nozzle 112, which may collectively surround a combustion zone cavity 122 of the combustion liner 10. To generate thrust for the rocket propulsion system, propellant may be introduced into the combustion zone cavity 122 via the inlet port 110 using an injector or other means for injecting propellant into the combustion zone cavity 122. Within the combustion zone cavity 122, the propellant is combusted to produce hot exhaust gases that can be used to generate thrust for the rocket propulsion system. For example, the thermal energy generated by the combustion of the propellant is converted to a large extent into kinetic energy. The combustor liner 10 may be manufactured to converge from the inlet port 110 to the throat 114 such that the cross-sectional area of the combustion zone cavity 122 decreases from the inlet port of the combustor liner 10 to the throat 114. Further aft from the throat 114, the cross-sectional area of the combustion zone cavity 122 may increase entirely across the nozzle to the nozzle outlet port 112. Due to the constriction of the combustion zone cavity 122 by the throat 114, the throat 114 of the combustion liner 10 may experience the maximum airflow per unit area and the highest pressure and temperature gradients of the combustion liner 10.

For rocket propulsion systems, it may be desirable for the velocity of the gas through the throat 114 to be sonic. Thus, as the hot gases exit the throat 114 and enter the diverging portion of the nozzle toward the nozzle outlet port 112, the pressure may decrease, resulting in an increase in velocity at supersonic conditions. This generates thrust for the rocket propulsion system, since a conversion of enthalpy into kinetic energy is achieved. The combustion of the propellant used to generate the thrust of the rocket propulsion system may result in high temperatures within the combustion zone cavity 122, and thus the combustion liner inner wall 102 of the combustion liner 10 may be subjected to significant temperatures and pressures during this process. Because the highest temperature and highest pressure of the airflow through the combustion zone cavity 122 of the combustor liner 10 is located within the throat 114, the combustor liner inner wall 102 of the combustor liner 10 may experience the greatest thermal and hoop stresses at the throat 114.

To mitigate the effects of thermal and hoop stresses on the combustor liner 10, regenerative cooling of the combustor liner 10 may be used. In one embodiment, the combustion liner 10 includes a plurality of cooling channels 116 cut into the combustion liner outer wall 104 of the combustion liner 10 to enable introduction of a propellant or other low temperature material to reduce a temperature gradient between the combustion liner inner wall 102 of the combustion liner 10 and the combustion liner outer wall 104 of the combustion liner 10. These cooling channels 116 may enable convective heat transfer so that the propellant may absorb heat from the combustion liner 10 to increase the initial temperature of the propellant and increase its energy level before passing through the injector and into the combustion zone cavity 122 through the inlet port 110.

In one embodiment, the cooling passage 116 is designed to minimize the rate of bubble formation and bubble size due to heat transfer from the combustor liner inner wall 102 of the combustor liner 10 to the combustor liner outer wall 104 of the combustor liner 10. As is known in the art, an excessive bubble formation rate may result in an unstable gas film along the combustor liner outer wall 104 of the combustor liner 10, resulting in an insulating layer that may cause rapid wall temperature increases. Such an increase in wall temperature may result in damage to the wall material and the combustor liner 10 itself.

In one embodiment, to reduce the rate of bubble formation within the cooling channel 116, the cooling channel 116 is cut into the combustor liner outer wall 104 of the combustor liner 10 in a spiral pattern along the circumference of the combustor liner outer wall 104 from the front of the inlet port 110 to the nozzle outlet port 112. This may ensure that the cooling passages 116 share a uniform length along the combustor liner outer wall 104 of the combustor liner 10 and reduce the amount of gradient along the combustor liner outer wall 104 from the inlet port 110, through the throat 114, and to the nozzle outlet port 112. Further, the spiral pattern may allow the cooling channels 116 to share a uniform width and depth within the combustor liner outer wall 104, thereby minimizing the potential for pressure drop along the combustor liner 10. The spiral pattern for the cooling channels 116 may also promote uniform flow of propellant through the cooling channels 116, thereby reducing turbulence and the occurrence of hot spots throughout the span of the cooling channels 116. The cooling channels 116 may be cut into the combustor liner outer wall 104 of the combustor liner 10 using CNC tools or other computer aided manufacturing tools.

In one embodiment, each cooling passage 116 of the plurality of cooling passages includes an inlet port chamfer 106 and a nozzle outlet port chamfer 108 proximate to the inlet port 110 of the combustion liner 10 and the nozzle outlet port 112 of the combustion liner 10, respectively. These chamfers 106, 108 may gradually introduce propellant into the cooling gallery 116 and enable propellant to flow out of the cooling gallery 116 smoothly. The inlet port chamfer 106 may begin at the full thickness of the combustor liner body 124 and cut into the combustor liner outer wall 104 of the combustor liner 10 at an angle until it reaches the full depth of the cooling channel 116, where the thickness of the combustor liner body 124 may be minimized. In one embodiment, each cooling passage 116 of the plurality of cooling passages has a uniform width along the length of the combustor liner 10. Further, in one embodiment, the distance between the cooling channels along the circumference of the combustor liner body 124 is at least equal to the uniform width of each cooling channel 116 of the plurality of cooling channels.

The combustor liner inner wall 102 of the combustor liner 10 may be smooth (e.g., without protrusions, bumps, ridges, or other non-planar elements that may prevent the presence of a planar surface) to prevent any disturbance to the gas flow through the combustion zone cavity 122 as it passes through the throat 114 and out of the nozzle outlet port 112. Thus, the combustor liner inner wall 102 may not have any obstructions that may cause perturbations in the gas flow through the combustion zone cavity 122 that may cause the downstream pressure to rise above the critical pressure value of the combustor liner 10. The combustor liner inner wall 102 and the combustor liner outer wall 104 may be separated by a wall thickness of the combustor liner body 124.

FIG. 2 is a perspective view of combustor liner 10, illustrating combustor liner outer wall 104 of combustor liner 10 and a plurality of spiral cooling channels 116 extending from nozzle outlet port 112 of combustor liner 10 to inlet port 110 of combustor liner 10. In the embodiment shown in FIG. 2, the combustion liner 10 includes a combustion liner outer wall 104 that circumferentially surrounds a combustion zone cavity within the combustion liner 10. The combustor liner outer wall 104 may narrow from the inlet port 110 at the throat 114 and then expand to the nozzle outlet port 112. In one embodiment, the combustor liner outer wall 104 is a combustor liner body, shown in detail in FIG. 1 above, which may include the combustor liner outer wall 104 and the combustor liner inner wall 102 separated from the combustor liner outer wall 104 by a wall thickness. Combustor liner inner wall 102 may circumferentially surround combustion zone cavity 122 within combustor liner 10.

In one embodiment, the combustor liner outer wall 104 of the combustor liner body 124 includes a plurality of spiral cut cooling channels 116 through which a propellant or other low temperature material may be introduced to reduce thermal and hoop stresses of the combustor liner 10 during operation. The spiral cut cooling channel 116 may cut into the combustion liner body 124 at the combustion liner outer wall 104, thereby reducing the thickness of the combustion liner body 124 from the combustion liner outer wall 104 until the cooling channel wall thickness is reached. In one embodiment, the spiral cut cooling channels 116 begin at an angle from the inlet port 110 and travel in a spiral pattern along the span and circumference of the combustion liner body 124 and terminate at another angle relative to the nozzle outlet port 112. The spiral pattern selected for each spiral cut cooling channel 116 may be selected to maintain a uniform width for each cooling channel 116 in the plurality of cooling channels. Further, the number of spiral cut cooling channels 116 may be selected such that along the circumference of the combustor liner body 124, the distance between the cooling channels 116 along the circumference of the combustor liner body 124 is at least equal to the uniform width of each cooling channel 116 of the plurality of cooling channels. The thickness of the combustor liner body 124 between the spiral cut cooling channels 116 may be at its maximum thickness.

In one embodiment, the cooling channel wall thickness varies along the length of each cooling channel 116, with a minimum thickness at the throat 114. For example, the cooling passage 116 may be cut into the combustor liner outer wall 104 of the combustor liner body 124 at a chamfer that begins at a distance from the inlet port 110 and extends to the throat 114, with the cooling passage wall thickness gradually decreasing. The gradient of cooling passage wall thickness created using the chamfer may be determined based at least in part on predicted or analyzed thermal stresses and temperature gradients along the circumference and span of the combustor liner 10. In one embodiment, the cooling passage 116 may also be cut into the combustor liner outer wall 104 of the combustor liner body 124 with another chamfer beginning at a distance from the nozzle outlet port 112 and extending to the throat 114 and the cooling passage wall thickness gradually decreasing with a minimum cooling passage wall thickness at the throat 114. In some embodiments, the angle of any chamfer may be different such that there is a minimum cooling channel wall thickness along a greater portion of the span of the cooling channel 116.

FIG. 3 is a longitudinal cross-sectional view of combustor liner 10 of FIG. 2, illustrating spiral cut cooling channels 116 along combustor liner outer wall 104 of combustor liner 10 and combustor liner inner wall 102 of combustor liner 10. In the embodiment shown in FIG. 3, the combustor liner inner wall 102 of the combustor liner body 124 is smooth, without protrusions, bumps, or other rough surfaces that may impair hot gas flow through the combustion zone cavity 122 of the combustor liner 10. The combustor liner outer wall 104 of the combustor liner body 124 may include a plurality of cooling channels 116 that may be used to convey a propellant or other low temperature material to reduce thermal stresses on the combustor liner body 124 during combustion of the propellant within the combustion zone cavity 122 inside the combustor liner 10.

The combustor liner body 124 may have an initial thickness 126 at the inlet port 110 and the nozzle outlet port 112 that is equal to the maximum allowable thickness of the combustor liner 10. The initial thickness 126 may be selected based at least in part on design considerations (e.g., weight, thermal conductivity, heat dissipation, etc.) of the combustor liner 10. In one embodiment, based at least in part on design considerations of the inlet port 110, a spiral cut cooling channel 116 is cut into the combustor liner outer wall 104 of the combustor liner body 124 at a distance from the inlet port 110 of the combustor liner 10. Further, in one embodiment, the spiral cut cooling channels 116 terminate a distance from the nozzle outlet port 112 of the combustion liner 10. The distance may be selected based at least in part on design considerations of the nozzle outlet port 112. These design considerations may be similar to the overall design considerations of the combustor liner 10 as described above.

In one embodiment, the cooling passage 116 is cut such that the cooling passage wall thickness 120 is less than the initial thickness 126 of the combustor liner body 124. The cooling channel wall thickness 120 may be uniform throughout the length of each spiral cut cooling channel 116. Alternatively, each cooling passage 116 may begin and end with a chamfer, which may result in a gradual decrease in combustor liner thickness along the span of the combustor liner 10 from an initial thickness 126 to a cooling passage wall thickness 120 along the length or a portion of the length of the cooling passage 116. The angles of these chamfers may be determined based at least in part on design considerations of the combustor liner 10, and preferably are determined based at least in part on determined thermal stresses along the span of the combustor liner 10.

As described above, cooling channels 116 are cut into the combustion liner outer wall 104 of the combustion liner body in a spiral pattern along the circumference of the combustion liner 10 from the inlet port 110 to the nozzle outlet port 112. The spiral pattern for the cooling passage 116 may be selected such that the combustor liner outer wall 104 of the combustor liner 10 has a cooling passage wall thickness 120 over a greater area of the throat 114. Since the throat 114 of the combustion liner 10 may exhibit the highest thermal and hoop stresses caused by combustion of the propellant in the combustion zone cavity 122, the increased area at the throat 114 with the cooling gallery wall thickness 120 may result in a greater reduction in these stresses. However, each cooling channel 116 may have a uniform width and depth along the length of the cooling channel 116. Thus, the longitudinal cross-sectional area of the combustor liner 10 may exhibit a larger area at the throat 114 where the combustor liner body thickness may be at its minimum.

FIG. 4 is a perspective view of the combustor liner 10 illustrating the combustor liner inner wall 102 of the combustor liner from the inlet port 110 to the throat 114 of the combustor liner 10. As described above, the combustion liner 10 may be constructed using a circumferential combustion liner body 124 that surrounds the combustion zone cavity 122 for propellant combustion. At the inlet port 110, the combustor liner body 124 may have an initial thickness 126 between the combustor liner inner wall 102 of the combustor liner 10 and the combustor liner outer wall 104 of the combustor liner 10. In one embodiment, at the inlet port 110 of the combustion liner 10, the cooling channels 116 are not present, as these cooling channels may be cut into the combustion liner outer wall 104 of the combustion liner body 124 at a distance from the inlet port 110 of the combustion liner 10 as described above.

The combustion liner inner wall 102 of the combustion liner body 124 may be smooth, free of protrusions, rough edges, bumps, etc., to enable an undisturbed flow of exhaust through the combustion zone cavity 122 during operation of the rocket propulsion system. Further, the combustor liner body 124 may converge from an initial cross-sectional diameter at the inlet port 110 to a minimum cross-sectional diameter at the throat 114 of the combustor liner 10 along the longitudinal span of the combustor liner 10. The amount of diameter taper along the longitudinal span of the combustor liner 10 may be determined based at least in part on desired exhaust flow conditions through the combustion zone cavity 122 of the combustor liner 10. For example, the diameter taper along the longitudinal span of the combustor liner 10 may be selected such that the exhaust flow through the throat 114 of the combustor liner 10 is sonic. This may result in the exhaust flow being supersonic at the nozzle and nozzle outlet port 112 aft of the throat 114, as the cross-sectional diameter aft of the throat 114 may expand to a maximum at the nozzle outlet port 112.

In one embodiment, the initial thickness 126 between the combustor liner inner wall 102 of the combustor liner 10 and the combustor liner outer wall 104 of the combustor liner 10 at the inlet port 110 is the maximum combustor liner body thickness of the combustor liner 10. At a distance from the inlet port 110 of the combustion liner 10, the combustion liner outer wall 104 of the combustion liner 10 is cut to introduce the cooling channel 116 along the circumference of the combustion liner outer wall 104 of the combustion liner body 124. In one embodiment, the cooling channel 116 is cut into the combustor liner outer wall 104 of the combustor liner 10 at a chamfer to gradually reduce the thickness of the combustor liner body 124 along the length of the cooling channel 116 from an initial thickness 126 to a minimum cooling channel wall thickness 120, for example, at the throat 114 of the combustor liner 10. The angle of the chamfer may be selected based at least in part on predicted thermal stresses along the longitudinal span of the combustor liner 10 and along the length of each cooling passage 116. Thus, based at least in part on these predicted thermal stresses, different fillets and thicknesses may be selected for each cooling passage.

FIG. 5 is a transverse cross-sectional view of combustor liner 10 from throat 114 of combustor liner 10, illustrating a plurality of spiral cut cooling channels 116 along combustor liner outer wall 104 of combustor liner body 124. At the throat 114, the cooling passage wall thickness 120 of each cooling passage 116 may be minimal, as this may alleviate the extreme thermal stresses experienced at the throat 114 when propellant or other low temperature material is pumped through the cooling passage 116 during operation of the rocket propulsion system. In one embodiment, the cooling channels 116 cut along the combustor liner outer wall 104 of the combustor liner body 124 may each be separated from the inlet port 110 by the combustor liner body 124 having an initial thickness 126, as described above. In some cases, the spacing between each cooling passage 116 may have a thickness that is different than the initial thickness 126. However, the different thickness may still be greater than the cooling channel wall thickness 120 of the cooling channel 116.

Each cooling channel 116 may have a uniform cooling channel width 118. The actual cooling channel width 118 may be selected based at least in part on various design considerations including, but not limited to, the amount of surface area along the circumference of the combustion liner 10 needed to mitigate the effects on thermal stress of the combustion liner 10 caused by combustion of the propellant within the combustion zone cavity 122. Further, as mentioned above, the distance between the cooling passages along the circumference of the combustor liner body 124 may be at least equal to the uniform width of each cooling passage 116 of the plurality of cooling passages. Since the highest thermal stresses may be experienced at the throat 114 of the combustor liner 10, these thermal stresses may be used as a basis for selecting the desired cooling passage width 118. The remaining circumferential area is reserved for the remainder of the combustor liner body 124, which may have an initial thickness 126 or any other thickness greater than the cooling channel wall thickness 120.

In one embodiment, the spiral pattern of the cooling passage 116 is selected such that the cooling passage 116 exhibits a uniform cross-sectional width along the length of the throat 114 through the longitudinal span of the combustion liner 10 that is about the cooling passage width 118. Thus, the spiral cut cooling channel 116 may be parallel to the longitudinal axis of the combustion liner 10 along the throat 114 of the combustion liner 10. This may provide a greater degree of relief from thermal stresses along the throat 114 of the combustion liner 10, as the propellant or other low temperature material may act as a larger heat sink at the throat 114 due to the cooling passage 116 being in the direction of the throat 114.

The specification and drawings are, accordingly, to be regarded in an illustrative rather than a restrictive sense. It will, however, be evident that various modifications and changes may be made thereto without departing from the broader spirit and scope of the invention as set forth in the claims. Other variations are also within the spirit of the present disclosure. Accordingly, while the disclosed technology is susceptible to various modifications and alternative constructions, certain illustrated embodiments thereof are shown in the drawings and have been described above in detail. It should be understood, however, that there is no intention to limit the invention to the specific form or forms disclosed, but on the contrary, the intention is to cover all modifications, alternative constructions, and equivalents falling within the spirit and scope of the invention as defined by the appended claims.

Additionally, various embodiments of the present disclosure may be described in view of the following clauses:

1. a combustor liner for a rocket propulsion system, comprising:

a combustor liner body extending between a first end and a second end;

a combustion liner inner wall defined by the combustion liner body, the combustion liner inner wall defining a combustion zone cavity extending between the first end and the second end;

a combustor liner outer wall defined by the combustor liner body, the combustor liner outer wall being opposite the combustor liner inner wall;

an inlet port defined by the combustor liner body at the first end;

a nozzle outlet port defined by the combustor liner body at the second end;

a throat defined by a portion of the combustor liner body between the first end and the second end; and

a plurality of spiral cooling passages defined by the combustor liner outer wall, the plurality of spiral cooling passages extending between the first end and the second end.

2. The combustion liner of clause 1, wherein each of the plurality of helical cooling channels includes a first chamfer at a first cooling channel end and a second chamfer at a second cooling channel end.

3. The combustion liner of clause 1 or 2, wherein each helical cooling channel of the plurality of helical cooling channels is defined by a uniform width.

4. The combustion liner of clause 3, wherein a distance between each of the plurality of helical cooling channels along the circumference of the combustion liner body is at least equal to the uniform width.

5. The combustion liner of any of clauses 1-4, wherein the combustion liner body comprises a material selected from the group consisting of copper, copper alloys, titanium, carbon fiber, and inconel.

6. The combustion liner of any of clauses 1-5, wherein a first area of a first portion of the combustion zone cavity at the throat is less than a second area of a second portion of the combustion zone cavity at the first end.

7. A method of manufacturing a combustor liner for a rocket propulsion system, comprising:

forming a material component into a combustion liner body extending between a first end and a second end, wherein the combustion liner body comprises:

a combustion liner inner wall defined by the combustion liner body, the combustion liner inner wall defining a combustion zone cavity extending between the first end and the second end;

a combustor liner outer wall defined by the combustor liner body, the combustor liner outer wall being opposite the combustor liner inner wall;

an inlet port defined by the combustor liner body at the first end;

a nozzle outlet port defined by the combustor liner body at the second end; and

a throat defined by a portion of the combustor liner body between the first end and the second end; and

cutting a plurality of helical cooling channels extending between the first end and the second end into the combustor liner outer wall.

8. The method of clause 7, wherein the combustor body is formed using a computer numerically controlled machine programmed based at least in part on the digital representation of the combustor body.

9. The method of clause 7 or 8, wherein the plurality of helical cooling channels are cut into the combustor liner outer wall such that the plurality of helical cooling channels are defined such that each helical cooling channel has a uniform width.

10. The method of clause 9, wherein a distance between each of the plurality of helical cooling channels along the circumference of the combustor liner body is at least equal to the uniform width.

11. The method of any of clauses 7-10, wherein cutting the plurality of helical cooling channels results in a first chamfer at a first cooling channel end of the plurality of helical cooling channels and a second chamfer at a second cooling channel end of the plurality of helical cooling channels.

12. The method of any of clauses 7-11, wherein the material component is selected from the group consisting of copper, copper alloys, titanium, carbon fiber, and inconel.

13. The method of any of clauses 7-12, wherein the combustion chamber body is formed such that a first area of a first portion of the combustion zone cavity at the throat is less than a second area of a second portion of the combustion zone cavity at the first end.

The use of the terms "a" and "an" and "the" and similar referents in the context of describing the disclosed embodiments (especially in the context of the following claims) is to be construed to cover both the singular and the plural, unless otherwise indicated herein or clearly contradicted by context. The terms "comprising," "having," "including," and "containing" are to be construed as open-ended terms (i.e., meaning "including, but not limited to,") unless otherwise noted. The term "connected," without modification and referring to physical connection, is to be construed as being partially or wholly contained within, attached to, or joined together, even if intervening elements are present. Recitation of ranges of values herein are merely intended to serve as a shorthand method of referring individually to each separate value falling within the range, unless otherwise indicated herein, and each separate value is incorporated into the specification as if it were individually recited herein. Unless otherwise indicated or contradicted by context, use of the term "set" (e.g., "set of items") or "subset" should be interpreted as a non-empty set comprising one or more members. Furthermore, unless otherwise indicated or contradicted by context, the term "subset" of a corresponding set does not necessarily denote a proper subset of the corresponding set, but rather the subset and the corresponding set may be equal.

Conjunctive languages, such as phrases in the form of "A, B, and at least one of C" or "A, B and C," are to be understood in conjunction with the context, and are generally used to indicate that an item, etc. can be either a or B or C, or any non-empty subset of the a and B and C collections, unless specifically stated otherwise or otherwise clearly contradicted by context. For example, in an illustrative example of a set having three members, a conjunctive phrase of "at least one of A, B, and C" and "at least one of A, B and C" refers to any of the following sets: { A }, { B }, { C }, { A, B }, { A, C }, { B, C }, and { A, B, C }. Thus, such conjunctive language is generally not intended to imply that certain implementations require the respective presence of at least one of a, at least one of B, and at least one of C. In addition, the term "plurality" indicates a plural state (e.g., "a plurality of items" indicates a plurality of items) unless otherwise stated or contradicted by context. The number of items is at least two, but may be more if explicitly or by context.

The use of any and all examples, or exemplary language (e.g., "such as") provided, is intended merely to better illuminate embodiments of the invention and does not pose a limitation on the scope of the invention unless otherwise claimed. No language in the specification should be construed as indicating any non-claimed element as essential to the practice of the invention.

Embodiments of this disclosure are described, including the best mode known to the inventors for carrying out the invention. Variations of those embodiments may become apparent to those of ordinary skill in the art upon reading the foregoing description. The inventors expect skilled artisans to employ such variations as appropriate, and the inventors intend for the embodiments of the disclosure to be practiced otherwise than as specifically described. Accordingly, the scope of the present disclosure includes all modifications and equivalents of the subject matter recited in the claims appended hereto as permitted by applicable law. Moreover, although elements described above may be described in the context of certain embodiments of the specification, unless otherwise indicated herein or otherwise clearly contradicted by context, such elements are not mutually exclusive of only those embodiments in which they are described; any combination of the above-described elements in all possible variations thereof is encompassed by the disclosure unless otherwise indicated herein or otherwise clearly contradicted by context.

All references, including publications, patent applications, and patents, cited herein are hereby incorporated by reference to the same extent as if each reference were individually and specifically indicated to be incorporated by reference and were set forth in its entirety herein.