Double wall for a combustion chamber of a gas turbine of an aircraft and method for manufacturing thereof
By installing protruding components on the inner wall of the combustion chamber of the gas turbine in the aircraft and utilizing additive manufacturing technology, the problem of bridge fracture caused by thermal expansion of the inner and outer walls was solved, achieving efficient cooling and extended service life, and reducing manufacturing costs.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- SAFRAN HELICOPTER ENGINES
- Filing Date
- 2021-12-20
- Publication Date
- 2026-05-29
AI Technical Summary
Due to thermal expansion, the inner and outer walls of the combustion chamber of existing aircraft gas turbines are relatively displaced, causing bridge connections to break, cooling airflow to change direction, affecting the cooling effect of the inner wall and shortening its lifespan. Moreover, existing reinforcement connection methods are costly.
An inner wall protrusion extends into the outer wall port, defining the regulating flow cross-section, enhancing the cooling airflow exchange surface area, and ensuring precise component positioning and thermal expansion adaptation through additive manufacturing.
It improves the cooling effect of the combustion chamber wall, extends service life, avoids thermal gradients and heat conduction, and reduces manufacturing costs.
Smart Images
Figure CN116601437B_ABST
Abstract
Description
Technical Field
[0001] This invention relates to the field of aircraft gas turbine combustion chambers, particularly for helicopters. Background Technology
[0002] Referring to FIG. 1, in a known manner, the combustion chamber 101 includes double walls 102, namely an inner wall 121 in contact with the combustion reaction R and an outer wall 122 forming a heat shield. To limit heat propagation from the inner wall 121, it is known to provide ports 103 in the outer wall 122 to allow circulation of a cooling airflow F, which cools the inner wall 121 by impact and thus increases its lifespan. To enable the fabrication of the double walls 2 and ensure the spacing between the inner wall 121 and the outer wall 122, it is known to provide a bridge 104 connecting the inner wall 121 and the outer wall 122, as shown in FIG. 1. In practice, the bridge 104 is installed in an assembled manner, particularly by welding to the walls 121, 122 of the double walls 102. Such a bridge is known, for example, from patent application FR3072448A1 in the prior art.
[0003] During operation, during the combustion reaction R, the temperature of the inner wall 121 is higher than that of the outer wall 122. Due to thermal expansion, this causes a relative displacement between the inner wall 121 and the outer wall 122. Bridge 104 may therefore break, as shown in Figure 2, resulting in a change in the distance between the inner wall 121 and the outer wall 122. Furthermore, the cooling airflow F is likely to be diverted in the broken area of bridge 104. As a result, the cooling of the inner wall 121 is not optimal. The inner wall 121 may include a high-temperature zone Z, which affects its lifespan.
[0004] A direct solution to eliminate this drawback is to strengthen the connections of each bridge, but this requires significant time and manufacturing costs. Therefore, the present invention aims to eliminate at least some of these drawbacks.
[0005] In the prior art derived from patent application US20130047618A1, it is known that the double walls for a gas turbine combustion chamber include protruding protrusions. Summary of the Invention
[0006] This invention relates to a double-walled combustion chamber for an aircraft gas turbine, comprising an inner wall configured to contact the combustion reaction and an outer wall spaced apart from the inner wall. The outer wall includes multiple ports to allow the flow of cooling airflow outside the outer wall, which cools the inner wall. The inner wall is not perforated to prevent airflow from flowing towards the center of the combustion chamber.
[0007] The significant feature of this invention is that the inner wall includes a plurality of members protruding outward from the outer wall, each protruding member extending into a port, thereby defining an adjustable flow cross section between the protruding member and the port for the passage of cooling airflow.
[0008] Advantageously, the multiple protruding members allow for an increased exchange surface between the cooling airflow and the inner wall, thereby improving the service life of the combustion chamber. Furthermore, the positioning of the protruding members in the port allows for the definition of an adjustable flow cross-section, thus allowing for precise regulation of the cooling airflow. Finally, these protruding members do not exhibit a significant thermal gradient during use, which further increases service life. Finally, these protruding members serve to support the inner wall during additive manufacturing.
[0009] Preferably, since each port has a peripheral edge, each protruding member extends away from the peripheral edge of that port. Therefore, there is no direct heat conduction between the protruding member and the outer wall. Furthermore, this allows for differential expansion during operation, taking into account differences in wall temperature.
[0010] Preferably, each protruding member is positioned away from the outer wall, i.e., not in contact, to avoid any heat conduction. The protruding members are advantageously free relative to the outer wall.
[0011] Advantageously, the cooling airflow circulates around the periphery of each protruding member, thereby improving cooling. Preferably, the regulated flow cross-section is circumferential, and more preferably annular.
[0012] According to one aspect of the invention, each protruding member has a flared cross-section facing the inner wall. Therefore, the protruding member has a robust base, thereby extending its service life.
[0013] Preferably, the outer wall includes an outer surface, and each protruding member has an end face that extends as an extension of the outer surface of the outer wall. This is advantageous because it improves the flow of cooling airflow by avoiding the formation of overflow patterns that could lead to turbulence. This characteristic is advantageous in additive manufacturing processes, as will be discussed later.
[0014] According to a preferred aspect of the invention, the double walls are additively manufactured. This manufacturing method ensures precise positioning of the protruding member within the port.
[0015] The present invention also relates to a method for manufacturing double walls as described above, wherein the inner wall and the outer wall are additively manufactured.
[0016] Preferably, the inner and outer walls are fixed to the temporary support by incrementally adding metal powder, and then separated from the temporary support by cutting at the interface between the walls and the temporary support. Preferably, the assembly is de-powdered before separation and then heat-treated.
[0017] The present invention also relates to an aircraft gas turbine combustion chamber comprising the double walls as described above, wherein the inner wall is configured to contact the combustion reaction.
[0018] The present invention also relates to gas turbines, particularly gas turbines for aircraft, including the combustion chamber as described above.
[0019] The present invention also relates to a method of using the aforementioned combustion chamber, comprising:
[0020] The combustion process in the combustion chamber increases the temperature of the inner walls, and
[0021] • The step of allowing cooling airflow from the outside to pass through each adjusted flow cross section of the outer wall, the cross section being defined between the protruding member and the port into which the protruding member extends, in order to cool the inner wall.
[0022] Preferably, when each protruding member thermally expands, each protruding member in the expanded state extends away from the peripheral edge of the port into which it extends. Attached Figure Description
[0023] The invention will be better understood by reading the following description (given as an example) and by referring to the following figures (given as a non-limiting example), wherein similar objects are given the same reference numerals.
[0024] Figure 1 is a schematic diagram of a double-walled combustion chamber with a bridge according to the prior art.
[0025] Figure 2 is a schematic diagram of the double wall in Figure 1 when the bridge connection is disconnected.
[0026] Figure 3 is a schematic diagram of the combustion chamber of a helicopter gas turbine.
[0027] Figure 4 is a schematic cross-sectional view of the double walls of the combustion chamber.
[0028] Figure 5 is a schematic diagram of the outer side of the double walls in Figure 4.
[0029] Figure 6 is a schematic cross-sectional view of the protruding member positioned in the port of the outer wall of the double wall.
[0030] Figure 7 is a schematic cross-sectional view of the additively manufactured double-walled structure, and
[0031] Figure 8 is a schematic diagram of the cross-section of the cooling airflow through the double-wall circulation when the combustion chamber is in use.
[0032] It should be noted that the accompanying drawings illustrate the invention in detail for the purpose of implementing the invention, and the drawings can, of course, be used to better define the invention if necessary. Detailed Implementation
[0033] This invention will be described in relation to the combustion chamber of an aircraft gas turbine. Referring to Figure 3, a combustion chamber 1 of a helicopter gas turbine is shown. It goes without saying that this invention is also applicable to other types of aircraft gas turbines.
[0034] As far as the combustion chamber is concerned, it advantageously means any shell in which the combustion reaction takes place and whose temperature should be controlled.
[0035] As schematically shown in Figure 3, the combustion chamber 1 includes double walls 2, which include an inner wall 21 and an outer wall 22 configured to contact the combustion reaction R. The outer wall 22 is spaced apart from the inner wall 21 to form a heat shield. Preferably, the walls 21 and 22 are metallic. The double walls 2 are shown in more detail in Figures 4 to 6.
[0036] As shown in Figures 4 and 5, the outer wall 22 includes a plurality of ports 3 to allow circulation of a cooling airflow F, which cools the inner wall 21 by flowing through the space between the two walls 21 and 22. The ports 3 are distributed on the outer wall 22 to allow for uniform cooling. Referring to Figure 5, the ports 3 are arranged in rows and columns. Preferably, the ports 3 have a circular cross-section with a radius of r3 (Figure 6), but it goes without saying that they may have different cross-sections. Each port 3 includes a peripheral edge 30, which in this example is circular. According to one aspect of the invention, the number of ports 3 is higher in the hottest region facing the inner wall 21.
[0037] The inner wall 21 is airtight, meaning it has no perforations, to prevent any flow of cooling airflow F towards the center of the combustion chamber 1 (which would affect combustion performance). Such an inner wall 21 can improve the combustion efficiency of the combustion chamber.
[0038] According to the invention, the inner wall 21 includes a plurality of protruding members 4 facing the outer wall 22, each protruding member 4 extending into a port 3, thereby defining a regulating flow cross-section through which the cooling airflow F passes. In this example, each port 3 is associated with a protruding member 4. It goes without saying that some ports 3 may not have protruding members 4.
[0039] Advantageously, the protruding member 4 allows for an increase in the heat exchange surface area between the inner wall 21 and the cooling airflow F, thereby improving the cooling of the inner wall 21. Furthermore, the adjusted flow cross-section allows for precise control of the cooling airflow F for economical use.
[0040] Referring to Figure 6, a schematic cross-sectional view of the protruding member 4 installed in port 3 is shown.
[0041] In this example, the protruding member 4 has a flared cross-section facing the inner wall 21. The flared cross-section allows the protruding member 4 to have a wide base, which ensures a firm connection with the inner wall 21.
[0042] Referring then to FIG. 6, for each protruding member 4, a foot 4a is defined near the inner wall 21, and a head 4b forming the free end of the protruding member 4, the head 4b extending into the port 3. The foot 4a has a larger cross-section than the head 4b. In this example, the foot 4a has a truncated conical cross-section providing high robustness. The head 4b is cylindrical and preferably has a circular cross-section with a radius r4. Preferably, the foot 4a has a cross-section that is at least 50%, preferably at least 100%, larger than the radius r4.
[0043] Referring to Figure 6, each protruding member 4 extends away from the peripheral edge 30 of the port 3 into which it extends. In other words, there is no contact between the protruding member 4 belonging to the inner wall 21 and the peripheral edge 30 of the port 3 belonging to the outer wall 22 that could cause heat conduction. There is no heat transfer between the inner wall 21 and the outer wall 22 through the protruding members.
[0044] Advantageously, the protruding member 4 is located at the center of the port 3, such that the adjustable cross-section is adapted between the head 4b and the port 3, preferably having an annular shape. As the combustion chamber temperature increases, thermal expansion will increase the cross-sectional area of the cooling airflow to achieve optimal cooling. The adjusted cross-section allows for adjustment of the cooling airflow rate, enabling careful use of the cooling airflow.
[0045] Preferably, the radius r3 of port 3 is larger than the radius r4 of protruding member 4 to define a sufficient flow cross-section for cooling air F. The radius r3 of port 3 is at least 10%, more preferably at least 30%, and more preferably at least 100% larger than the radius r4. The space between the protruding member 4 and the peripheral edge 30 of port 3 defines a gap that allows the protruding member 4 to expand. As will be explained later, in the expanded state, each protruding member 4 extends away from the peripheral edge 30 of the port 3 into which it extends. Therefore, any heat conduction between the protruding member 4 and the outer wall 22 is avoided. Preferably, as shown in FIG. 6, each protruding member 4 appears as a body of revolution about axis X, which is locally orthogonal to walls 21, 22.
[0046] Referring again to Figure 6, the head 4b has a planar end face 40. In this example, the planar end face 40 extends uniformly with the outer surface of the outer wall 22, as shown in Figure 6. Therefore, the protruding member 4 does not extend outward relative to the outer wall 22, which avoids the formation of vortices and improves the flow of cooling air.
[0047] Preferably, referring to FIG. 7, the double wall 2 is additively manufactured to achieve optimal alignment between the protruding member 4 and the port 3. As shown in FIG. 7, the double wall 2 is formed on the temporary support 5, and then formed by gradual and continuous deposition along the vertical direction FA. Thus, the outer wall 22 and the head 4b of the protruding member 4 are formed before the foot 4a of the protruding member 4 and the inner wall 21. The protruding member 4 advantageously fulfills its function of supporting the inner wall 21 during the additive manufacturing process, thereby achieving optimal alignment between the protruding member 4 and the port 3.
[0048] According to an exemplary implementation, during the manufacturing process, walls 21 and 22 are fixed to the temporary support 5 by incrementally adding metal powder. The assembly is then de-powdered and heat-treated. Walls 21 and 22 are separated from the temporary support 5 by cutting at the interface between walls 21 and 22 and the temporary support 5. This additive manufacturing advantageously makes it possible to obtain original and innovative geometries while reducing thickness. Furthermore, this additive manufacturing eliminates the need for molds, which is a source of savings. It goes without saying that walls 21 and 22 can also be manufactured by combining parts obtained through mechanical welding or casting.
[0049] The inner wall 21 and outer wall 22 are then installed in the combustion chamber 1 to provide a space E between them, as shown in Figure 8. Due to precision manufacturing, each protruding member 4 is optimally fitted with the port 3 to provide an regulated flow cross-section between the protruding member 4 and the port 3 for the passage of cooling airflow F. Installation with digons, etc., can be implemented.
[0050] An exemplary example of the invention will now be described with reference to FIG8. During turbomachinery operation, the method includes:
[0051] • The combustion step R in combustion chamber 1 increases the temperature of the inner wall 21, and
[0052] • The step of allowing cooling airflow F from the outside to circulate through each regulating flow cross section of the outer wall, the cross section being defined between the protruding member 4 and its extending port 3, in order to cool the inner wall 21.
[0053] Advantageously, the cooling airflow F moves within the space E formed between the inner wall 21 and the outer wall 22 through an adjusted flow cross-section. The cooling airflow F allows contact with the entire surface area of the protruding member 4, thereby maximizing heat exchange.
[0054] Preferably, during operation, each protruding member 4 thermally expands. In the expanded state, each protruding member 4 extends away from the peripheral edge 30 of the port 3 into which it extends. Therefore, any heat conduction between the protruding member 4 and the outer wall 22 is avoided.
[0055] With this invention, the double walls 2 can be optimally cooled by the cooling airflow F without the risk of creating weak points or breaks. The presence of the protruding member 4 increases the heat exchange surface area and regulates the flow cross-section of the cooling airflow F.
Claims
1. A double wall (2) for a combustion chamber (1) of an aircraft gas turbine, the double wall (2) comprising an inner wall (21) configured to contact a combustion reaction and an outer wall (22) spaced apart from the inner wall (21), the outer wall (22) comprising a plurality of ports (3) to allow the flow of a cooling airflow (F) outside the outer wall (22) to cool the inner wall (21), the inner wall (21) being non-perforated to prevent any flow of the cooling airflow (F) toward the center of the combustion chamber (1), wherein the double wall (2) is characterized in that the inner wall (21) 1) Includes a plurality of protruding members (4) facing an outer wall (22), each protruding member (4) including a foot (4a) and a cylindrical head (4b) of a circular cross section extending into a port (3) of a circular cross section to define an regulated flow cross section for passage of cooling airflow (F) between the protruding member (4) and the port (3), wherein the outer wall (22) includes an outer surface, and each protruding member (4) includes an end face (40) extending as an extension of the outer surface of the outer wall (22).
2. The double-walled (2) according to claim 1, wherein, Each port (3) has a peripheral edge (30), and each protruding member (4) extends away from the peripheral edge (30) of the port (3).
3. The double-walled (2) according to claim 1, wherein, Each protruding component (4) is away from the outer wall (22).
4. The double-walled (2) according to claim 1, wherein, The adjusted flow cross-section is circumferential.
5. The double-walled (2) according to claim 1, wherein, The foot (4a) has a trumpet-shaped cross section facing the inner wall (21).
6. A method for manufacturing the double-walled (2) according to claim 1, wherein, The inner wall (21) and the outer wall (22) are additively manufactured.
7. The method for manufacturing double-walled (2) according to claim 6, wherein, The inner wall (21) and outer wall (22) are fixed to the temporary support (5) by incrementally adding metal powder, and then separated from the temporary support (5) by cutting at the interface between the wall (21, 22) and the temporary support (5).
8. A combustion chamber (1) for an aircraft gas turbine comprising a double-walled (2) according to any one of claims 1 to 5, wherein, The inner wall (21) is configured to be in contact with the combustion reaction.
9. An aircraft gas turbine comprising a combustion chamber (1) according to claim 8.
10. A method of using the combustion chamber (1) according to claim 8, comprising: The combustion process in the combustion chamber (1) increases the temperature of the inner wall (21), and The step of allowing cooling airflow (F) from the outside to circulate through each adjustable flow cross section of the outer wall (22), the adjustable flow cross section being defined between the protruding member (4) and the port (3) into which the protruding member (4) extends, in order to cool the inner wall (21).