An integrated flame tube wall for an aircraft engine

Abstract

The present application belongs to the technical field of design of the wall of aero-engine flame duct, and specifically relates to an integrally formed flame duct wall of an aero-engine, which is formed by stacking layers in a vertical direction upwards using a metal additive manufacturing technology and a laser selective melting process; a plurality of impact cooling holes are formed on the outer wall of the duct, and the radial size of each impact cooling hole gradually shrinks and is conical, with a side wall angle of 45° to 75°; a plurality of air film holes are formed on the inner wall of the duct, and each air film hole is tilted backward with an inclination angle of 45° to 75°; a plurality of groups of suction holes are formed on the inner wall of the duct, and each group of suction holes surrounds each air film hole, and the inlet end extends to the inner wall of the inlet of each air film hole; an annular impact cavity is formed in the duct, and the annular impact cavity connects each impact cooling hole and air film hole, and a plurality of spoiler columns are formed in the annular impact cavity, and each spoiler column has chamfered ends with a chamfer angle of 45° to 75°.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft engine flame duct wall design, and specifically relates to an aircraft engine integrally formed flame duct wall. Background Art

[0002] In pursuit of higher thrust, the temperature in front of the aircraft engine turbine needs to be increased, which causes the temperature inside the flame tube in the combustion chamber to rise sharply. The flame tube wall is subjected to a high temperature load and is prone to ablation. For this reason, the flame tube wall is designed to adopt porous layer plate cooling technology to efficiently cool the flame tube wall.

[0003] At present, the flame tube wall adopts the porous layer plate cooling technology, which includes an annular outer wall and an annular inner wall arranged inside the outer wall. The outer wall has multiple impingement cooling holes on the side wall and multiple spoiler columns formed on the inner side; the inner wall has multiple air film holes on the side wall; each spoiler column is connected to the outer side of the inner wall by diffusion welding or brazing, such as Figure 1 As shown, the cooling air can enter between the outer wall and the inner wall through each impact cooling hole for impact cooling, and is discharged through each air film hole after being disturbed by each spoiler column, forming an air film on the inner side of the inner wall for air film cooling, thereby being able to efficiently cool the flame tube wall and protect the flame tube wall from ablation.

[0004] Due to the limitations of processing technology, the flame tube wall currently using porous layer plate cooling technology mainly adopts a circumferential segmentation + welding structure, or a floating tile split structure.

[0005] The flame tube wall adopts the structure of circumferential segmentation + welding, and the outer wall and the inner wall are divided into multiple corresponding fan-shaped segments. The edges of each fan-shaped segment are welded together by electron beam welding. Figure 2 As shown, the circumferential joint surface is a solid structure, and spoiler columns cannot be arranged. The local cooling effect is poor, and ablation is prone to occur in actual applications.

[0006] The flame tube wall adopts the floating tile split structure. The outer wall is designed to maintain the whole ring structure, and the inner wall is designed to be a plurality of corresponding fan-shaped segments. Each fan-shaped segment is connected to the inner side of the outer wall by bolt fasteners. There is a floating gap between each fan-shaped segment. Figure 3 As shown, many parts are required, assembly is difficult, and the overall mass of the structure is large, which is inconsistent with the requirements of aircraft engine weight reduction. In addition, the local cooling effect at each floating gap is poor, and ablation is prone to occur in actual applications.

[0007] This application is proposed in view of the above-mentioned technical defects.

[0008] It should be noted that the disclosure of the above background technology content is only used to assist in understanding the inventive concept and technical solution of the present invention, and it does not necessarily belong to the prior art of this patent application. In the absence of clear evidence that the above content has been disclosed on the filing date of this application, the above background technology should not be used to evaluate the novelty and creativity of this application. Summary of the Invention

[0009] The purpose of the present application is to provide an integrally formed flame tube wall of an aircraft engine to overcome or alleviate at least one of the existing technical defects.

[0010] The technical solution of this application is:

[0011] An integrally formed flame tube wall of an aircraft engine is formed by stacking layers vertically upwards using a metal additive manufacturing technology and a laser selective melting process;

[0012] There are multiple impingement cooling holes formed on the outer wall of the cylinder. The radial size of each impingement cooling hole gradually shrinks and becomes conical, with the side wall angle of 45° to 75°.

[0013] A plurality of air film holes are formed on the inner wall of the cylinder wall, and each air film hole is tilted backward at an angle of 45° to 75°; a plurality of groups of suction holes are formed on the inner wall of the cylinder wall, and each group of suction holes surrounds each air film hole, and the inlet end extends to the inner wall of the inlet of each air film hole;

[0014] An annular impact cavity is formed in the cylinder wall, which is connected to each impact cooling hole and air film hole. Multiple spoiler columns are formed in the annular impact cavity, and both ends of each spoiler column are chamfered with a chamfer angle of 45° to 75°.

[0015] According to at least one embodiment of the present application, in the above-mentioned integrally formed flame tube wall of the aircraft engine, the inlet end of each group of suction holes is located within the reflow zone at the inlet of each air film hole.

[0016] According to at least one embodiment of the present application, in the above-mentioned integrally formed flame tube wall of the aircraft engine, the number of each group of suction holes is no less than 6.

[0017] According to at least one embodiment of the present application, in the above-mentioned integrally formed flame tube wall of the aircraft engine, the various impingement cooling holes, spoiler columns, and film holes are periodically distributed along the circumferential and axial directions.

[0018] According to at least one embodiment of the present application, in the above-mentioned integrally formed flame tube wall of the aircraft engine, multiple groups of spoiler ribs are formed on the inner side wall of the annular impact cavity in the tube wall, and each group of spoiler ribs is located between adjacent spoiler columns in the oblique direction, wherein each spoiler rib gradually changes from an outer arc shape to a straight line from the outside to the inside.

[0019] According to at least one embodiment of the present application, in the above-mentioned integrally formed flame tube wall of the aircraft engine, the cross-section of each group of spoiler ribs on the inner side wall of the annular impact cavity in the tube wall is trapezoidal.

[0020] According to at least one embodiment of the present application, in the above-mentioned integrally formed flame tube wall of the aircraft engine, the height H of each group of spoiler ribs on the inner side wall of the annular impact cavity in the tube wall does not exceed one-fourth of the thickness L of the annular impact cavity, the bottom width W is 1 to 2 times the height H, the inclination angles F on both sides are 15° to 45°, and the spacing P between them is 1 to 3 times the bottom width W. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 This is a schematic diagram of the current porous plate cooling of the flame tube wall of an aircraft engine;

[0022] Figure 2 This is a schematic diagram of the current porous plate cooling aircraft engine flame tube wall adopting the circumferential segmentation + welding structure;

[0023] Figure 3 This is a schematic diagram of the floating tile split structure used in the current porous plate cooling aircraft engine flame tube wall;

[0024] Figure 4 This is a schematic diagram of the integrally formed flame tube wall of an aircraft engine provided by an embodiment of the present application;

[0025] Figure 5 This is a schematic diagram of the operation of the suction holes in the wall of the integrally formed flame tube of an aircraft engine provided by an embodiment of the present application;

[0026] Figure 6 Schematic diagram of the position distribution of various impingement cooling holes, spoiler columns, and film holes, and the flow direction of airflow in the annular impingement cavity when there are no spoiler ribs in the wall of the integrally formed flame tube of an aircraft engine provided by an embodiment of the present application;

[0027] Figure 7 This is a partial schematic diagram of the provision of spoiler ribs in the wall of an integrally formed flame liner of an aircraft engine provided by an embodiment of the present application;

[0028] Figure 8 This is a schematic diagram of the form of spoiler ribs provided in the wall of the integrally formed flame duct of an aircraft engine and their relative position distribution according to an embodiment of the present application;

[0029] in:

[0030] 1-impact cooling hole; 2-air film hole; 3-suction hole; 4-spoiler column; 5-spoiler rib.

[0031] In order to better illustrate this embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of the actual product. In addition, the drawings are only used for illustrative purposes and should not be understood as limitations on this patent. DETAILED DESCRIPTION

[0032] To make the technical solution and its advantages of the present application clearer, the technical solution of the present application will be described in further detail below in conjunction with the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of the present application and are only used to explain the present application, not to limit the present application. It should be noted that, for ease of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments of the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0033] In addition, unless otherwise defined, the technical or scientific terms used in the description of this application should have the ordinary meanings understood by those of ordinary skill in the art to which this application belongs. The words "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer" used in the description of this application are only used to indicate relative directions or positional relationships, and do not imply that the device or component must have a specific orientation, be constructed, or operate in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. Therefore, they should not be understood as limitations on this application. The words "first," "second," "third," and similar terms used in the description of this application are used only for descriptive purposes to distinguish different components and should not be understood to indicate or imply relative importance. The words "one," "an," or "the" used in the description of this application should not be understood as absolute limitations on quantity, but should be understood as meaning the presence of at least one. The words "include" or "comprises" used in the description of this application mean that the element or object listed before the word includes the elements or objects listed after the word and their equivalents, but does not exclude other elements or objects.

[0034] In addition, it should be noted that, unless otherwise clearly stipulated and limited, the words "install", "connect", "connect" and similar terms used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, an indirect connection through an intermediate medium, or a connection between two components. Technical personnel in the field can understand their specific meanings in this application according to the specific circumstances.

[0035] In order to overcome the problems of poor local cooling effect, easy ablation, difficult assembly, and large mass in the porous layer plate cooling of the flame tube wall of the aircraft engine using the circumferential segmentation + welding structure and the floating tile split structure, the embodiment of the present application provides an aircraft engine integrally formed flame tube wall, such as Figure 4 As shown, the metal additive manufacturing technology and the laser selective melting process are used for integral molding, wherein a plurality of impact cooling holes 1 are formed on the outer wall of the cylinder wall, a plurality of air film holes 2 are formed on the inner wall of the cylinder wall, and an annular impact cavity is formed in the cylinder wall, which connects the various impact cooling holes 1 and air film holes 2, and a plurality of spoiler columns 4 are formed in the annular impact cavity. The cooling gas can enter the annular impact cavity through the various impact cooling holes 1 for impact cooling, and is discharged through the various air film holes 2 after being disturbed by the various spoiler columns 4, forming an air film on the inner wall of the cylinder wall for air film cooling, thereby efficiently cooling the wall of the flame tube and protecting the wall of the flame tube from ablation.

[0036] The flame tube is a rotating body around the engine axis. In order to ensure the roundness of the flame tube and facilitate the control of thermal deformation during processing, the metal additive manufacturing technology laser selective melting process is used to integrally form the flame tube wall, and the parts are stacked layer by layer in the vertical direction.

[0037] During the one-piece forming process of the laser selective melting process of metal additive manufacturing technology, the already formed layers of the structure need to provide support for the "layer" to be formed, otherwise the structure will be affected by gravity and cause "collapse", bending deformation and surface quality problems. There are certain requirements for the angle between the "cantilever" type structure and the main structure. In the integrally formed flame tube wall of the aircraft engine disclosed in the above embodiment, the impact cooling holes 1, the air film holes 2, and the spoiler columns 4 have the characteristics of a "cantilever" type structure, and the angle between the edge of the structure and the horizontal direction is required to be greater than 45°, and should not exceed 75°. In order to ensure the angle requirements, each impingement cooling hole 1 is designed to be a cone with a large inlet and a small outlet, and the air film hole 2 is designed to be a backward beveled hole with an angle between the center line and the wall surface, and the two ends of each spoiler column 4 are chamfered. The chamfering of the two ends of each spoiler column 4 can also reduce the stress concentration of the structure and improve the reliability of the structure.

[0038] In the wall of the integrally formed flame tube of the aircraft engine disclosed in the above-mentioned embodiment, if each air film hole 2 adopts a small hole of equal cross-section, separation will easily occur at the inlet and outside the outlet. If a trumpet-shaped small hole with a small inlet and a large outlet is adopted, the outlet velocity of the airflow will be reduced. Although separation outside the outlet can be suppressed, the separation zone will move to the outlet, which is not ideal for suppressing separation. For this reason, in some optional embodiments, the SLM process can be used to design the hole shape of each air film hole 2 into a form that meets the optimal flow state of the airflow to suppress the separation of the airflow in the hole and at the outlet, and multiple groups of suction holes 3 are designed to be formed on the inner wall of the tube wall. Each group of suction holes 3 surrounds each air film hole 2, and the inlet end extends to the inner wall of the inlet of each air film hole 2. Within the reflux area at the inlet, the airflow within the reflux area is sucked out, such as Figure 5 As shown, this can effectively suppress the separation at the inlet of the air film hole 2, and the sucked and discharged air flow enters the inner side of the inner wall of the cylinder wall, which can ensure the cooling effect of the structure and make the wall temperature on the inner wall of the cylinder wall more uniform, which helps to reduce the temperature gradient and improve the thermal fatigue life. In addition, the profiles at the inlet of each air film hole 2 and the inlet of the suction hole 3 can be further optimized to improve the diversion effect and reduce the pressure loss. In actual application, the number of each group of suction holes 3 can be designed to be no less than 6, and the suction amount of the air flow is less than 10% of the air flow of the corresponding air film hole 2.

[0039] In some optional embodiments, in the integrally formed flame tube wall of the aircraft engine disclosed in the above embodiment, each impingement cooling hole 1, spoiler column 4, and film hole 2 is periodically distributed along the circumferential direction and the axial direction, such as Figure 6 , to ensure the cooling effect.

[0040] Furthermore, in the above embodiment disclosed in the integrally formed flame tube wall of the aircraft engine, a plurality of groups of spoiler ribs 5 are formed on the inner side wall of the annular impact cavity in the tube wall, such as Figure 7 As shown, in order to increase the disturbance of the airflow in the annular impact chamber and ensure the cooling effect, further, each group of spoiler ribs 5 can be designed to be located between each adjacent spoiler column 4 in the oblique direction, and each spoiler rib 5 gradually changes from an outer arc shape to a straight line shape from the outside to the inside, as shown in FIG. Figure 8 As shown, it is roughly perpendicular to the flow direction of the airflow in the annular impact cavity, thereby achieving the maximum turbulence effect and thus ensuring the cooling effect.

[0041] In order to adapt to the laser selective melting process of metal additive manufacturing technology, the cross-section of each group of spoiler ribs 5 in the wall of the integrally formed flame tube of the aircraft engine disclosed in the above embodiment is further designed to be trapezoidal, with a height H not exceeding one-quarter of the thickness L of the annular impact cavity, a bottom width W being 1 to 2 times the height H, an inclination angle F on both sides being 15° to 45°, and a spacing P therebetween being 1 to 3 times the bottom width W.

[0042] The various embodiments in the specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to in detail.

[0043] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or replacements to the relevant technical features, and the technical solutions after these changes or replacements will fall within the scope of protection of the present application.

Claims

1. An integrally formed flame tube wall of an aircraft engine, characterized in that: The metal additive manufacturing technology laser selective melting process is used to stack layers vertically upward to form an integrated structure. A plurality of impingement cooling holes (1) are formed on the outer wall of the cylinder wall, and the radial size of each impingement cooling hole (1) gradually shrinks and is tapered, with a side wall angle of 45° to 75°; A plurality of air film holes (2) are formed on the inner side wall of the cylinder wall, and each air film hole is tilted backward with an inclination angle of 45° to 75°; a plurality of groups of suction holes (3) are formed on the inner side wall of the cylinder wall, and each group of suction holes (3) surrounds each air film hole (2), and the inlet end extends to the inner side wall of the inlet of each air film hole (2); An annular impact cavity is formed in the cylinder wall, the annular impact cavity is connected to each impact cooling hole (1) and the air film hole (2), and a plurality of spoiler columns (4) are formed in the annular impact cavity. Both ends of each spoiler column (4) are chamfered, and the chamfer angle is 45° to 75°.

2. The integrally formed flame liner wall of an aircraft engine according to claim 1, characterized in that: The inlet end of each group of suction holes (3) is located within the reflow zone at the inlet of each air film hole (2).

3. The integrally formed flame liner wall of an aircraft engine according to claim 1, characterized in that: The number of each group of suction holes (3) is not less than 6.

4. The integrally formed flame liner wall of an aircraft engine according to claim 1, characterized in that: The impact cooling holes (1), the spoiler columns (4), and the air film holes (2) are periodically distributed along the circumferential direction and the axial direction.

5. The integrally formed flame liner wall of an aircraft engine according to claim 1, characterized in that: A plurality of groups of spoiler ribs (5) are formed on the inner side wall of the annular impact cavity in the cylinder wall, and each group of spoiler ribs (5) is located between adjacent spoiler columns (4) in an oblique direction, wherein each spoiler rib (5) gradually changes from an outer arc shape to a straight line shape from the outside to the inside.

6. The integrally formed flame liner wall of an aircraft engine according to claim 1, characterized in that: The cross section of each group of flow-turbulating ribs (5) on the inner side wall of the annular impact cavity in the cylinder wall is trapezoidal.

7. The integrally formed flame liner wall of an aircraft engine according to claim 6, characterized in that: The height H of each group of spoiler ribs (5) on the inner side wall of the annular impact cavity in the cylinder wall does not exceed one quarter of the thickness L of the annular impact cavity, the bottom width W is 1 to 2 times the height H, the inclination angles F on both sides are 15° to 45°, and the spacing P between them is 1 to 3 times the bottom width W.

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