Aero-engine radial diffuser structure and manufacturing method thereof
By designing a split radial diffuser structure, and using the expanded and stable split structure to add small blades at the leading edge of the blade, the problem of expanding the throat area in the prior art will be solved, and the effect of rapidly increasing the throat area and reducing costs is achieved.
Patent Information
- Application Number
- CN202510169991.4
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-02-17
- Publication Date
- 2025-05-13
AI Technical Summary
When the performance of existing radial diffusers is attenuated, the throat area expansion method needs to be disassembled and machined, which is time-consuming and cost-effective, and the blade leading edge disassembly consistency is poor.
A split radial diffuser structure is designed, including a dilated and stable split structure. By designing the compression relationship between the small blades and the roots at the leading edge of the blade, the throat area is quickly increased and the dismantling machine processing is avoided.
It realizes rapid increase in the throat area of the radial diffuser when performance decays, reduces repair costs, and improves blade leading edge consistency.
Smart Images
Figure CN119982651A_ABST
Abstract
Description
Technical field:
[0001] The invention relates to an aeroengine radial diffuser structure and a manufacturing method thereof, and belongs to the field of impeller machinery such as aeroengines and gas turbines. Background technology:
[0002] The radial diffuser of an aircraft engine is commonly found in centrifugal compressors. It is usually located downstream of the centrifugal impeller. It is a section of stator flow channel with a large cross-sectional area. Its main function is to decelerate and increase the pressure of the high-speed gas at the outlet of the centrifugal impeller, that is, the gas velocity passing through the radial diffuser will be greatly reduced, and the static pressure will be further increased, providing favorable conditions for the gas to further participate in combustion. Due to the high inlet velocity and large adverse pressure gradient of the radial diffuser, it is easy to cause flow instability, such as flow separation, shock wave-boundary layer interference, etc., which in turn affects the stable operating range of the entire compressor, and may cause severe machine surge. One of the important factors affecting the stable operating range of the centrifugal compressor is the matching degree between the radial diffuser and the centrifugal impeller. As mentioned above, the radial diffuser itself has a narrow stable operating range, which undoubtedly brings great difficulty to the performance matching of the two. Further research shows that the throat area of the radial diffuser has a huge impact on the matching performance of the two.
[0003] Figure 1 A schematic diagram of a typical radial diffuser of a centrifugal compressor is given. The typical radial diffuser consists of a diffuser body 1 with blades 2 and a cover plate 3. Figure 2 A schematic diagram of the assembly relationship of the components of the radial diffuser is given. The assembly of the components of the radial diffuser is relatively simple, that is, the cover plate 3 is installed on the diffuser body 1, and the connection relationship between the two can be bolted. Figure 3 A schematic diagram of the radial diffuser blade channel is given. The inlet air flow velocity is Vin, the outlet air flow velocity is Vout, the area sandwiched by the two blades 2 is the air flow channel, and an inscribed circle is drawn from the inlet to the outlet along the blade channel. There is a minimum inscribed circle a near the inlet of the air flow channel, that is, the throat section, and the diameter of the minimum inscribed circle is D. Figure 4 A schematic diagram of the key parameters of the radial diffuser meridian section is given. The inlet channel width of the meridian section is H1, and the outlet channel width is H2. Generally speaking, H1 = H2 = H. According to the geometric relationship, the throat area of the radial diffuser (i.e., the minimum throttling area between two blades 2) is A = H*D.
[0004] During the long-term operation of the engine, the centrifugal impeller and radial diffuser will experience a certain degree of performance degradation, which is mainly due to unavoidable factors such as blade surface wear, contamination, and corrosion. When performance degradation occurs, the pressure ratio of the centrifugal impeller will decrease, and the gas density will decrease. At this time, under the premise that the channel cross-sectional area remains unchanged, the gas mass flow rate passing through will decrease compared to the state before attenuation, making the compressor operating point close to the surge boundary, and the risk of engine instability is greatly increased. Generally speaking, without replacing parts, when the above-mentioned performance degradation occurs, one of the common practices is to open the radial diffuser throat to maintain a gas mass flow rate close to the original one, so as to keep the compressor operating point basically unchanged.
[0005] At present, the main method to expand the throat area of radial diffuser is to "cut and grind" the leading edge of the blade, that is, to cut off the leading edge of the radial diffuser blade and polish it to make it smooth by machine tool processing + manual polishing to achieve the effect of increasing the throat area. Figure 5 As shown, by cutting off the leading edge R of the blade L , thereby reducing the diameter of the throat section from D o Increase to D M , thereby increasing the throat area.
[0006] The existing method of expanding the throat area of the radial diffuser requires the radial diffuser to be disassembled and then processed on a machine tool, which is time-consuming and increases costs. At the same time, the manual polishing method is used, and the consistency of the blade leading edge polishing is poor.
[0007] Therefore, it is necessary to improve the existing technology to solve the shortcomings of the existing technology. Summary of the invention:
[0008] The present invention aims to solve the problems existing in the above-mentioned prior art and provides an aircraft engine radial diffuser structure and a manufacturing method thereof. When the radial diffuser has performance degradation, the throat area of the radial diffuser can be effectively expanded to increase the working stability, and the consistency of the blade leading edge is better after the throat area is increased.
[0009] The present invention adopts the following technical scheme: an aircraft engine radial diffuser structure, including a diffuser body, a cover plate and a stabilizing split structure, wherein a plurality of spaced-apart blades are integrally formed on the diffuser body, and the area sandwiched between two adjacent blades is an airflow channel, and the stabilizing split structure includes a root and small blades formed on the root and spaced apart, and when the cover plate is installed on the diffuser body and the root of the stabilizing split structure is pressed against the inlet of the airflow channel, the small blades are pressed together with the leading edge of the blades.
[0010] Furthermore, there is a one-to-one correspondence between the small blades and the blades.
[0011] Furthermore, the thickness B1 of the root portion decreases gradually from the thickness B1 far from the end of the small blade to the thickness B2 close to the end of the small blade.
[0012] Further, the thickness B2 adjacent to the tip of the bladelet is ≮2 mm.
[0013] Furthermore, an inscribed circle is drawn from the inlet to the outlet along the air flow channel. There is a minimum inscribed circle a, i.e., the throat cross section, near the inlet of the air flow channel. The diameter of the minimum inscribed circle is D, and the inlet width between adjacent blades is H. The throat area of the radial diffuser is A=H*D.
[0014] The present invention also adopts the following technical solution: Step 1: Determine the throat area value ΔA that needs to be added when the performance is attenuated, and leave an additional flow margin of 5 to 10% to cope with the engine performance attenuation. The speed of the radial diffuser inlet remains unchanged before and after the performance attenuation, and the newly added throat area value ΔA is:
[0015] ΔA=(1.1P 1t / P 2t (T 2t / T 1t ) 0.5 -1)·H·D0
[0016] Among them, P 1t and P 2t Respectively represent the total pressure at the radial diffuser inlet before and after performance degradation, T 1t and T 2t They represent the total temperature of the radial diffuser inlet before and after performance attenuation, D0 represents the small circle diameter at the initial throat of the diffuser, and H represents the inlet width between adjacent blades;
[0017] Step 2: Add the throat area value ΔA as needed and the small circle diameter D at the initial throat of the diffuser o Or the radius R o , calculate the thickness ΔD that needs to be cut off at the leading edge of the blade and the diameter D of the small circle at the new throat M Or the radius R M :
[0018] ΔD=ΔA / H
[0019] D M =D o +ΔD
[0020] R M =R o +0.5ΔD
[0021] Step 3: After determining the leading edge cutting thickness ΔD, starting from the new leading edge point after cutting, use a straight line transition to determine the blade part that needs to be cut off. H1 is the inlet width between adjacent blades of the radial diffuser, and H 2i is the blade width between adjacent blade cutting boundaries of the radial diffuser, set H1 = H 2i =H;
[0022] Step 4: During the assembly phase, first fix the radial diffuser body, and then install the diffuser split structure and cover plate in sequence;
[0023] Step 5: When performance degradation occurs and stabilization is required, disassemble the radial diffuser, decompose the lower cover plate and the stabilization split structure in turn, remove the stabilization split structure and discard it, install the stabilization split structure without the small blades and only the root, and reassemble it according to the requirements of step 3. At this time, the throat area of the leading edge of the radial diffuser blade has been increased.
[0024] Furthermore, in step 2, it is assumed that the radial cutting length of the blade meridian plane is L1 and the meridian chord length of the blade is L T , L1≯0.1L T .
[0025] The present invention has the following beneficial effects: the present invention designs the radial diffuser inlet part to be split, and when performance degradation occurs and stabilization is required, the stabilization split structure is removed, thereby quickly achieving the effect of increasing the throat area of the radial diffuser, effectively reducing the repair cost, and the consistency of the leading edge of the blade after the throat area is increased is good. Description of the drawings:
[0026] Figure 1 Schematic diagram of a typical centrifugal compressor radial diffuser.
[0027] Figure 2 This is a schematic diagram of the assembly relationship of the components of the radial diffuser.
[0028] Figure 3 Schematic diagram of the radial diffuser blade channel.
[0029] Figure 4 Schematic diagram of key parameters of radial diffuser meridian section.
[0030] Figure 5 Schematic diagram of the radial diffuser throat area before and after expansion.
[0031] Figure 6 It is a schematic diagram of the structure of the radial diffuser of the present invention.
[0032] Figure 7 for Figure 6 A is a partial enlarged schematic diagram of the middle part.
[0033] Figure 8 It is a meridian schematic diagram of the radial diffuser assembly with a stabilizing split structure of the present invention.
[0034] Fig. 9 This is a front view of the assembly relationship of the radial diffuser with a stabilizing split structure of the present invention.
[0035] Fig.10 It is a schematic diagram of the disassembly and assembly process when repairing the radial diffuser structure of the present invention. Specific implementation method:
[0036] The present invention will be further described below in conjunction with the accompanying drawings.
[0037] The radial diffuser structure of an aircraft engine of the present invention comprises a diffuser body 1', a cover plate 3' and a diffuser and stabilizer split structure 4', wherein a plurality of spaced blades 2' are integrally formed on the diffuser body 1', and the area sandwiched between two adjacent blades 2' is an airflow channel. The diffuser and stabilizer split structure 4' comprises a root 41' and small blades 40' formed on the root 41' and spaced apart.
[0038] When the root 41 ' of the expansion and stabilization split structure 4 ' is installed at the inlet of the airflow channel, the small blade 40 ' and the leading edge of the blade 2 ' are pressed tightly together.
[0039] The design flow chart of the radial diffuser structure of an aircraft engine of the present invention is as follows:
[0040] Step 1: Determine the additional throat area value ΔA required when performance is attenuated. During the design stage of the radial diffuser of the compressor, an additional 5-10% flow margin should be considered to cope with the engine performance attenuation. For centrifugal compressors, this is mainly due to the decrease in the pressure ratio of the centrifugal impeller upstream of the radial diffuser, which leads to a decrease in the mass flow rate at the inlet of the radial diffuser, resulting in a significant attenuation of the power or thrust of the entire engine. At this time, it is necessary to increase the throat area of the radial diffuser to compensate for the performance attenuation caused by the reduction in flow. Taking a 10% flow margin as an example, generally speaking, the speed of the radial diffuser inlet remains basically unchanged before and after performance attenuation, and the additional diffuser throat area value ΔA is:
[0041] ΔA=(1.1P 1t / P 2t (T 2t / T 1t ) 0.5 -1)·H·D0
[0042] Among them, P 1t and P 2t Respectively represent the total pressure at the radial diffuser inlet before and after performance degradation, T 1t and T 2tThey represent the total temperature of the radial diffuser inlet before and after performance attenuation, D0 represents the small circle diameter at the initial throat of the diffuser, and H represents the inlet width between adjacent blades.
[0043] Step 2: If Figure 5 As shown, according to the required additional throat area value ΔA and the small circle diameter D at the initial throat of the diffuser o Or the radius R o , calculate the thickness ΔD that needs to be cut off at the leading edge of the blade and the diameter D of the small circle at the new throat M Or the radius R M :
[0044] ΔD=ΔA / H
[0045] D M =D o +ΔD
[0046] R M =R o +0.5ΔD
[0047] Step 3: After determining the cutting thickness ΔD of the leading edge, it is recommended to use a straight line transition from the new leading edge point after cutting to determine the blade part that needs to be cut off. The cutting parameters in this step are as follows: Figure 7 As shown, H1 is the inlet width between adjacent blades of the radial diffuser, H 2i It is the blade width between adjacent blade cutting boundaries of the radial diffuser. Generally speaking, H1=H 2i In this paper, H1=H 2i =H, L1 is the radial cutting length of the blade meridian plane, B1 and B2 are the thickness of the root 41' of the expansion and stabilization split structure 4', where L1≯0.1L T , B1>B2 and B2≮2mm, L T is the meridian chord length of the blade.
[0048] Step 4: Assembly stage, such as Figure 8 As shown, first fix the radial diffuser body 1', then install the expansion and stabilization split structure 4' and the cover plate 3' in sequence. Fig. 9 The assembly relationship of the radial diffuser with the stabilizing split structure 4' in the front view is given.
[0049] Step 5: When performance degradation occurs and stabilization is required, the radial diffuser is disassembled, and the lower cover plate 3' and the stabilization split structure 4' are disassembled in sequence. The stabilization split structure is removed and discarded, and the stabilization split structure without the small blades and only the root is installed. Fig.10 As shown, and reassemble according to the requirements of step three, at this time, the throat area of the leading edge of the radial diffuser blade has been increased.
[0050] It should be noted that, according to specific needs and different throat area increase requirements, the stamping die can be designed with different expansion and stabilization split structures, and the principle is exactly the same as above.
[0051] The aero-engine radial diffuser structure of the present invention designs the radial diffuser inlet portion to be split, and adds a stabilizing split structure. A portion of the radial diffuser leading edge (the portion that needs to be cut off when stabilizing is required) grows on the stabilizing split structure. When performance degradation occurs and stabilizing is required, the stabilizing split structure is removed, thereby quickly achieving the effect of increasing the throat area of the radial diffuser.
[0052] The above description is only a preferred embodiment of the present invention. It should be pointed out that a person skilled in the art can make several improvements without departing from the principle of the present invention, and these improvements should also be regarded as within the protection scope of the present invention.
Claims
1. An aero-engine radial diffuser structure, characterized in that: The invention comprises a diffuser body (1'), a cover plate (3') and a stabilizing split structure (4'); the diffuser body (1') is integrally formed with a plurality of spaced-apart blades (2'); the region sandwiched between two adjacent blades (2') is an airflow channel; the stabilizing split structure (4') comprises a root (41') and small blades (40') spaced-apart and formed on the root (41'); when the cover plate (3') is mounted on the diffuser body (1'), the root (41') of the stabilizing split structure (4') is pressed against the inlet of the airflow channel, and the small blades (40') are pressed together with the leading edge of the blades (2').
2. The aerospace engine radial diffuser structure according to claim 1, characterized in that: The small blades (40') and the blades (2') are in a one-to-one correspondence.
3. The aerospace engine radial diffuser structure according to claim 2, characterized in that: The thickness B1 of the root portion (41') gradually decreases from the thickness B1 far from the end of the small blade (40') to the thickness B2 close to the end of the small blade (40').
4. The aerospace engine radial diffuser structure according to claim 3, characterized in that: The thickness B2 adjacent to the tip of the small blade (40') is 2 mm.
5. The aeroengine radial diffuser structure according to claim 4, characterized in that: Draw an inscribed circle from the inlet to the outlet along the air flow channel. There is a minimum inscribed circle a near the inlet of the air flow channel, i.e. the throat cross section. The diameter of the minimum inscribed circle is D. The inlet width between adjacent blades is H. The throat area of the radial diffuser is A=H*D.
6. A method for manufacturing an aircraft engine radial diffuser structure, characterized in that: Here are the steps: Step 1: Determine the additional throat area value ΔA required when performance is degraded, and leave an additional 5-10% flow margin to cope with engine performance degradation. The speed at the radial diffuser inlet remains unchanged before and after performance degradation, so the additional throat area value ΔA is: ΔA=(1.1P 1t / P 2t (T 2t / T 1t ) 0.5 -1)·H·D0 Among them, P 1t and P 2t Respectively represent the total pressure at the radial diffuser inlet before and after performance degradation, T 1t and T 2t They represent the total temperature of the radial diffuser inlet before and after performance attenuation, D0 represents the small circle diameter at the initial throat of the diffuser, and H represents the inlet width between adjacent blades; Step 2: Add the throat area value ΔA as needed and the small circle diameter D at the initial throat of the diffuser o Or the radius R o , calculate the thickness ΔD that needs to be cut off at the leading edge of the blade and the diameter D of the small circle at the new throat M Or the radius R M : ΔD=ΔA / H D M =D o +ΔD R M =R o +0.5ΔD Step 3: After determining the leading edge cutting thickness ΔD, starting from the new leading edge point after cutting, use a straight line transition to determine the blade part that needs to be cut off. H1 is the inlet width between adjacent blades of the radial diffuser, and H 2i is the blade width between adjacent blade cutting boundaries of the radial diffuser, set H1 = H 2i =H; Step 4: During the assembly phase, first fix the radial diffuser body (1'), and then install the diffuser split structure (4') and the cover plate (3'); Step 5: When performance degradation occurs and stabilization is required, the radial diffuser is disassembled, and the lower cover plate (3') and the stabilization split structure (4') are disassembled in turn. The stabilization split structure is removed and discarded, and the stabilization split structure without the small blades and only the root is installed, and reassembled according to the requirements of step 3. At this time, the throat area of the leading edge of the radial diffuser blade has been increased.
7. The method for manufacturing an aircraft engine radial diffuser structure according to claim 6, characterized in that: In step 2, assume that the radial cutting length of the blade meridian plane is L1 and the meridian chord length of the blade is L T , L1≯0.1L T .