Design method of acoustic lining hole of aero-engine combustion chamber and cap cover and aero-engine

Abstract

The application discloses a kind of acoustic lining hole design method of aero-engine combustion chamber, cap and aero-engine, aero-engine combustion chamber includes outer casing, inner casing and cap, the outer casing, inner casing and cap form a sound lining cavity, the cap is provided with several acoustic lining holes, the sound lining cavity is communicated with the combustion cavity of aero-engine combustion chamber by the acoustic lining hole.The aero-engine combustion chamber is communicated by being provided with several acoustic lining holes on the cap, to form the porous sound lining sound damping structure, which suppresses the pressure oscillation in the combustion chamber, while also having better sound absorption characteristics in the target frequency range, and can also regulate the air at the inlet of the flame tube to alleviate the large-scale counter-current pulsation of the main flow in the main flow direction when the oscillatory combustion occurs in the combustion chamber, reducing impact and noise.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aero-engines, in particular to a design method of acoustic lining holes of a combustion chamber and a cap cover of an aero-engine. BACKGROUND

[0002] With the improvement of environmental awareness, it has become the focus of research and development designers to minimize the pollutants emitted by aero-engines during combustion. The European Aviation Research and Innovation Advisory Council (ACARE) has set future targets for carbon dioxide and nitrogen oxide emissions in the Flight path 2050 report. The goal is to reduce carbon dioxide emissions by 75% and nitrogen oxide emissions by 90% by 2050. The main factor in reducing nitrogen oxides (NOx) is to reduce the temperature of the combustion flame, which will adversely affect the efficiency of the combustion and carbon dioxide emissions. In order to achieve lower nitrogen oxide (NOx) emissions without increasing the concentration of carbon monoxide and unburned hydrocarbons in the exhaust gas, lean premixed combustion has been more researched and applied in gas turbines and aero-engines.

[0003] Compared with the traditional diffusion flame combustion chamber, the lean premixed combustion chamber is prone to oscillatory combustion, the main reasons are as follows: first, the combustion process deviates from the chemical equilibrium state, and small equivalence ratio fluctuations will cause large heat release rate changes. Second, there are fewer cooling air and dilution air holes, which reduces the absorption of acoustic pressure fluctuations on the combustion chamber wall by cooling air. Third, the premixed combustion mode causes the mixed fuel and air to be easily affected by upstream flow unsteady fluctuations and aero-engine combustion chamber unsteady fluctuations, resulting in changes in equivalence ratio fluctuations. When heat release rate fluctuations and pressure fluctuations form a positive feedback, combustion instability occurs.

[0004] Based on the above three reasons, to suppress combustion instability, the following two aspects can be considered. First, add an acoustic liner to the combustion chamber. An acoustic liner is a structural device with several through holes of a certain diameter on its surface and a certain geometric space on its back. When minute pressure pulsations cause airflow velocity pulsations, the airflow passes through the small holes on the acoustic liner surface, forming vortices, and further attenuates in the space on the back of the acoustic liner. This principle absorbs the weak aerodynamic motion generated by the pressure waves produced by thermoacoustic oscillations, thereby weakening the amplitude of pressure pulsations driven by combustion and producing the effect of thermoacoustic oscillation pressure pulsations. Second, rectify the incoming airflow to the combustion chamber to eliminate the possibility of triggering oscillatory combustion caused by upstream inlet air pulsations. In current combustion chamber designs, due to the short flow distance and large swirl number of the swirler, the short-distance deflection of airflow in the swirler blades easily causes airflow separation, forming various vortices. These vortices will continuously detach from the swirler blades, causing unsteady changes in the air in the swirler channel, triggering fuel-air ratio pulsations after fuel injection. Therefore, it is necessary to regulate the flow field characteristics of the airflow in the hydrocyclone or suppress the unsteady flow phenomena caused by the airflow in the hydrocyclone.

[0005] Therefore, suppressing pressure pulsations during oscillating combustion when the engine is operating in a lean combustion chamber, thereby reducing thermal and acoustic instability in aero-engine combustion, is a key focus for R&D designers. Summary of the Invention

[0006] The technical problem to be solved by the present invention is to overcome the defect of unstable combustion in the flame combustion chamber of aero-engines in the prior art, and to provide a design method for the acoustic liner hole of the combustion chamber and hood of an aero-engine, as well as an aero-engine.

[0007] The present invention solves the above-mentioned technical problems through the following technical solution:

[0008] An aero-engine combustion chamber includes an outer casing, an inner casing, and a cap, wherein the outer casing, inner casing, and cap form a sound liner cavity, and the cap is provided with a plurality of sound liner holes, and the sound liner cavity is connected to the combustion chamber of the aero-engine combustion chamber through the sound liner holes.

[0009] In the scheme, the aero-engine combustion chamber is communicated with the acoustic lining cavity surrounded by the combustion chamber, the outer casing, the inner casing and the cap cover through the acoustic lining holes arranged on the cap cover, thereby forming a porous acoustic lining sound damping structure which has better sound absorption characteristics in the target frequency range while suppressing pressure oscillation in the combustion chamber; and the airflow partially flows into the head swirler from the acoustic lining holes in the process of flowing through the cap cover and the acoustic lining cavity, so as to compensate for the problem of uneven air distribution of the swirler caused by the part of the swirler being blocked by the cap cover, reduce the risk of oscillating combustion caused by unsteady airflow pulsation at the swirler inlet, and regulate the air at the inlet of the flame tube; at the same time, the acoustic lining holes can also act as bleed holes to relieve large-scale reverse flow pulsation of the main combustion stage airflow in the main flow direction when oscillating combustion occurs in the combustion chamber, so that the pulsating airflow at the inlet of the combustion chamber is dissipated in different directions in the acoustic lining cavity through the acoustic lining holes, reducing the impact and noise.

[0010] Preferably, the acoustic lining holes are arranged on the outer circumferential surface of the cap cover.

[0011] Preferably, the acoustic lining holes are arranged on the side surface of the cap cover.

[0012] In the scheme, the above structure is arranged to form a channel in the radial direction and / or the axial direction of the flame tube, thereby improving the sound damping and flow suppression effect.

[0013] Preferably, the acoustic lining holes have the same diameter, and the acoustic lining holes are uniformly and spacedly arranged on the cap cover.

[0014] In the scheme, the above structure is arranged to balance the stress on the cap cover.

[0015] Preferably, the acoustic lining holes include first acoustic lining holes and second acoustic lining holes, the diameter of the first acoustic lining holes is larger than that of the second acoustic lining holes, the first acoustic lining holes and the second acoustic lining holes are uniformly and spacedly arranged along the circumferential direction of the cap cover, and the first acoustic lining holes and the second acoustic lining holes are spacedly arranged along the axial direction of the cap cover.

[0016] In the scheme, the above structure is arranged, and the first acoustic lining holes are mainly used for airflow regulation, and the second acoustic lining holes are mainly used for sound damping effect.

[0017] Preferably, the diameter of the first acoustic lining holes is not less than 4mm, and the diameter of the second acoustic lining holes is not less than 2mm.

[0018] Preferably, the porosity of the cap cover is less than 0.3.

[0019] In the scheme, the above structure is arranged to suppress airflow pulsation in the oscillating combustion frequency range of 300-1500Hz.

[0020] Preferably, the number of the outer casing, the inner casing and the cap is two, so as to form two sound lining cavities respectively located at the inner side and the outer side of the combustion cavity.

[0021] In this scheme, the combustion cavity is arranged between two sound lining cavities, and each sound lining cavity is annular.

[0022] A design method of a sound lining hole of a cap, the design method being used for manufacturing the cap in the aero-engine combustion chamber as described above, and the design method comprising the steps of:

[0023] S1, calculating the acoustic impedance ζ of the cap according to preset parameters of the sound lining hole, and the calculation formula of the acoustic impedance ζ is as follows:

[0024]

[0025] wherein ρ is the gas density, c is the gas sound speed, σ is the porosity, μ is the fluid viscosity, ω is the angular frequency, h is the sound lining hole height, d is the equivalent diameter of the sound lining hole, Mb is the sound lining flow Mach number, Mg is the sweeping flow Mach number, k=ω / c, and i is the imaginary symbol;

[0026] S2, calculating the sound reflection coefficient R by using the value of the acoustic impedance ζ, and the calculation formula of the sound reflection coefficient R is as follows:

[0027]

[0028] S3, calculating the sound absorption coefficient α by using the value of the sound reflection coefficient R, and the calculation formula of the sound absorption coefficient α is as follows:

[0029] α=1-|R| 2 ;

[0030] S4, comparing the sound absorption coefficient α with the design parameters of the combustion chamber, if the sound absorption coefficient α is not less than the design parameters of the aero-engine combustion chamber, then entering step S5, otherwise, adjusting and updating the preset parameters of the sound lining hole, and entering step S1;

[0031] S5, machining the sound lining hole on the cap according to the preset parameters.

[0032] In this scheme, the sound absorption coefficient α is compared with the design parameters of the combustion chamber, so as to judge the damping effect of the cap, and through the above steps, it can be verified whether the sound lining hole on the cap meets the design requirements, and when the preset parameters of the sound lining hole meet the requirements of the damping effect, the hole is machined on the cap, thereby reducing the trial production times of the cap, improving the machining efficiency of the cap, and further obtaining the aero-engine combustion chamber meeting the preset damping characteristics.

[0033] Preferably, the preset parameters of the acoustic holes include the porosity of the cap, the equivalent diameter of the acoustic holes, the height of the acoustic holes, and the spacing between adjacent acoustic holes.

[0034] An aero-engine comprising the aero-engine combustion chamber as described above.

[0035] On the basis of common knowledge in the art, the above-mentioned preferred conditions can be combined arbitrarily to obtain preferred examples of the present application.

[0036] The positive progress effect of the present application is that: the aero-engine combustion chamber is connected with the acoustic cavity surrounded by the combustion chamber, the outer casing, the inner casing and the cap through the acoustic holes arranged on the cap, thereby forming a porous acoustic lining sound damping structure which has better sound absorption characteristics in the target frequency range while suppressing pressure oscillation in the combustion chamber; and in the process of airflow passing through the cap and the acoustic cavity, part of the airflow flows into the head vortex finder from the acoustic holes, which compensates for the problem of uneven air distribution of the vortex finder caused by the part of the vortex finder being blocked by the cap, reduces the risk of oscillatory combustion caused by unsteady airflow pulsation at the inlet of the vortex finder, and regulates the air at the inlet of the flame tube; at the same time, the acoustic holes can also act as relief holes to alleviate the large-scale reverse flow pulsation of the main combustion stage airflow in the main flow direction when oscillatory combustion occurs in the combustion chamber, so that the pulsating airflow at the inlet of the combustion chamber is dissipated in different directions in the acoustic cavity through the acoustic holes, reducing the impact and noise. BRIEF DESCRIPTION OF DRAWINGS

[0037] Figure 1 It is a schematic diagram of the sector cross-sectional structure of the aero-engine combustion chamber of embodiment 1 of the present application (one).

[0038] Figure 2 It is a schematic diagram of the local structure of the annular combustion chamber sector of the aero-engine combustion chamber of embodiment 1 of the present application.

[0039] Figure 3 It is a schematic diagram of the cross-sectional structure of the aero-engine combustion chamber of embodiment 1 of the present application (one).

[0040] Figure 4 It is a schematic diagram of the cross-sectional structure of the aero-engine combustion chamber of embodiment 1 of the present application (two).

[0041] Figure 5 It is a schematic diagram of the local structure of the cap of embodiment 1 of the present application.

[0042] Figure 6 It is a flowchart of the acoustic hole design method of the cap of embodiment 1 of the present application.

[0043] Figure 7 It is a schematic diagram of the local structure of the cap of embodiment 2 of the present application.

[0044] Reference Signs List:

[0045] outer casing 1a, outer casing 1b

[0046] inner casing 2a, inner casing 2b

[0047] sound lining cavity 3a, sound lining cavity 3b

[0048] swirler 4a, swirler 4b

[0049] cap 5a, cap 5b

[0050] main combustion stage passage 6

[0051] fuel pipe 7

[0052] diffuser 8

[0053] sound lining hole 10

[0054] first sound lining hole 101

[0055] second sound lining hole 102

[0056] combustion chamber 20

[0057] flame tube 30

[0058] gas flow direction 50

[0059] region A

[0060] region B DETAILED DESCRIPTION

[0061] The present application will be more clearly understood and more fully appreciated by referring to the following detailed description in conjunction with the accompanying drawings.

[0062] As Figures 1-5 shown in the drawings, the present embodiment discloses an aero-engine combustion chamber, which comprises an outer casing 1a, an inner casing 2a and a cap 5a. The outer casing 1a is sleeved outside the inner casing 2a, and the cap 5a is installed at the inlet end of the inner casing 2a. The outer casing 1a, the inner casing 2a and the cap 5a enclose a sound lining cavity 3a, which has an annular structure surrounding the inner casing 2a. The cap 5a is provided with a plurality of sound lining holes 10, and the sound lining cavity 3a is in communication with a combustion chamber 20 of the aero-engine combustion chamber through the sound lining holes 10.

[0063] In the present embodiment, the number of outer casings, inner casings and caps is two, and the two outer casings, two inner casings and two caps form two acoustic cavities, and the two acoustic cavities 3a and 3b are respectively located on the inner side and the outer side of the combustion chamber 20. In addition, the swirler 4a and the swirler 4b are also installed between the end portions of the two inner casings 2a and 2b, and the two outer casings, two inner casings, two caps and swirler form a flame tube 30 structure. The outer casing 1b, the inner casing 2b and the cap 5b surround the acoustic cavity 3b, which is in the form of an annular structure around the outer casing 1b. The cap 5b is provided with a plurality of acoustic holes 10, and the acoustic cavity 3b is in communication with the combustion chamber 20 of the aero-engine combustion chamber through the acoustic holes 10.

[0064] In the present embodiment, the aero-engine combustion chamber is in communication with the acoustic cavity 3a surrounded by the combustion chamber 20, the outer casing 1a, the inner casing 2a and the cap 5a by providing a plurality of acoustic holes 10 on the cap 5a, thereby forming a porous acoustic damping structure, which has better sound absorption characteristics in the target frequency range while suppressing pressure oscillation in the combustion chamber 20; and in the process of flowing through the cap 5a and the acoustic cavity 3a, part of the airflow flows into the head swirler 4a from the acoustic holes 10, which is used to compensate for the problem of uneven air distribution of the swirler 4a caused by the part of the swirler 4a being blocked by the cap 5a, to reduce the risk of oscillatory combustion caused by unsteady airflow pulsation at the inlet of the swirler 4a, to regulate the airflow at the inlet of the flame tube 30; at the same time, the acoustic holes 10 can also act as a bleed hole to relieve the large-scale reverse flow pulsation of the main combustion stage airflow in the main flow direction when oscillatory combustion occurs in the combustion chamber 20, so that the pulsating airflow at the inlet of the combustion chamber 20 is dissipated in different directions in the acoustic cavity 3a through the acoustic holes 10, reducing the impact and noise. The gas flow direction 50 is shown by the arrow in Figure 1 .

[0065] Since the acoustic cavities 3a and 3b are both in communication with the combustion chamber 20, by the same reasoning, the acoustic cavity 3b has the same effect as the acoustic cavity 3a, which is used to compensate for the problem of uneven air distribution on the upper and lower sides of the swirler caused by the part of the swirler being blocked by the cap, to reduce the risk of oscillatory combustion caused by unsteady airflow pulsation at the inlet of the swirler.

[0066] In the present embodiment, the outer peripheral surface and the side surface of the cap 5a and the cap 5b are provided with acoustic holes 10 to form channels in the radial direction and the axial direction of the flame tube 30, thereby improving the sound attenuation and flow suppression effects.

[0067] In other alternative embodiments, acoustic holes can also be provided on the outer peripheral surface or the side surface of the cap to form a sound attenuation structure or a flow channel in a certain direction.

[0068] As Figure 5As shown, in order to facilitate the equalization of the stress on the caps 5a and 5b, the sound lining holes 10 have the same diameter, and the sound lining holes 10 are evenly spaced on the cap 5a.

[0069] In the embodiment, the porosity of the caps 5a and 5b is less than 0.3, and the frequency range of the airflow pulsation that is effectively inhibited by the oscillating combustion is 300-1500 Hz.

[0070] The gas flow direction 50 is shown by the arrows in Figure 1 , Figure 3 and Figure 4 . In the embodiment, as shown in Figure 1 and Figure 2 , the gas flow at the inlet enters the combustion chamber from the diffuser 8. After the sudden expansion section, the gas flow is divided into three streams, the middle stream of air directly opposite the flame tube 30 first passes through the sound lining holes 10 on the cap 5a and the sound lining holes 10 on the cap 5b, enters the primary combustion stage swirler 4a, and the other two streams of air are guided by the caps 5a and 5b and flow into the sound lining cavities 3a and 3b, respectively, and then pass through the wall cooling holes of the inner casing 2a and 2b to enter the combustion chamber 20, and finally the three streams of gas flow together into the combustion chamber 20 of the flame tube 30 for combustion. The flow rate of the gas flow can be calculated in proportion to the total hole area on the cap and the cross-sectional area of the sound lining cavity on the outside.

[0071] As shown in Figure 3 and Figure 4 , the fuel enters the swirler 4a and 4b through the fuel pipe 7, partially mixes in the swirler, and chemically reacts at regions A and B, and the burned gas is discharged from the outlet of the flame tube 30. As shown in Figure 3 , in the axial direction, the primary combustion stage gas flow directly passes through the caps 5a and 5b, enters the axial primary combustion stage swirler 4a, and enters the combustion chamber 20 for combustion. As shown in Figure 4 , in the radial direction, the primary combustion stage gas flow passes through the caps 5a and 5b, is deflected, enters the swirler 4b, and then enters the combustion chamber 20 for combustion.

[0072] As shown in Figure 3 and Figure 4 , the sound lining holes 10 on the cap 5a communicate the combustion chamber 20 with the sound lining cavity 3a on the outside, and the sound lining holes 10 on the cap 5b communicate the combustion chamber 20 with the sound lining cavity 3b on the inside. When the flame tube 30 of the combustion chamber oscillates and burns, the gas flow can flow out through the sound lining holes 10 on the caps 5a and 5b to relieve the large-scale gas flow pulsation in the axial direction of the flame tube 30, and has a certain damping effect on the pulsation of the entire flow field caused by the oscillating combustion.

[0073] The embodiment also discloses an aero-engine comprising the aero-engine combustion chamber as described above.

[0074] As Figure 6 shown, the embodiment also discloses a design method of a hat cover acoustic baffle hole, which is used for manufacturing the hat cover in the aero-engine combustion chamber as described above, and the design method comprises the steps of:

[0075] S1, calculating the acoustic impedance ζ of the hat cover according to preset parameters of the acoustic baffle hole, and the calculation formula of the acoustic impedance ζ is as follows:

[0076]

[0077] wherein ρ is the gas density, c is the gas sound speed, σ is the porosity, μ is the fluid viscosity, ω is the angular frequency, h is the acoustic baffle hole height, d is the acoustic baffle hole equivalent diameter, Mb is the acoustic baffle flow Mach number, Mg is the sweeping flow Mach number, k = ω / c, and i is the imaginary symbol;

[0078] S2, calculating the acoustic reflection coefficient R by using the value of the acoustic impedance ζ, and the calculation formula of the acoustic reflection coefficient R is as follows:

[0079]

[0080] S3, calculating the sound absorption coefficient α by using the value of the acoustic reflection coefficient R, and the calculation formula of the sound absorption coefficient α is as follows:

[0081] α = 1 - |R| 2 ;

[0082] S4, if the sound absorption coefficient α is not less than the design parameter of the aero-engine combustion chamber, then entering step S5, otherwise adjusting and updating the preset parameters of the acoustic baffle hole, and entering step S1;

[0083] S5, processing the acoustic baffle hole on the hat cover according to the preset parameters.

[0084] The preset parameters of the acoustic baffle hole include the porosity of the hat cover, the acoustic baffle hole equivalent diameter, the acoustic baffle hole height, and the spacing between adjacent acoustic baffle holes. In the method, the sound absorption coefficient α is compared with the design parameter of the combustion chamber to judge the damping effect of the hat cover. The sound absorption coefficient and the design parameter of the aero-engine combustion chamber are compared to judge the damping effect of the hat cover. If the conclusion of the sound absorption effect is that the design requirement cannot be met, then the preset parameters of the acoustic baffle hole on the hat cover are adjusted, and then recalculated until the requirements are met. Compared with directly drilling holes on the hat cover, the above design method can reduce the trial production times of the hat cover, improve the processing efficiency of the hat cover, and then obtain the aero-engine combustion chamber meeting the preset damping characteristics.

[0085] Embodiment 2

[0086] As Figure 7As shown, the embodiment is basically the same as that of Embodiment 1, except that the sound insulation holes 10 on the cap 5a include first sound insulation holes 101 and second sound insulation holes 102, the aperture of the first sound insulation holes 101 is larger than that of the second sound insulation holes 102, the first sound insulation holes 101 and the second sound insulation holes 102 are uniformly spaced along the circumferential direction of the cap respectively, and the first sound insulation holes 101 and the second sound insulation holes 102 are arranged in a spaced manner along the axial direction of the cap.

[0087] The first sound insulation holes 101 are mainly used for gas rectification, and the second sound insulation holes 102 are mainly used for the effect of sound damping. Preferably, the aperture of the first sound insulation holes 101 is not less than 4 mm, and the aperture of the second sound insulation holes 102 is not less than 2 mm. The sound insulation holes with different apertures optimize the uniformity of the gas flow after the gas flow passes through the cap from upstream, and reduce the non-uniform pulsation of the flame zone in the flame tube when the combustion chamber is stably working. Of course, the sound insulation holes on the two caps can be set to the same or different hole structures.

[0088] Although the specific embodiments of the present application are described above, those skilled in the art should understand that this is only an example, the protection scope of the present application is defined by the appended claims. Those skilled in the art can make various changes or modifications to these embodiments without departing from the principles and essence of the present application, and these changes and modifications all fall within the protection scope of the present application.

Claims

1. An aircraft engine combustion chamber, characterized in that, It includes an outer casing, an inner casing, and a cap, which together form a sound liner cavity. The cap is provided with a plurality of sound liner holes, and the sound liner cavity is connected to the combustion chamber of the aero-engine combustion chamber through the sound liner holes. The acoustic liner holes include a first acoustic liner hole and a second acoustic liner hole. The diameter of the first acoustic liner hole is larger than that of the second acoustic liner hole. The first acoustic liner hole and the second acoustic liner hole are evenly spaced along the circumferential direction of the cap, and the first acoustic liner hole and the second acoustic liner hole are spaced along the axial direction of the cap.

2. The aero-engine combustion chamber as described in claim 1, characterized in that, The acoustic lining hole is located on the outer peripheral surface of the cap; And / or, the acoustic liner hole is located on the side of the cap.

3. The aero-engine combustion chamber as described in claim 1, characterized in that, The sound liner holes all have the same diameter, and several of the sound liner holes are evenly spaced on the cap.

4. The aircraft engine combustion chamber as described in claim 1, characterized in that, The diameter of the first acoustic liner hole is not less than 4 mm, and the diameter of the second acoustic liner hole is not greater than 2 mm.

5. The aero-engine combustion chamber as described in claim 1, characterized in that, The porosity of the cap is less than 0.

3.

6. The aircraft engine combustion chamber as described in claim 1, characterized in that, The number of the outer casing, the inner casing, and the cap are all two, forming two acoustic lining cavities, which are located on the inner and outer sides of the combustion chamber, respectively.

7. A method for designing the sound liner holes of a hat cover, characterized in that, The acoustic liner design method for the hood is used to manufacture a hood in the combustion chamber of an aero-engine as described in any one of claims 1-6, the design method comprising the steps of: S1. Calculate the acoustic impedance ζ of the cap based on the preset parameters of the acoustic liner hole. The formula for calculating the acoustic impedance ζ is as follows: ， Where ρ is the gas density, c is the gas sound velocity, σ is the porosity, μ is the fluid viscosity, ω is the angular frequency, h is the height of the acoustic liner orifice, d is the equivalent diameter of the acoustic liner orifice, and M... b For the acoustic lining deflection Mach number, M g For the sweeping Mach number, k = ω / c, where i is the imaginary number. S2. Calculate the sound reflection coefficient R using the value of the acoustic impedance ζ. The formula for calculating the sound reflection coefficient R is as follows: ； S3. Calculate the sound absorption coefficient α using the value of the sound reflection coefficient R. The formula for calculating the sound absorption coefficient α is as follows: α＝1-|R| 2 ； S4. Compare the sound absorption coefficient α with the design parameters of the combustion chamber. If the sound absorption coefficient α is not less than the design parameters of the aero-engine combustion chamber, proceed to step S5. Otherwise, adjust and update the preset parameters of the acoustic liner hole, and proceed to step S1. S5. Machine the sound liner holes on the cap according to the preset parameters.

8. The method for designing the sound liner hole of the cap as described in claim 7, characterized in that, The preset parameters of the acoustic liner holes include the porosity of the cap, the equivalent diameter of the acoustic liner holes, the height of the acoustic liner holes, and the spacing between adjacent acoustic liner holes.

9. An aircraft engine, characterized in that, It includes the aircraft engine combustion chamber as described in any one of claims 1-6.

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