A method for designing a super-sonic rotor blade elementary airfoil, a rotor blade and an axial compressor

Abstract

The application relates to a supersonic rotor blade basic element airfoil design method, a rotor blade and an axial flow compressor, and belongs to the technical field of compressor blade design. Three sections of front, middle and rear are divided by adopting two shock wave surfaces in a rotor blade basic element channel, each section turning angle and relative Mach number are calculated, the rotor blade basic element airfoil is geometrically modeled based on the section turning angles, and the geometric modeling is verified and fine-tuned based on the relative Mach numbers, so that the problems of low efficiency and complex process of the supersonic blade basic element airfoil design method of the existing axial flow compressor are solved.

Description

Technical Field

[0001] This invention relates to the field of compressor blade design technology, and in particular to a method for designing the basic airfoil of a supersonic rotor blade, rotor blades, and an axial compressor. Background Technology

[0002] Currently, axial compressors for aero engines continue to evolve towards ultra-high loads, reducing the axial and radial dimensions of the compressor to further improve its thrust-to-weight ratio or power-to-weight ratio. Increasing the rotor blade tangential velocity can increase the load on the axial compressor stage, but it also leads to a higher relative Mach number at the rotor blade tips, further enhancing shock wave intensity and resulting in relatively high shock wave losses. After the airflow passes through the strong shock wave surface, the differences in airflow velocity abruptly in different directions cause significant airflow deflection, increasing the derailment angle of the airflow on the rotor blade tip suction surface, and even leading to flow separation, further increasing flow losses and affecting the matching with the next blade row.

[0003] Improving the curvature distribution of the basic airfoil profile of an axial compressor supersonic rotor blade is crucial for controlling the shock wave surface and improving the flow before and after the shock wave. Current technologies often focus on the impact of the pre-compression airfoil geometry parameters on performance and the design input parameters for supersonic compressor blades. However, there is a lack of corresponding solutions for the Mach number, shock wave, and deflection variations within the blade channel. This leads to a wide range of possible shapes when designing the basic airfoil profile of the rotor blade, increasing design difficulty and requiring repeated adjustments and calculations, resulting in low design efficiency and a complex process. Summary of the Invention

[0004] Based on the above analysis, the present invention aims to provide a method for designing the basic airfoil profile of a supersonic rotor blade, a rotor blade, and an axial compressor, in order to solve the problems of low efficiency and complex process in existing supersonic compressor blade basic airfoil profile design methods.

[0005] On one hand, embodiments of the present invention provide a method for designing the basic airfoil of a supersonic rotor blade, generating a preliminary profile and design parameters for the basic airfoil of the rotor blade, dividing the basic channel of the rotor blade into three segments—front, middle, and rear—based on the first and second shock surfaces through which the airflow direction passes between the blades, and designing the basic airfoil of the rotor blade through the following steps:

[0006] The relative Mach number of the incoming flow in the front section is calculated based on the rotor blade basic airfoil design parameters.

[0007] Calculate the total turning angle and the corresponding relative Mach number for each region in sequence;

[0008] Based on the design parameters and the total turning angle of each region, the basic airfoil of the rotor blade is geometrically modeled;

[0009] Based on the obtained relative Mach numbers, verify whether each part of the geometric shape meets the index requirements, and fine-tune the parts that do not meet the requirements until the rotor blade basic airfoil is finally obtained.

[0010] The beneficial effects of the above technical solution are as follows: By segmenting the rotor blade element channel and calculating the total turning angle and corresponding relative Mach number of each region in sequence, the geometric shape, verification and fine-tuning of the rotor blade element airfoil are performed. Compared with the existing technology, which has the problem of broad shape design when designing rotor blade element airfoil because the Mach number, shock wave and turning changes of the airflow inside the blade channel are not quantitatively analyzed during the design, the embodiment of the present invention quantifies the changes of airflow through the blade element channel, realizes the verification of the geometric shape and more precise fine-tuning, effectively improves design efficiency and reduces design complexity.

[0011] Based on a further improvement of the above method, the relative Mach number of the incoming flow in the front section is calculated using the following formula based on the rotor blade element airfoil design parameters:

[0012] In the formula, ω is the angular velocity, r is the radius of the basic plane airfoil, and c z α1 is the inlet axial velocity of the blade, γ is the inlet absolute airflow angle, R is the adiabatic index, T is the air constant, T is the inlet static temperature, and β1 is the inlet relative airflow angle.

[0013] The beneficial effect of the above-mentioned further improvement scheme is that it can calculate the relative Mach number of the incoming flow in the front section, laying the foundation for subsequent quantitative calculations in the blade element channel.

[0014] Based on a further improvement of the above method, the step of sequentially calculating the total turning angle and the corresponding relative Mach number of each region specifically refers to sequentially obtaining the compression angle of the current region based on the relative Mach number of the previous position point for the front exit, the rear of the first shock surface, the middle exit, and the rear of the second shock surface, and then calculating the total turning angle of the current region based on the compression angle of the current region, and finally calculating the relative Mach number of the current position point based on the total turning angle of the current region.

[0015] The total turning angle of the current region is calculated based on the current region's compression angle using the following formula, which is specifically expressed as follows:

[0016] In the formula,

[0017] θ N This is the current region compression angle, where N represents the region that the airflow sequentially passes through from the rotor blade element channel inlet.

[0018] δ NIt is the total turning angle of the current region.

[0019] M rev，Lpre It is the relative Mach number of the previous position point, where L represents the position points through which the airflow passes sequentially from the inlet of the rotor blade element channel, and Lpre refers to the position point before L.

[0020] The beneficial effects of the above-mentioned further improvement scheme are: it can achieve efficient geometric modeling by quantifying the current region compression angle and the current region total turning angle obtained through calculation, and can verify and fine-tune the geometric modeling based on the relative Mach number of the current position point, effectively improving design efficiency and reducing design complexity; it can calculate the current region turning angle of the corresponding segment in the rotor blade element channel, thereby quantifying the change of the turning angle of the airflow entering the rotor blade element channel when it reaches the corresponding segment of the channel, so that the design efficiency can be improved based on the quantitative calculation results when performing geometric modeling of the professional blade element airfoil, and can be used for subsequent calculation of the corresponding current relative Mach number.

[0021] Based on a further improvement of the above method, the relative Mach number of the current position point is calculated using the following formula based on the total turning angle of the current region. The formula is specifically expressed as follows:

[0022] In the formula, M rev，L It is the relative Mach number of the current position point.

[0023] The beneficial effects of the above-mentioned further improvement scheme are: it is possible to calculate the relative Mach number of the current position point of the corresponding segment in the rotor blade element channel, thereby quantifying the change of the relative Mach number of the airflow entering the rotor blade element channel when it reaches the position point of the corresponding segment of the channel, so that the design efficiency can be improved based on the quantitative calculation results when verifying and fine-tuning the geometric shape of the professional blade element.

[0024] A further improvement to the above method, specifically, involves geometrically shaping the rotor blade element airfoil based on the rotor blade element airfoil design parameters and the total turning angle of each region, including:

[0025] A suction surface profile with multiple control points is generated based on B-spline curves, and the turning angles of each control point in the front, middle, and rear sections are obtained based on the total turning angle of each region. The curvatures corresponding to the turning angles of each control point are then distributed sequentially on the suction surface profile.

[0026] Set the thickness of the blade element;

[0027] Generate a mid-arc line with multiple control points based on B-spline curves;

[0028] The blade element is used to generate a pressure surface profile including multiple control points based on a Bezier curve;

[0029] The suction surface, pressure surface, leading edge contact point, trailing edge contact point, and mid-arc leading edge endpoint and trailing edge endpoint of the blade element are constrained.

[0030] The beneficial effects of the above-mentioned further improvement scheme are: it achieves a more efficient geometric design for the rotor blade element airfoil compared to the existing technology. Specifically, by constraining the suction surface, pressure surface, leading edge contact point, trailing edge contact point, and leading edge and trailing edge endpoints of the blade element airfoil based on the blade element airfoil design parameters, it avoids excessive divergence in the geometric design of the blade element airfoil; it generates the suction surface profile and mid-arc line with multiple control points based on B-spline curves for the blade element, which facilitates subsequent fine-tuning and avoids the impact of adjusting a single control point. This causes a change in the overall shape of the suction surface profile; the turning angles of each control point are calculated and the curvatures corresponding to the turning angles of each control point are sequentially distributed onto the suction surface profile, enabling the suction surface profile to be basically close to the design requirements through a single geometric modeling, avoiding the need for significant adjustments to the geometric modeling later; the pressure surface profile including multiple control points is generated for the blade element based on the Bezier curve, and the thickness of the blade element airfoil is obtained, thus obtaining the geometric modeling of the rotor blade element airfoil more efficiently than the prior art.

[0031] Based on a further improvement of the above method, the turning angle of each control point in the front segment is calculated using the following formula, which is specifically expressed as:

[0032] In the formula, δ A It is the total turning angle of the front section. It is the geometric angle of the aforementioned front entrance control point. It is the geometric angle of the aforementioned front exit control point, δ i It is the turning angle corresponding to each control point in the preceding segment, where i is an integer greater than or equal to 1, and as i increases, δ i Increasing sequentially.

[0033] The beneficial effect of the above-mentioned further improvement scheme is that it can obtain the turning angle of each control point in the front section, which facilitates efficient subsequent geometric modeling.

[0034] Based on a further improvement of the above method, the turning angle of each control point in the middle section is calculated using the following formula, which is specifically expressed as:

[0035] In the formula, δ B δ1 is the total turning angle of the middle section, δ2 is the total turning angle of the first shock surface, and δ2 is the total turning angle of the second shock surface. It is the geometric angle of the mid-section exit control point, δ j It is the turning angle corresponding to the control point of the suction surface profile in the middle section, where j is an integer greater than or equal to 1, and as j increases, δ j Decreasing sequentially.

[0036] The beneficial effect of the above-mentioned further improvement scheme is that it can obtain the geometric angles of each control point in the middle section, which facilitates efficient subsequent geometric modeling.

[0037] Based on a further improvement of the above method, the turning angle of each control point in the latter segment is calculated using the following formula, which is specifically expressed as:

[0038] In the formula, δ C It is the total turning angle of segment C mentioned above. It is the geometric angle of the downstream exit control point, δ k It is the turning angle corresponding to the control point of the suction surface profile of the rear section, where k is an integer greater than or equal to 1, and as k increases, δ k Increasing sequentially.

[0039] The beneficial effect of the above-mentioned further improvement scheme is that it can obtain the geometric turning angle of each control point in the latter section, which facilitates efficient subsequent geometric modeling.

[0040] On the other hand, embodiments of the present invention also provide a supersonic rotor blade for an axial compressor, wherein the rotor blade is designed using a supersonic rotor blade basic airfoil design method provided in embodiments of the present invention.

[0041] The beneficial effects of the above technical solution are as follows: The axial compressor supersonic rotor blade designed by the supersonic rotor blade basic airfoil design method provided by the present invention has a smoother relative Mach number gradient distribution and better supersonic aerodynamic performance compared with the prior art.

[0042] This invention also provides an axial compressor, which includes a supersonic rotor blade for an axial compressor provided in this invention.

[0043] The beneficial effects of the above technical solution are as follows: the axial flow compressor has better performance in dealing with supersonic speeds compared with existing axial flow compressors of the same type.

[0044] In this invention, the above-described technical solutions can be combined with each other to achieve more preferred combinations. Other features and advantages of this invention will be set forth in the following description, and some advantages may become apparent from the description or be learned by practicing the invention. The objects and other advantages of this invention can be realized and obtained from what is particularly pointed out in the description and drawings. Attached Figure Description

[0045] The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Throughout the drawings, the same reference numerals denote the same parts.

[0046] Figure 1 This is a schematic diagram of the distribution of blade profile parameters of the supersonic rotor blade element in Embodiment 1 of the present invention.

[0047] Figure 2 This is a schematic diagram of the constraint between the endpoint of the arc line and the junction of the suction surface profile and the pressure surface profile in the basic blade shape of the supersonic rotor blade of Embodiment 1 of the present invention.

[0048] Figure 3 This is a schematic diagram comparing the relative Mach number gradient distribution of the supersonic rotor blade profile designed by the method of Embodiment 1 of the present invention with that of the existing blade profile design. Detailed Implementation

[0049] Preferred embodiments of the present invention will now be described in detail with reference to the accompanying drawings, which form part of this application and are used together with the embodiments of the present invention to illustrate the principles of the present invention, but are not intended to limit the scope of the present invention.

[0050] Example 1:

[0051] A specific embodiment of the present invention discloses a method for designing the basic airfoil of a supersonic rotor blade, such as... Figure 1 As shown.

[0052] The preliminary profile and design parameters of the rotor blade element airfoil are generated. Based on the first and second shock wave surfaces through which the airflow passes between the blades, the rotor blade element channel is divided into three segments: front, middle, and rear. The rotor blade element airfoil is then designed through the following steps:

[0053] The relative Mach number of the incoming flow in the front section is calculated based on the rotor blade basic airfoil design parameters.

[0054] The total turning angle and corresponding relative Mach number for each region are calculated sequentially, including:

[0055] The compression angle of the current region is obtained based on the relative Mach number of the previous position point for the front exit, the rear of the first shock surface, the middle exit, and the rear of the second shock surface, respectively. The total turning angle of the current region is calculated based on the compression angle of the current region, and the relative Mach number of the current position point is calculated based on the total turning angle of the current region.

[0056] Based on the design parameters and the turning angles of each region, the basic airfoil of the rotor blade is geometrically modeled;

[0057] Based on the obtained relative Mach numbers, verify whether each part of the geometric shape meets the index requirements, and fine-tune the parts that do not meet the requirements until the rotor blade basic airfoil is finally obtained.

[0058] Typically, the design of compressor rotor blade basic airfoil begins with generating a preliminary profile and design parameters of the rotor blade through quasi-three-dimensional design based on the design objectives. The preliminary profile includes the shape of the rotor blade basic airfoil adopted to achieve the required thrust-to-weight ratio and relative Mach number of the airflow. The design parameters are parameters of the rotor blade basic airfoil generated through the design requirements, such as: angular velocity, basic plane airfoil radius, blade inlet axial velocity, inlet absolute airflow angle, adiabatic index, air constant, blade inlet static temperature, and blade inlet relative airflow angle.

[0059] like Figure 1 As shown, the axial compressor supersonic rotor blade basic airfoil design method disclosed in this embodiment adopts a two-shock wave design, specifically, as follows: Figure 1 As shown, the rotor blade element channel is divided into three sections—front, middle, and rear—according to the airflow direction within the rotor blade element channel. The front section (… Figure 1 The middle section (marked as section A) is the pre-compression section for a higher relative Mach number inflow, suitable for supersonic airflow with a relative Mach number greater than 1.4; the middle section ( Figure 1 The section marked B in the middle is the section for enhanced compression and suction surface rectification, suitable for flows with relative Mach numbers of 1.1 to 1.3; the subsequent section ( Figure 1 The section marked C is the outlet rectifier section, which basically ensures the bending angle of the rotor element airfoil and the outflow angle of the airflow to achieve aerodynamic matching with the rear blades; the first shock wave surface is located at the junction of the front section outlet and the middle section inlet, and the second shock wave surface is located at the junction of the middle section outlet and the rear section inlet; in order to meet the requirements of airflow variation of the three sections and the first and second shock wave surfaces in the rotor blade element channel, the rotor blade element airfoil adopts an "S" shape design.

[0060] When designing the rotor blade basic airfoil, according to the relevant theories of supersonic flow and airfoil structure disturbance, as the airflow passes through the front outlet, the first shock surface, the middle outlet, and the second shock surface in sequence, the supersonic speed of the incoming flow in the front section gradually decreases, and it decreases to subsonic speed when it enters the rear inlet after the second shock surface. Therefore, combining the design parameters obtained through the quasi-three-dimensional design and the preliminary profile of the blade basic airfoil, the current compression angle is obtained based on the previous relative Mach number for the front outlet, the first shock surface, the middle outlet, and the second shock surface in sequence, and the current turning angle is calculated based on the current compression angle. Finally, the current relative Mach number is calculated based on the current turning angle.

[0061] It should be noted that "previous" specifically refers to the previous position point that the airflow passes through sequentially in the blade element channel, and "current" specifically refers to the current position point that the airflow arrives at sequentially in the blade element channel. The position points are as follows: front inlet is marked A1, front outlet (i.e., in front of the first shock wave) is marked A2, behind the first shock wave (i.e., mid-section inlet) is marked B1, in front of the second shock wave (i.e., mid-section outlet) is marked B2, behind the second shock wave (i.e., rear inlet) is marked C1, and rear outlet is marked C2. The areas that the airflow passes through sequentially in the blade element channel are as follows: front section is marked A, mid section is marked B, rear section is marked C, first shock wave is marked W1, and second shock wave is marked W2.

[0062] Specifically, the relative Mach number of the incoming flow in the front section is calculated based on the rotor blade element airfoil design parameters using the following formula, which is specifically expressed as:

[0063] In the formula, ω is the angular velocity, r is the radius of the basic plane airfoil, and c z α1 is the inlet axial velocity of the blade, γ is the inlet absolute airflow angle, R is the adiabatic index, T is the air constant, T is the inlet static temperature, and β1 is the inlet relative airflow angle.

[0064] The relative Mach number M of the incoming flow in the front section rev，A1 The relative Mach number of the airflow entering the inlet is used to determine the current compression angle θ of the inlet. This is based on the Mach number-compression angle-turning angle characteristic chart of the compression wave. A .

[0065] Furthermore, based on the current region's compression angle, the total turning angle of the current region is calculated using the following formula, which is specifically expressed as:

[0066] In the formula,

[0067] θ N It is the current region compression angle, where N refers to the region identifiers that the airflow passes through sequentially from the inlet of the rotor blade element channel, N∈[A,W1,B,W2,C];

[0068] δ N It is the total turning angle of the current area;

[0069] M rev，Lpre It is the relative Mach number of the previous position point, where L represents the position points that the airflow passes through sequentially from the inlet of the rotor blade element channel, L∈[A1,A2,B1,B2,C1,C2], and Lpre refers to the previous position point of L.

[0070] For example, such as Figure 1As shown, for the front segment, the compression angle based on the front segment region is expressed as θ. A The total turning angle of the front segment, calculated using the above formula, is expressed as δ. A Specifically, calculate δ A The specific expression of the formula is as follows:

[0071]

[0072] Furthermore, based on the total turning angle of the current region, the relative Mach number of the current position point is calculated using the following formula, which is specifically expressed as:

[0073] In the formula, M rev，L It is the relative Mach number of the current position point.

[0074] Specifically, based on δ A Then, applying the above formula, the relative Mach number of the front exit is calculated as M. rev，A2 , indicating the relative Mach number M of the incoming flow in the preceding section. rev，A1 After passing through a compression wave in the front region, the relative Mach number M at the front exit is reduced. rev，A2 Calculate M rev，A2 The formula is specifically expressed as:

[0075]

[0076] Next, M rev，A2 After passing the first shock surface, a relative Mach number change occurs. Similarly, based on the Mach number-compression angle-turning angle characteristic chart of the compression wave, the compression angle of the first shock surface is expressed as θ. W1 .

[0077] At this point, the airflow is still in a supersonic state. Therefore, the formula for calculating the total turning angle of the current region based on the current region's compression angle still applies. Specifically, the turning angle after calculating the first shock surface is expressed as δ. W1 The specific expression of the formula is as follows:

[0078] Based on δ W1 The relative Mach number after calculating the first shock surface is expressed as M. rev，B1 Specifically, calculate M rev，B1 The specific expression of the formula is as follows:

[0079]

[0080] Similarly, by analogy, the mid-section compression angle can be successively obtained as θ. B The total turning angle of the middle section is expressed as δ. B The mid-section exit Mach number is denoted as M.rev，B2 The compression angle of the second shock surface is expressed as θ. W2 The total turning angle of the second shock surface is expressed as δ. W2 The relative Mach number behind the second shock surface is denoted as M. rev，C1 .

[0081] Next, based on the rotor blade element airfoil design parameters and each of the aforementioned turning angles, the rotor blade element airfoil is geometrically modeled, specifically including:

[0082] A suction surface profile with multiple control points is generated based on B-spline curves, and the turning angles of each control point in the front, middle, and rear sections are obtained based on the total turning angles of each region. The curvatures corresponding to the turning angles of each control point are then distributed sequentially onto the suction surface profile.

[0083] Set the thickness of the blade element;

[0084] Generate a mid-arc line with multiple control points based on B-spline curves;

[0085] The blade element is used to generate a pressure surface profile including multiple control points based on a Bezier curve;

[0086] The suction surface, pressure surface, leading edge contact point, trailing edge contact point, and mid-arc leading edge endpoint and trailing edge endpoint of the blade element are constrained.

[0087] Specifically, the geometric modeling of the basic airfoil of a supersonic rotor blade includes shaping the suction surface profile, pressure surface profile, and intermediate arc line of the basic airfoil; determining the thickness of the basic airfoil; and constraining the leading edge contact point and trailing edge contact point of the suction and pressure surfaces, as well as constraining the leading edge endpoint and trailing edge endpoint of the intermediate arc line. The geometric modeling process of the basic airfoil of the rotor blade is very flexible and can be freely combined as needed.

[0088] For example, the following illustrates one embodiment of the basic blade geometry of the supersonic rotor blade, specifically including:

[0089] Step 1: Generate a suction surface profile with multiple control points for the blade element based on a B-spline curve, and obtain the turning angles of each control point in the front, middle and rear sections. Then, distribute the curvatures corresponding to the turning angles of each control point onto the suction surface profile in sequence.

[0090] The geometric modeling involves sequentially distributing the curvature of each control point onto the generated curve using a curve generation method.

[0091] Specifically, the turning angles of the front segment are calculated using the following formula, which is expressed as follows:

[0092] In the formula, δ A It is the total turning angle of the front section. It is the geometric angle of the aforementioned front entrance control point. It is the geometric angle of the aforementioned front exit control point, δ i It is the turning angle corresponding to each control point in the preceding segment, where i is an integer greater than or equal to 1, and as i increases, δ i In ascending order. It is one of the design parameters of the rotor blade basic airfoil, which is obtained through quasi-three-dimensional design and the selected angle of attack before designing the rotor blade basic airfoil.

[0093] From the inlet of the rotor blade element airfoil to the first shock surface, the total turning angle of the leading section, i.e., the turning angle δ of the leading section, is... A The pressure surface is distributed in ascending order from the front inlet to the front outlet (the front outlet being the front of the first shock wave) onto the suction surface profile and the pressure surface profile.

[0094] The turning angles of each control point in the middle section are calculated using the following formula, which is specifically expressed as follows:

[0095] In the formula, δ B δ1 is the total turning angle of the middle section, δ2 is the total turning angle of the first shock surface, and δ2 is the total turning angle of the second shock surface. It is the geometric angle of the mid-section exit control point, δ j It is the turning angle corresponding to the control point of the suction surface profile in the middle section, where j is an integer greater than or equal to 1, and as j increases, δ j Decreasing sequentially.

[0096] The middle section is located between the first shock wave surface and the second shock wave surface. When the airflow passes through both the first and second shock wave surfaces, abrupt changes occur. The suction surface of the rotor blade's basic airfoil will correspondingly deform to adapt to the shock wave changes. Specifically, the total turning angle δ of the middle section... B It needs to be superimposed with the first shock surface turning angle δ1 and the second shock surface turning angle δ2, and distributed on the suction surface line in a regular decreasing manner from the middle section inlet to the middle section outlet.

[0097] When the airflow reaches the inlet of the rear section, i.e. the second shock surface, the airflow has attenuated to subsonic speed. In order to better adapt to the rear exhaust blades, the geometric design of the rear section inlet to the rear section outlet needs to be smooth, until it attenuates to the relative Mach number required by the airflow outlet design.

[0098] Specifically, the turning angles of each control point in the latter segment are calculated using the following formula, which is expressed as follows:

[0099] In the formula, δ C It is the total turning angle of segment C mentioned above. It is the geometric angle of the downstream exit control point, δ k It is the turning angle corresponding to the control point of the suction surface profile of the rear section, where k is an integer greater than or equal to 1, and as k increases, δ k In ascending order. It is one of the design parameters of the rotor blade basic airfoil, which is obtained through quasi-three-dimensional design and the selected lag angle before designing the rotor blade basic airfoil.

[0100] Step 2: Set the initial value of the blade element thickness based on experience, and generate the mid-arc line of the blade based on the B-spline curve.

[0101] Step 3: Based on the arc line in the blade and the initial value of the blade element thickness, generate a pressure surface profile including multiple control points for the blade element based on the Bezier curve.

[0102] Step 4: As Figure 2 As shown, the suction surface profile, pressure surface profile, and mid-curve of the basic airfoil are constrained. The suction surface profile is labeled S, the pressure surface profile is labeled P, and the mid-curve is labeled M, where A S1 A represents the point where the leading edges of the suction surface profile meet. M1 A represents the endpoint of the leading edge of the middle arc. P1 A represents the point where the leading edges of the pressure surface profile meet. Sn A represents the point where the trailing edges of the suction surface lines meet. Mn Indicates the point where the trailing edges of the middle arc meet, A Pn This indicates the point where the tail edges of the pressure surface profile meet, specifically including:

[0103] The constraint element blade's suction surface and pressure surface's front edge contact point are specifically constrained to have equal positions, slopes, and curvatures at the first control point of the suction surface at the front inlet and the first control point of the pressure surface at the front inlet.

[0104] The connection point between the trailing edges of the suction and pressure surfaces of the constrained basic blade is specifically constrained so that the suction surface at the last control point of the rear outlet and the pressure surface at the last control point of the rear outlet have equal positions, slopes, and curvatures.

[0105] The endpoint of the leading edge of the constrained arc line specifically corresponds to the junction of the front edges of the suction surface and the pressure surface. It is the first control point of the leading edge position of the arc line. The position and slope of the endpoint of the leading edge of the arc line are constrained based on the junction of the front edges of the suction surface and the pressure surface.

[0106] The endpoint of the trailing edge of the middle arc is constrained, specifically corresponding to the junction of the trailing edges of the suction surface and the pressure surface. It is the last control point for the position of the trailing edge of the middle arc. The position and slope of the endpoint of the trailing edge of the middle arc are constrained based on the junction of the trailing edges of the suction surface and the pressure surface.

[0107] Furthermore, based on the obtained relative Mach numbers, it is verified whether each part of the geometric shape meets the index requirements. Fine adjustments are made to the parts that do not meet the requirements until the rotor blade basic airfoil is finally obtained.

[0108] Specifically, the relative Mach number corresponding to each segment of the rotor blade element channel calculated through the embodiments of the present invention is a theoretical value obtained by calculating the flow and deflection of the airflow in each segment of the rotor blade element channel, and is correlated with each deflection angle related to the geometric shape of the rotor blade element. In the prior art, the airflow and relative Mach number changes in the rotor blade element channel are not quantified, and only the relative Mach number of the incoming flow in the front section and the relative Mach number of the outlet section in the rear section are considered. This results in a very wide range of adjustments when shaping the geometric shape of the rotor blade element, and the adjustment is very difficult. In this embodiment, a theoretical baseline value is established by calculating the relative Mach numbers within the rotor blade element channel. After completing the geometric model, the geometric model is substituted into a software tool to obtain the relative Mach numbers of each segment within the rotor blade element channel within the geometric model. Compared with the theoretical baseline value of the relative Mach number, the difference is basically very small. Then, some control points that do not meet the standard are fine-tuned. Specifically, the control points that do not meet the standard refer to those whose relative Mach numbers differ significantly from the theoretical baseline value of the relative Mach number. Generally, the control points that do not meet the standard are located on the suction surface profile and pressure surface profile of the rotor blade element airfoil geometry. The curvature of the control points that do not meet the standard is adjusted accordingly so that their corresponding relative Mach numbers are close to or reach the theoretical baseline value of the relative Mach number, thereby obtaining a better relative Mach number gradient change.

[0109] Figure 3 The figure shows the relative Mach number distribution of the basic airfoil of a supersonic rotor blade in different regions of the blade, designed by the prior art and the embodiment of the present invention, respectively. The left side is the design of the prior art, and the right side is the design of the embodiment of the present invention. The left side of both is the leading edge of the blade, and the right side is the trailing edge of the blade. As can be seen from the figure, the basic airfoil of the embodiment of the present invention on the right side has a smoother gradient of relative Mach number decrease in each region compared with the prior art design.

[0110] This embodiment discloses a method for designing the basic airfoil of a supersonic rotor blade. Based on the first and second shock surfaces through which the airflow passes between the blades, the rotor blade basic channel is divided into three segments: front, middle, and rear. The current compression angle is obtained sequentially at the outlet of the front segment, after the first shock surface, the outlet of the middle segment, and after the second shock surface, based on the previous relative Mach number. The current turning angle is then calculated based on the current compression angle, and the current relative Mach number is calculated based on the current turning angle. Finally, the rotor blade basic airfoil is geometrically modeled and fine-tuned based on each turning angle and the rotor blade basic airfoil design parameters. Compared to existing technologies, this embodiment calculates theoretical baseline values ​​for airflow changes and relative Mach numbers within the rotor blade basic channel, significantly improving efficiency and reducing design complexity during geometric modeling and fine-tuning. This results in a smoother relative Mach number gradient change when supersonic airflow flows through the designed airfoil, enhancing the airfoil's ability to handle supersonic airflow and improving the matching of the rear blades from a geometric design perspective.

[0111] Example 2:

[0112] This embodiment discloses a supersonic rotor blade for an axial compressor. The rotor blade is designed using a supersonic rotor blade basic airfoil design method disclosed in Embodiment 1. This embodiment includes all the technical features of Embodiment 1.

[0113] Compared with existing supersonic rotor blades, the supersonic rotor blades disclosed in this embodiment make the relative Mach number gradient change smoother when the supersonic airflow flows in the supersonic rotor blades disclosed in this embodiment, have better ability to cope with supersonic airflow, and have better matching with the rear blades.

[0114] Example 3:

[0115] This embodiment discloses an axial flow compressor, which includes the supersonic rotor blades of an axial flow compressor disclosed in Embodiment 2, and includes all the technical features of Embodiment 2.

[0116] The axial compressor disclosed in this embodiment has better supersonic performance compared to existing axial compressors of the same type.

[0117] Those skilled in the art will understand that all or part of the processes of the methods described in the above embodiments can be implemented by a computer program instructing related hardware, and the program can be stored in a computer-readable storage medium. The computer-readable storage medium may be a disk, optical disk, read-only memory, or random access memory, etc.

[0118] The above description is only a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for designing the basic airfoil shape of a supersonic rotor blade, characterized in that, The preliminary profile and design parameters of the rotor blade element airfoil are generated. Based on the first and second shock wave surfaces through which the airflow passes between the blades, the rotor blade element channel is divided into three segments: front, middle, and rear. The rotor blade element airfoil is then designed through the following steps: The relative Mach number of the incoming flow in the front section is calculated based on the rotor blade basic airfoil design parameters. Calculate the total turning angle and the corresponding relative Mach number for each region in sequence; Based on the design parameters and the total turning angle of each region, the basic airfoil of the rotor blade is geometrically modeled; Based on the obtained relative Mach numbers, verify whether each part of the geometric shape meets the index requirements, and fine-tune the parts that do not meet the requirements until the rotor blade basic airfoil is finally obtained.

2. The method for designing the basic airfoil of a supersonic rotor blade according to claim 1, characterized in that, Based on the rotor blade element airfoil design parameters, the relative Mach number of the incoming flow in the front section is calculated using the following formula, which is specifically expressed as follows: In the formula, It is the relative Mach number of the airflow entering the front inlet. It's angular velocity. The radius of the basic planar airfoil. The axial velocity at the blade inlet. The absolute airflow angle at the inlet. The adiabatic index, air constant, To ensure the inlet is at a quiet temperature. The inlet relative airflow angle.

3. The method for designing the basic airfoil of a supersonic rotor blade according to claim 1, characterized in that, The step of sequentially calculating the total turning angle and corresponding relative Mach number of each region specifically refers to obtaining the compression angle of the current region based on the relative Mach number of the previous position point for the front exit, the rear of the first shock surface, the middle exit, and the rear of the second shock surface, respectively, and then calculating the total turning angle of the current region based on the compression angle of the current region, and finally calculating the relative Mach number of the current position point based on the total turning angle of the current region. The total turning angle of the current region is calculated based on the current region's compression angle using the following formula, which is specifically expressed as follows: In the formula, This is the current region compression angle, where N represents the region that the airflow sequentially passes through from the rotor blade element channel inlet. It is the total turning angle of the current region. It is the relative Mach number of the previous position point, where L represents the position points through which the airflow passes sequentially from the inlet of the rotor blade element channel, and Lpre refers to the position point before L.

4. The method for designing the basic airfoil of a supersonic rotor blade according to claim 3, characterized in that, Based on the total turning angle of the current region, the relative Mach number of the current position point is calculated using the following formula, which is specifically expressed as follows: In the formula, It is the relative Mach number of the current position point.

5. The method for designing the basic airfoil of a supersonic rotor blade according to claim 1, characterized in that, The geometric modeling of the rotor blade element airfoil based on the rotor blade element airfoil design parameters and the total turning angle of each region specifically includes: A suction surface profile with multiple control points is generated based on B-spline curves, and the turning angles of each control point in the front, middle, and rear sections are obtained based on the total turning angle of each region. The curvatures corresponding to the turning angles of each control point are then distributed sequentially on the suction surface profile. Set the thickness of the blade element; Generate a mid-arc line with multiple control points based on B-spline curves; The blade element is used to generate a pressure surface profile including multiple control points based on a Bezier curve; The suction surface, pressure surface, leading edge contact point, trailing edge contact point, and mid-arc leading edge endpoint and trailing edge endpoint of the blade element are constrained.

6. The method for designing the basic airfoil of a supersonic rotor blade according to claim 5, characterized in that, The turning angles of each control point in the preceding segment are calculated using the following formula, which is specifically expressed as follows: In the formula, It is the total turning angle of the front section. It is the geometric angle of the aforementioned front entrance control point. It is the geometric angle of the aforementioned front exit control point. These are the turning angles corresponding to each control point in the preceding segment, where i is an integer greater than or equal to 1. As i increases, Increasing sequentially.

7. The method for designing the basic airfoil of a supersonic rotor blade according to claim 5, characterized in that, The turning angles of each control point in the middle section are calculated using the following formula, which is specifically expressed as follows: In the formula, It is the total turning angle of the middle section. It is the total turning angle of the first shock wave surface. It is the total deflection angle of the second shock surface. It is the geometric angle of the mid-section exit control point. It is the turning angle corresponding to the control point of the suction surface profile in the middle section, where j is an integer greater than or equal to 1. As j increases, Decreasing sequentially.

8. The method for designing the basic airfoil of a supersonic rotor blade according to claim 5, characterized in that, The turning angles of each control point in the latter segment are calculated using the following formula, which is specifically expressed as follows: In the formula, It is the total turning angle of the latter section. It is the geometric angle of the downstream exit control point. It is the turning angle corresponding to the control point of the rear suction surface profile, where k is an integer greater than or equal to 1. As k increases, Increasing sequentially.

9. A supersonic rotor blade for an axial compressor, characterized in that, The rotor blades are designed using the supersonic rotor blade basic airfoil shaping design method according to any one of claims 1-8.

10. An axial flow compressor, characterized in that, The axial compressor includes the supersonic rotor blades of the axial compressor as described in claim 9.

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