Supersonic rotor blade primitive blade profile design method, rotor blade and axial flow compressor

Abstract

The invention relates to a supersonic rotor blade primitive blade profile design method, a rotor blade and an axial flow compressor, and belongs to the technical field of compressor blade design. Two shock wave surfaces are adopted in a rotor blade primitive channel to divide the interior of the rotor blade primitive channel into a front segment, a middle segment and a rear segment, and the turning angle and the relative Mach number of each segment are calculated respectively; and geometric modeling is carried out on the rotor blade primitive blade profile on the basis of the segmented turning angles, and the problems that in the prior art, an axial flow compressor supersonic blade primitive blade profile design method is low in efficiency and complex in process are solved on the basis of verification and fine adjustment of the geometric modeling of the relative Mach numbers.

Description

Technical Field

[0001] The invention relates to the technical field of compressor blade design, and in particular to a supersonic rotor blade primitive blade profile design method, a rotor blade and an axial flow compressor. Background Art

[0002] At present, the axial compressor of aircraft engines continues to develop in the direction of ultra-high load, reducing the axial and radial dimensions of the compressor, thereby further improving its thrust-to-weight ratio or power-to-weight ratio. Increasing the tangential speed of the rotor blades can achieve an increase in the load of the axial compressor stage, but it will also lead to a higher relative Mach number at the tip of the rotor blades, further increasing the shock wave intensity and causing relatively high shock wave losses. After the airflow passes through the strong shock wave surface, due to the difference in the airflow velocity jump in each direction, the airflow is significantly deflected, the derailment angle of the airflow on the suction surface of the rotor blade tip increases, and even causes flow separation, further increasing the flow loss and affecting the matching with the next blade row.

[0003] Improving the curvature distribution of the blade profile of the axial flow compressor supersonic rotor blade is the key to controlling the shock wave surface and improving the flow in front of and behind the shock wave surface. In the prior art, the design focus is often on the study of the influence of the pre-compression blade profile geometry parameters of the compressor supersonic blade on the performance and the design input parameters, but there is a lack of corresponding solutions for the Mach number, shock wave and deflection changes of the airflow inside the blade slot, resulting in a wide shape when designing the rotor blade profile, increasing the design difficulty, requiring repeated adjustments and calculations, resulting in low design efficiency and complex processes. Summary of the invention

[0004] In view of the above analysis, an embodiment of the present invention aims to provide a supersonic rotor blade primitive blade profile design method, a rotor blade and an axial flow compressor, so as to solve the problems of low efficiency and complex process of the existing compressor supersonic blade primitive blade profile design method.

[0005] On the one hand, an embodiment of the present invention provides a method for designing a supersonic rotor blade primitive profile, generating a preliminary outline and design parameters of the rotor blade primitive profile, dividing the rotor blade primitive channel into three sections: front, middle, and rear based on a first shock wave surface and a second shock wave surface through which an airflow passes in a flow direction between blades, and designing the rotor blade primitive profile by the following steps:

[0006] The relative Mach number of the front section incoming flow is calculated based on the rotor blade primitive blade profile design parameters;

[0007] Calculate the total turning angle and the corresponding relative Mach number of each area in turn;

[0008] Performing geometric modeling on the rotor blade primitive blade profile based on the design parameters and the total turning angle of each region;

[0009] Based on the obtained relative Mach numbers, it is verified whether each part of the geometric shape meets the index requirements, and the parts that do not meet the requirements are fine-tuned until the rotor blade primitive blade profile is finally obtained.

[0010] The beneficial effects of the above technical solution are as follows: by segmenting the rotor blade primitive slot, and calculating the total turning angle and the corresponding relative Mach number of each area in turn, it is used for geometric modeling, verification and fine-tuning of the rotor blade primitive blade profile. Compared with the prior art, which does not perform quantitative analysis on the Mach number, shock wave and turning changes of the airflow inside the blade slot during design, resulting in a wide modeling problem when designing the rotor blade primitive blade profile, the embodiment of the present invention quantifies the changes of the airflow through the blade primitive slot, realizes the verification of the geometric modeling and more precise fine-tuning, effectively improves the design efficiency and reduces the design complexity.

[0011] Based on the further improvement of the above method, the relative Mach number of the front section incoming flow is calculated based on the rotor blade primitive blade design parameters by the following formula, and the formula is specifically expressed as:

[0012] Where ω is the angular velocity, r is the radius of the element plane blade, c z is the blade inlet axial velocity, α1 is the inlet absolute airflow angle, γ is the adiabatic index, R 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 further improvement scheme is that the relative Mach number of the front section incoming flow can be calculated, laying the foundation for the subsequent quantitative calculation in the blade element slot.

[0014] Based on the further improvement of the above method, the total turning angle of each area and the corresponding relative Mach number are calculated in sequence, specifically referring to obtaining the compression angle of the current area based on the relative Mach number of the previous position point for the front section outlet, the first shock wave surface, the middle section outlet, and the second shock wave surface, and calculating the total turning angle of the current area based on the compression angle of the current area, and then calculating the relative Mach number of the current position point based on the total turning angle of the current area;

[0015] The total turning angle of the current region is calculated based on the compression angle of the current region by the following formula, and the formula is specifically expressed as:

[0016] In the formula,

[0017] θ N is the compression angle of the current region, where N refers to the region identifier that the airflow passes through in sequence from the inlet of the rotor blade primitive slot,

[0018] δ Nis the total turning angle of the current area,

[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 in sequence from the inlet of the rotor blade element slot, and Lpre refers to the previous position point of L.

[0020] The beneficial effects of the above-mentioned further improvement scheme are: efficient geometric modeling can be achieved through the calculated quantitative results of the current area compression angle and the current area total turning angle, and the geometric modeling can be verified and more accurately fine-tuned based on the relative Mach number of the current position point, effectively improving the design efficiency and reducing the design complexity; the current area turning angle of the corresponding segment in the rotor blade primitive slot can be calculated, thereby quantifying the turning angle change of the airflow entering the rotor blade primitive slot when it reaches the corresponding segment of the slot, so that the subsequent geometric modeling of the professional blade primitive blade profile can improve the design efficiency based on the quantitative calculation results, and can be used for the subsequent calculation of the corresponding current relative Mach number.

[0021] Based on the further improvement of the above method, the relative Mach number of the current position point is calculated based on the total turning angle of the current area by the following formula, and the formula is specifically expressed as:

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

[0023] The beneficial effect of the above-mentioned further improvement scheme is that it can calculate the relative Mach number of the current position point of the corresponding segment in the rotor blade primitive slot, thereby quantifying the relative Mach number change of the airflow entering the rotor blade primitive slot when it reaches the corresponding segment position point of the slot, so that the design efficiency can be improved based on the quantitative calculation results when the geometric shape of the professional blade primitive blade is subsequently verified and fine-tuned.

[0024] Based on the further improvement of the above method, the geometric modeling of the rotor blade primitive profile based on the rotor blade primitive profile design parameters and the total turning angle of each region specifically includes:

[0025] Generate a suction surface profile including a plurality of control points based on a B-spline curve, obtain the turning angles of the control points of the front section, the middle section and the rear section based on the total turning angles of the regions, and then distribute the curvatures corresponding to the turning angles of the control points to the suction surface profile in sequence;

[0026] Setting the thickness of the blade primitive profile;

[0027] Generate a median arc line including a plurality of control points based on a B-spline curve;

[0028] Generating a pressure surface profile including a plurality of control points for the blade primitive based on a Bezier curve;

[0029] The suction surface of the blade element, the connection point of the front edge and the connection point of the pressure surface, the connection point of the trailing edge, and the leading edge endpoint and the trailing edge endpoint of the mid-camber line are constrained.

[0030] The beneficial effect of the above-mentioned further improvement scheme is: it realizes a more efficient geometric modeling design of the rotor blade primitive blade profile compared with the prior art. Specifically, based on the blade primitive blade profile design parameters, the blade primitive suction surface, the pressure surface front edge connection point, the trailing edge connection point, and the leading edge endpoint and trailing edge endpoint of the mid-arc line can avoid excessive divergence of the geometric modeling of the blade primitive blade profile; the suction surface profile line and the mid-arc line including multiple control points are generated for the blade primitive based on the B-spline curve, so as to avoid the adjustment of a certain control point during subsequent fine-tuning. The overall shape of the suction surface profile line is changed; the turning angles of the control points are calculated and the curvatures corresponding to the turning angles of the control points are distributed sequentially on the suction surface profile line, so that the suction surface profile line can be basically close to the design requirements through one geometric modeling, avoiding the need to make major adjustments to the geometric modeling later; the pressure surface profile line including multiple control points is generated for the blade primitive based on the Bezier curve to obtain the thickness of the blade primitive profile, so as to obtain the geometric modeling of the rotor blade primitive profile more efficiently than the prior art.

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

[0032] In the formula, δ A is the total turning angle of the front section, is the geometric angle of the front-end entry control point, is the geometric angle of the front section exit control point, δ i is the turning angle corresponding to each control point in the front section, where i is an integer greater than or equal to 1. As i increases, δ i In increasing order.

[0033] The beneficial effect of the above further improved solution is that the turning angle of each control point in the front section can be obtained, which is convenient for subsequent efficient geometric modeling.

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

[0035] In the formula, δ B is the total turning angle of the middle section, δ1 is the total turning angle of the first shock wave surface, δ2 is the total turning angle of the second shock wave surface, is the geometric angle of the middle section exit control point, δ j is the turning angle corresponding to the control point of the middle suction surface profile, where j is an integer greater than or equal to 1. As j increases, δ j In descending order.

[0036] The beneficial effect of the above further improved solution is that the geometric angles of each control point in the middle section can be obtained, which is convenient for subsequent efficient geometric modeling.

[0037] Based on the further improvement of the above method, the turning angle of each control point in the rear section is calculated by the following formula, which is specifically expressed as:

[0038] In the formula, δ C is the total turning angle of the middle C segment, is the geometric angle of the control point at the rear exit, δ k 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, δ k In increasing order.

[0039] The beneficial effect of the above further improved solution is that the geometric turning angles of each control point in the rear section can be obtained, which is convenient for subsequent efficient geometric modeling.

[0040] On the other hand, an embodiment of the present invention further provides a supersonic rotor blade for an axial compressor, wherein the rotor blade is designed by a supersonic rotor blade primitive blade profile design method provided by an embodiment of the present invention.

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

[0042] An embodiment of the present invention further provides an axial flow compressor, and the axial flow compressor includes an axial flow compressor supersonic rotor blade also provided in an embodiment of the present invention.

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

[0044] In the present invention, the above-mentioned technical solutions can also be combined with each other to achieve more preferred combination solutions. Other features and advantages of the present invention will be described in the subsequent description, and some advantages can become obvious from the description, or can be understood by practicing the present invention. The purpose and other advantages of the present invention can be realized and obtained through the contents particularly pointed out in the description and the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] The accompanying drawings are only used for the purpose of illustrating specific embodiments and are not to be considered as limiting the present invention. In the entire drawings, the same reference symbols represent the same components;

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

[0047] Figure 2 This is a schematic diagram of the constraints on the connection points between the endpoints of the arc line of the basic element blade profile of a supersonic rotor blade in Example 1 of the present invention and the suction surface profile line and the pressure surface profile line.

[0048] Figure 3 Schematic diagram comparing the relative Mach number gradient distribution of the supersonic rotor blade profile designed by the method of Example 1 of the present invention and the existing designed blade profile. DETAILED DESCRIPTION

[0049] The preferred embodiments of the present invention are described in detail below in conjunction with the accompanying drawings, wherein the accompanying drawings constitute a 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 used to limit the scope of the present invention.

[0050] Embodiment 1:

[0051] A specific embodiment of the present invention discloses a method for designing a supersonic rotor blade element profile. Figure 1 shown.

[0052] Generate a preliminary outline and design parameters of the rotor blade primitive blade profile, divide the rotor blade primitive channel into three sections: front, middle and rear based on the first shock wave surface and the second shock wave surface through which the airflow passes between the blades, and design the rotor blade primitive blade profile through the following steps:

[0053] The relative Mach number of the front section incoming flow is calculated based on the rotor blade primitive blade profile design parameters;

[0054] Calculate the total turning angle and the corresponding relative Mach number of each area in turn, including:

[0055] Obtaining the current region compression angle based on the relative Mach number of the previous position point for the front section exit, the first shock wave surface, the middle section exit, and the second shock wave surface, respectively, and calculating the current region total turning angle based on the current region compression angle, and then calculating the current position point relative Mach number based on the current region total turning angle;

[0056] Performing geometric modeling on the rotor blade primitive profile based on the design parameters and the turning angles of each region;

[0057] Based on the obtained relative Mach numbers, it is verified whether each part of the geometric shape meets the index requirements, and the parts that do not meet the requirements are fine-tuned until the rotor blade primitive blade profile is finally obtained.

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

[0059] like Figure 1 As shown, the method for designing the basic blade profile of the axial flow compressor supersonic rotor blade disclosed in this embodiment adopts a two-wave shock wave design. Specifically, Figure 1 As shown, the rotor blade primitive channel is divided into three sections: front, middle and rear according to the flow direction of the airflow in the rotor blade primitive channel, wherein the front section ( Figure 1 The middle section (marked as section A) is the pre-compression section of the incoming flow with a relatively high relative Mach number, which is suitable for supersonic airflow with a relative Mach number greater than 1.4; the middle section ( Figure 1 The section marked as B in the middle is the section with enhanced compression and suction surface, which is suitable for the flow with relative Mach number of 1.1 to 1.3; the rear section ( Figure 1 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 realize the airflow flow change requirements of the three sections and the first shock wave surface and the second shock wave surface in the rotor blade element slot, the rotor blade element blade adopts an "S" type design.

[0060] When designing the rotor blade element profile, according to the supersonic flow and blade structure disturbance related theories, when the airflow passes through the front section outlet, the first shock wave surface, the middle section outlet, and the second shock wave surface in sequence, the supersonic speed of the front section flow gradually decays, and after reaching the second shock wave surface, it decays to subsonic speed when entering the rear section inlet. Therefore, combined with the design parameters obtained through the quasi-three-dimensional design and the preliminary outline of the blade element profile, the current compression angle is obtained based on the previous relative Mach number for the front section outlet, the first shock wave surface, the middle section outlet, and the second shock surface, and the current turning angle is calculated based on the current compression angle, and the current relative Mach number is calculated based on the current turning angle.

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

[0062] Specifically, the relative Mach number of the front-end incoming flow is calculated based on the rotor blade primitive blade profile design parameters by the following formula, and the formula is specifically expressed as:

[0063] Where ω is the angular velocity, r is the radius of the element plane blade, c z is the blade inlet axial velocity, α1 is the inlet absolute airflow angle, γ is the adiabatic index, R 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 front flow rev，A1 is the relative Mach number of the airflow entering the front section inlet, and the current compression angle θ of the front section is obtained based on the Mach number-compression angle-turning angle characteristic diagram of the compression wave A .

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

[0066] In the formula,

[0067] θ N is the current region compression angle, where N refers to the region identifier through which the airflow passes from the inlet of the rotor blade primitive slot in sequence, N∈[A,W1,B,W2,C];

[0068] δ N 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 where the airflow passes through the inlet of the rotor blade element slot in sequence, L∈[A1,A2,B1,B2,C1,C2], and Lpre refers to the previous position point of L.

[0070] For example, Figure 1As shown, for the front section, the compression angle based on the front section area is expressed as θ A , then the total turning angle of the front section is calculated by the above formula and expressed as δ A Specifically, calculate δ A The specific formula is expressed as:

[0071]

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

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

[0074] Specifically, based on δ A , then the relative Mach number of the front section outlet is calculated by the above formula and expressed as M rev，A2 , represents the relative Mach number M of the front flow rev，A1 After passing through a series of compression waves in the front section area, it is reduced to the relative Mach number M at the front section outlet. rev，A2 , calculate M rev，A2 The specific formula is:

[0075]

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

[0077] At this time, the airflow is still in a supersonic state. Therefore, the formula for calculating the total turning angle of the current area based on the compression angle of the current area is still applicable. Specifically, the turning angle after calculating the first shock wave surface is expressed as δ W1 The specific formula is expressed as:

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

[0079]

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

[0081] Next, geometric modeling of the rotor blade primitive profile is performed based on the rotor blade primitive profile design parameters and each of the turning angles, specifically including:

[0082] Generate a suction surface profile including a plurality of control points based on a B-spline curve, obtain the turning angles of each control point in the front section, the middle section, and the rear section based on the total turning angles of each area, and then distribute the curvatures corresponding to the turning angles of each control point in sequence to the suction surface profile;

[0083] Setting the thickness of the blade primitive profile;

[0084] Generate a median arc line including a plurality of control points based on a B-spline curve;

[0085] Generating a pressure surface profile including a plurality of control points for the blade primitive based on a Bezier curve;

[0086] The suction surface of the blade element, the connection point of the front edge and the connection point of the pressure surface, the connection point of the trailing edge, and the leading edge endpoint and the trailing edge endpoint of the mid-camber line are constrained.

[0087] Specifically, for the geometric modeling of the primitive blade of the supersonic rotor blade, it specifically includes geometric modeling of the suction surface profile, pressure surface profile, and mid-arc line of the primitive blade, determining the thickness of the primitive blade, and constraining the leading edge connection point and trailing edge connection point of the suction surface and the pressure surface, and constraining the leading edge endpoint and trailing edge endpoint of the mid-arc line. The geometric modeling process of the primitive blade of the rotor blade is very flexible and can be freely combined according to needs.

[0088] Exemplarily, one implementation of the geometric modeling of the supersonic rotor blade element blade profile is shown below, specifically including:

[0089] Step 1: Generate a suction surface profile including multiple control points for the blade primitive based on the B-spline curve, and obtain the turning angles of the control points at the front section, middle section and rear section, and then distribute the curvatures corresponding to the turning angles of the control points in sequence to the suction surface profile.

[0090] The geometric modeling is to distribute the curvature of each control point to the generated curve in sequence through the adopted curve generation method.

[0091] Specifically, the turning angles of the front section are calculated by the following formula, which is specifically expressed as:

[0092] In the formula, δ A is the total turning angle of the front section, is the geometric angle of the front-end entry control point, is the geometric angle of the front section exit control point, δ i is the turning angle corresponding to each control point in the front section, where i is an integer greater than or equal to 1. As i increases, δ i In increasing order. ; It is one of the design parameters of the rotor blade primitive profile, and is obtained through quasi-three-dimensional design and selected angle of attack before designing the rotor blade primitive profile.

[0093] From the entrance of the front section of the rotor blade element blade profile to the first shock wave surface, the total turning angle of the front section, that is, the turning angle of the front section δ A , is distributed on the suction surface profile and the pressure surface profile in a regular increasing order from small to large from the front section entrance to the front section exit (the front section exit is on the first shock wave front).

[0094] The turning angle of each control point in the middle section is calculated by the following formula, which is specifically expressed as:

[0095] In the formula, δ B is the total turning angle of the middle section, δ1 is the total turning angle of the first shock wave surface, δ2 is the total turning angle of the second shock wave surface, is the geometric angle of the middle section exit control point, δ j is the turning angle corresponding to the control point of the middle suction surface profile, where j is an integer greater than or equal to 1. As j increases, δ j In descending order.

[0096] The middle section is located between the first shock wave surface and the second shock wave surface. When the airflow passes through the first shock wave surface and the second shock wave surface, a sudden change will occur. The suction surface of the rotor blade primitive blade shape will correspondingly deform to adapt to the shock wave change. Specifically, the total turning angle δ of the middle section is B It needs to be superimposed with the first shock wave surface turning angle δ1 and the second shock wave surface turning angle δ2, and distributed on the suction surface profile line in a regular decreasing order from large to small from the middle section entrance to the middle section exit.

[0097] When the airflow reaches the inlet of the rear section, that is, the second shock wave surface, the airflow has decayed to subsonic speed. In order to better adapt to the rear blade surface, a smooth transition geometric design is required from the inlet of the rear section to the outlet of the rear section until it decays to the relative Mach number required by the airflow outlet design.

[0098] Specifically, the turning angle of each control point in the rear section is calculated by the following formula, which is specifically expressed as:

[0099] In the formula, δ C is the total turning angle of the middle C segment, is the geometric angle of the control point at the rear exit, δ k 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, δ k In increasing order. ; It is one of the design parameters of the rotor blade primitive profile, and is obtained through quasi-three-dimensional design and the selected backward angle before designing the rotor blade primitive profile.

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

[0101] Step 3: Based on the blade center arc line and the initial value of the blade base thickness, a pressure surface profile including multiple control points is generated for the blade primitive based on a Bezier curve.

[0102] Step 4: Figure 2 As shown, the suction surface profile, pressure surface profile, and camber line of the primitive blade are constrained, the suction surface profile is marked as S, the pressure surface profile is marked as P, and the camber line is marked as M, wherein A S1 Indicates the point where the front edge of the suction surface line meets, A M1 Indicates the front edge endpoint of the mid-arc line, A P1 Indicates the point where the leading edge of the pressure surface profile meets, A Sn Indicates the point where the trailing edge of the suction surface line meets, A Mn Indicates the point where the trailing edge of the mid-arc meets, A Pn Indicates the point where the trailing edge of the pressure surface profile meets, including:

[0103] Constraining the connection point between the suction surface and the front edge of the pressure surface of the primitive blade, specifically constraining the position, slope and curvature of the suction surface at the first control point of the front section entrance and the pressure surface at the first control point of the front section entrance to be equal;

[0104] Constraining the connection point of the trailing edges of the suction surface and the pressure surface of the primitive blade, specifically constraining the position, slope and curvature of the last control point of the suction surface at the outlet of the rear section and the last control point of the pressure surface at the outlet of the rear section to be equal;

[0105] The leading edge endpoint of the constrained mid-camber line specifically corresponds to the point where the suction surface and the front edge of the pressure surface meet, and is the first control point of the leading edge position of the mid-camber line. The position and slope of the leading edge endpoint of the mid-camber line are constrained based on the point where the suction surface and the front edge of the pressure surface meet;

[0106] The endpoint of the trailing edge of the constrained mid-arc line specifically corresponds to the junction point of the trailing edges of the suction surface and the pressure surface, and is the last control point of the trailing edge position of the mid-arc line. The position and slope of the endpoint of the trailing edge of the mid-arc line are constrained based on the junction point 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, and the parts that do not meet the requirements are fine-tuned until the rotor blade primitive blade profile is finally obtained.

[0108] Specifically, the relative Mach number corresponding to each segment of the rotor blade primitive channel calculated by the embodiment of the present invention is a theoretical value obtained by calculating the flow and deflection of the airflow in the rotor blade primitive channel in each segment, and has a correlation with each deflection angle related to the geometric modeling of the rotor blade primitive blade. In the prior art, the airflow flow and relative Mach number change in the rotor blade primitive channel are not quantified, and only the relative Mach number of the front section incoming flow and the relative Mach number of the rear section outlet are concerned, resulting in a wide range of adjustments for the geometric modeling of the rotor blade primitive blade, and great difficulty in adjustment. In this embodiment, a theoretical baseline value is established by calculating the relative Mach numbers in the rotor blade primitive slot. After completing the geometric modeling, the geometric modeling is substituted into the software tool to obtain the relative Mach numbers of each segment in the rotor blade primitive slot in the geometric modeling. Compared with the theoretical baseline value of the relative Mach number, the difference is basically very small, and then some control points that do not meet the standard are fine-tuned. Specifically, the control point that does not meet the standard refers to the control point whose relative Mach number is significantly different from the theoretical baseline value of the relative Mach number. Generally, the control point that does not meet the standard is specifically located on the suction surface line and the pressure surface line of the rotor blade primitive blade geometry. Then, the curvature of the control point that does not meet the standard is adjusted accordingly so that the corresponding relative Mach number is close to or reaches the theoretical baseline value of the relative Mach number, thereby obtaining a better relative Mach number gradient change.

[0109] Figure 3 The relative Mach number distribution of the supersonic rotor blade elemental blade profiles designed in the prior art and the embodiment of the present invention in different areas of the blade is shown, wherein the left side is the prior art design and the right side is the embodiment of the present invention design, the left side of both are the leading edge of the blade and the right side are the trailing edge of the blade. It can be seen from the figure that the relative Mach number drop gradient of the elemental blade profile designed in the embodiment of the present invention on the right side is smoother than that of the prior art design in each area.

[0110] The present embodiment discloses a method for designing a supersonic rotor blade primitive blade profile. The rotor blade primitive slot is divided into three sections: front, middle and rear sections based on a first shock wave surface and a second shock wave surface through which the airflow passes in the inter-blade flow direction. The current compression angle is obtained based on the previous relative Mach number for the front section outlet of the rotor blade primitive slot, the first shock wave surface, the middle section outlet and the second shock wave surface. The current turning angle is calculated based on the current compression angle, and the current relative Mach number is calculated based on the current turning angle. The rotor blade primitive blade profile is geometrically modeled and fine-tuned based on each turning angle and the rotor blade primitive blade profile design parameters. Compared with the prior art, the present embodiment obtains a theoretical baseline value for the airflow change and the relative Mach number in the rotor blade primitive slot by calculating, greatly improves efficiency and reduces design complexity in the geometric modeling and fine-tuning process, makes the relative Mach number gradient change smoother when the supersonic airflow flows in the designed blade profile, and enhances the designed blade profile's ability to cope with supersonic airflow and the matching of the rear row of blades from the geometric structure design level.

[0111] Embodiment 2:

[0112] This embodiment discloses a supersonic rotor blade for an axial compressor. The rotor blade is designed by a supersonic rotor blade primitive blade profile design method disclosed in Example 1. This embodiment includes all the technical features of Example 1.

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

[0114] Embodiment 3:

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

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

[0117] Those skilled in the art will appreciate that all or part of the processes of the above-mentioned embodiments can be implemented by instructing related hardware through a computer program, and the program can be stored in a computer-readable storage medium, wherein the computer-readable storage medium is a disk, an optical disk, a read-only storage memory, or a random access memory, etc.

[0118] The above description is only a preferred specific implementation manner of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by any technician familiar with the technical field within the technical scope disclosed by the present invention should be covered within the protection scope of the present invention.

Claims

1. A method for designing a supersonic rotor blade element profile, characterized in that: Generate a preliminary outline and design parameters of the rotor blade primitive blade profile, divide the rotor blade primitive channel into three sections: front, middle and rear based on the first shock wave surface and the second shock wave surface through which the airflow passes between the blades, and design the rotor blade primitive blade profile through the following steps: The relative Mach number of the front section incoming flow is calculated based on the rotor blade primitive blade profile design parameters; Calculate the total turning angle and the corresponding relative Mach number of each area in turn; Performing geometric modeling on the rotor blade primitive blade profile based on the design parameters and the total turning angle of each region; Based on the obtained relative Mach numbers, it is verified whether each part of the geometric shape meets the index requirements, and the parts that do not meet the requirements are fine-tuned until the rotor blade primitive blade profile is finally obtained.

2. A supersonic rotor blade element profile design method according to claim 1, characterized in that: Based on the rotor blade element profile design parameters, the front section incoming flow relative Mach number is calculated by the following formula, which is specifically expressed as: Where ω is the angular velocity, r is the radius of the element plane blade, c z is the blade inlet axial velocity, α1 is the inlet absolute airflow angle, γ is the adiabatic index, R is the air constant, T is the inlet static temperature, and β1 is the inlet relative airflow angle.

3. A supersonic rotor blade element profile design method according to claim 1, characterized in that: The sequentially calculating the total turning angle of each region and the corresponding relative Mach number 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 section exit, the first shock wave surface, the middle section exit, and the second shock wave surface, and calculating the total turning angle of the current region based on the compression angle of the current region, and then 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 compression angle of the current region by the following formula, and the formula is specifically expressed as: In the formula, θ N is the compression angle of the current region, where N refers to the region identifier that the airflow passes through in sequence from the inlet of the rotor blade primitive slot, δ N is the total turning angle of the current area, 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 in sequence from the inlet of the rotor blade element slot, and Lpre refers to the previous position point of L.

4. A supersonic rotor blade element profile design method according to claim 3, characterized in that: Based on the total turning angle of the current area, the relative Mach number of the current position point is calculated by the following formula, which is specifically expressed as: In the formula, M rev，L is the relative Mach number of the current position point.

5. The method for designing a supersonic rotor blade element profile according to claim 1, characterized in that: The step of geometrically shaping the rotor blade primitive profile based on the rotor blade primitive profile design parameters and the total turning angle of each region specifically includes: Generate a suction surface profile including a plurality of control points based on a B-spline curve, obtain the turning angles of the control points of the front section, the middle section and the rear section based on the total turning angles of the regions, and then distribute the curvatures corresponding to the turning angles of the control points to the suction surface profile in sequence; Setting the thickness of the blade primitive profile; Generate a median arc line including a plurality of control points based on a B-spline curve; Generating a pressure surface profile including a plurality of control points for the blade primitive based on a Bezier curve; The suction surface of the blade element, the connection point of the front edge and the connection point of the pressure surface, the connection point of the trailing edge, and the leading edge endpoint and the trailing edge endpoint of the mid-camber line are constrained.

6. A supersonic rotor blade element profile design method according to claim 5, characterized in that: The turning angle of each control point in the front section is calculated by the following formula, which is specifically expressed as: In the formula, δ A is the total turning angle of the front section, is the geometric angle of the front-end entry control point, is the geometric angle of the front section exit control point, δ i is the turning angle corresponding to each control point in the front section, where i is an integer greater than or equal to 1. As i increases, δ i In increasing order.

7. A supersonic rotor blade element profile design method according to claim 5, characterized in that: The turning angle of each control point in the middle section is calculated by the following formula, which is specifically expressed as: In the formula, δ B is the total turning angle of the middle section, δ1 is the total turning angle of the first shock wave surface, δ2 is the total turning angle of the second shock wave surface, is the geometric angle of the middle section exit control point, δ j is the turning angle corresponding to the control point of the middle suction surface profile, where j is an integer greater than or equal to 1. As j increases, δ j In descending order.

8. The method for designing a supersonic rotor blade element profile according to claim 5, characterized in that: The turning angle of each control point in the rear section is calculated by the following formula, which is specifically expressed as: In the formula, δ C is the total turning angle of the middle C segment, is the geometric angle of the control point at the rear exit, δ k 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, δ k In increasing order.

9. A supersonic rotor blade for an axial flow compressor, characterized in that: The rotor blade is designed by the supersonic rotor blade primitive blade profile design method according to any one of claims 1-8.

10. An axial flow compressor, characterized in that: The axial flow compressor comprises the axial flow compressor supersonic rotor blades according to claim 9.

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