A method for analyzing the angle-of-attack characteristics of a front-ledge shedding compressor blade profile

Abstract

The application discloses a method for analyzing the attack angle characteristics of a front edge drop-off compressor blade profile, comprising the following steps: constructing a front edge drop-off compressor blade profile geometry; calculating the aerodynamic characteristics of a normal compressor blade profile and a front edge drop-off compressor blade profile under a limited attack angle; and correlating the change of the aerodynamic characteristics of the front edge drop-off compressor blade profile under the limited attack angle with the change of the aerodynamic characteristics of the front edge drop-off compressor blade profile under other inlet attack angles. The application can quickly calculate the change of the aerodynamic characteristics of the front edge drop-off compressor blade profile under different attack angles on the premise of calculating the loss and the lag angle of the front edge drop-off compressor blade profile under only three attack angles, can quickly predict the influence of the front edge drop-off on the aerodynamic characteristics of the compressor blade profile, and has a wide application prospect in the evaluation of the influence of the hard object hitting the blade on the aerodynamic characteristics of the compressor.

Description

Technical Field

[0001] This invention relates to a method for predicting the impact of blade damage from hard objects on the aerodynamic characteristics of compressors, specifically to a method for analyzing the angle of attack characteristics of compressor blades with leading-edge debris. Background Technology

[0002] During the service life of aero-engines, it is inevitable that hard objects will be ingested and damage the compressor blades. Leading-edge spalling is a common form of blade damage caused by hard objects, which affects the aerodynamic characteristics of the compressor. Airfoil profile is the foundation of compressor blade design and manufacturing; therefore, analyzing the impact of leading-edge spalling on compressor aerodynamic characteristics often begins with analyzing the impact of leading-edge spalling on the aerodynamic characteristics of the compressor airfoil profile. Existing analytical methods mainly employ numerical simulation to calculate the aerodynamic characteristics of the compressor airfoil profile under different operating conditions, comparing the losses and trailing angles of the compressor airfoil profile before and after leading-edge spalling to obtain the impact of leading-edge spalling on the aerodynamic characteristics of the compressor airfoil profile.

[0003] For a given compressor airfoil, when the incoming Mach number remains constant, the inlet angle of attack is the key factor causing aerodynamic changes in the airfoil. To obtain the impact of leading-edge debris on the aerodynamic characteristics of the compressor airfoil, it is necessary to calculate the aerodynamic characteristics of the airfoil at different angles of attack. Since the geometric characteristics of leading-edge debris are random, studying its impact on the aerodynamic characteristics of the compressor airfoil requires calculating a large number of samples and statistically determining the distribution law of the impact of leading-edge debris on the aerodynamic characteristics of the airfoil, which consumes significant computational resources. Therefore, exploring and developing a rapid prediction method for the impact of leading-edge debris on the aerodynamic characteristics of the compressor airfoil at different angles of attack is of great significance for assessing the impact of leading-edge debris on compressor aerodynamic characteristics. Summary of the Invention

[0004] To address the problem of compressor aerodynamic characteristics caused by blade leading-edge chipping due to hard object damage, this invention proposes a method for analyzing the angle-of-attack characteristics of compressor blades with leading-edge chipping, which specifically includes the following steps;

[0005] I. Constructing the leading-edge chipped compressor blade geometry;

[0006] Based on the geometric features of leading edge chipping caused by hard objects damaging the blades, the leading edge curve of the compressor blade is regenerated on the basis of the existing compressor blade geometry to construct the leading edge chipping compressor blade geometry.

[0007] II. Calculate the aerodynamic characteristics of the compressor blade profile under normal and limited angle of attack conditions after leading edge decapitation;

[0008] For a given compressor airfoil, determine its operating inlet Mach number and available angle of attack range; keeping the inlet Mach number constant, set the inlet angle and allow the compressor airfoil to operate at the designed inlet angle of attack, the minimum inlet angle of attack within the available inlet angle of attack range, and the maximum inlet angle of attack within the available inlet angle of attack range; calculate the losses and lag angle of the normal compressor airfoil through experiments and numerical simulations; for a compressor airfoil with leading-edge decapitation, keep the inlet angle constant and calculate the losses and lag angle of the compressor airfoil with leading-edge decapitation; compare the losses and lag angles of the normal compressor airfoil and the compressor airfoil with leading-edge decapitation to obtain the changes in the aerodynamic characteristics of the compressor airfoil after leading-edge decapitation at the designed inlet angle of attack, the minimum inlet angle of attack within the available inlet angle of attack range, and the maximum inlet angle of attack within the available inlet angle of attack range.

[0009] 3. Correlate the changes in the aerodynamic characteristics of the leading-edge slugging compressor airfoil under a limited angle of attack with the changes in the aerodynamic characteristics of the leading-edge slugging compressor airfoil under other inlet angles of attack;

[0010] For a compressor airfoil with leading-edge spalling, it is assumed that the airfoil loss and lag angle vary with the inlet angle of attack according to a quadratic function. Step two obtains the airfoil loss and lag angle of the compressor with leading-edge spalling at three inlet angles of attack. Using these three sets of data, the quadratic function relationship between the airfoil loss and lag angle of the compressor with leading-edge spalling and the inlet angle of attack is obtained. The difference between the quadratic function of the airfoil loss and lag angle of the compressor with leading-edge spalling and the quadratic function of the airfoil loss and lag angle of the normal compressor with the angle of attack is obtained to obtain the change of the aerodynamic characteristics of the compressor with leading-edge spalling relative to the normal compressor airfoil at different inlet angles of attack.

[0011] In one embodiment of the present invention, step one specifically includes the following steps:

[0012] Step 1: Determine the starting positions C1 and C6 of the pressure surface and suction surface when the leading edge blocks fall off;

[0013] Normally, the suction and pressure surface curves of a compressor blade profile converge at the leading edge point A. When a leading edge delamination occurs, the starting point of the suction surface curve changes from point A to point C1, and the starting point of the pressure surface curve changes from point A to point C6. The distance between point C1 and point A is denoted as DSm, and the distance between point C6 and point A is denoted as DPm. Given points C1 and C6, a cubic B-spline curve is used to determine the leading edge geometry of the delamination compressor blade profile. A total of 6 control points are needed to generate the leading edge geometry of the delamination compressor blade profile. The distance between points C1 and C6 is denoted as Rt.

[0014] Step 2: Generate control points C2 and C5 for the cubic B-spline curve based on points C1 and C6;

[0015] Point C2 is located outside the line connecting points C1 and C6, outside point C1. The angle between the line connecting points C2 and C1 and the tangent of the suction surface curve at point C1 is denoted as α1. For a given compressor blade profile, the direction of the line connecting points C1 and C2 is known, and the distance between points C1 and C2 is taken as k1 times Rt. Based on the coordinates of point C1, the coordinates of point C2 are obtained.

[0016] Point C5 is located outside the line connecting points C1 and C6, and outside point C6. The angle between the line connecting points C5 and C6 and the tangent of the pressure surface curve at point C6 is denoted as α2. For a given compressor blade profile, the direction of the line connecting points C5 and C6 is known, and the distance between points C5 and C6 is taken as k2 times Rt. Given the direction of the line connecting points C5 and C6 and the distance between them, the coordinates of point C5 are obtained based on the coordinates of point C6.

[0017] Step 3: Generate control points C3 and C4 for the cubic B-spline curve based on points C2 and C5;

[0018] Connecting points C2 and C5, we obtain a line segment with C2 as the starting point and C5 as the ending point. Divide this line segment into three equal parts, obtaining two points on it. One point is 1 / 3 of the line segment length from C2, denoted as C3b, and the other point is 2 / 3 of the line segment length from C2, denoted as C4b. Draw perpendicular lines from C3b and C4b to the line segment, and take points on these two perpendicular lines, namely C3 and C4. Points C3 and C4 extend in a direction parallel to the x-axis. To determine points C3 and C4, we need to know the distances d3 and d4 between C3 and C4 and C3b and C4b, respectively. Let d3 = k3 * Rt and d4 = k4 * Rt. Within the range of values ​​for k3 and k4, we use a random number generator to determine the values ​​of k3 and k4. When the value is negative, the point extends in the negative x-axis direction; when the value is positive, the point extends in the positive x-axis direction.

[0019] Step 4: Use points C1 to C6 to generate a cubic B-spline curve to obtain the leading edge curve of the scrambled compressor blade profile;

[0020] C1, C2, C3, C4, C5, and C6 are designated as six control points for the leading edge cubic B-spline curve. For a cubic B-spline curve using six control points, a node vector NV containing 10 elements needs to be constructed, denoted as NV = [n1, n2, n3, n4, n5, n6, n7, n8, n9, n10]. The values ​​of n5 and n6 are determined within the range of 0 to 1, ensuring that 0... <n5<n6<1;

[0021] This yields the leading edge curve of the chipped compressor blade profile. By splicing this curve with the suction and pressure surface curves, the geometry of the chipped compressor blade profile is obtained.

[0022] In another embodiment of the present invention, the values ​​of α1 and α2 are both in the range of 0° to 180°; k1 = 0.1 to 1, k2 = 0.1 to 1.

[0023] In a specific embodiment of the present invention, the ratios of DSm and DPm to the blade chord length are both no greater than 12%; α1 and α2 are both 0°; k1 = k2 = 0.2; the range of k3 and k4 is [-0.2, +0.2]; let n1 = n2 = n3 = n4 = 0, n7 = n8 = n9 = n10 = 1; let n5 = 1 / 3, n6 = 2 / 3.

[0024] In yet another embodiment of the present invention, step two specifically includes the following steps;

[0025] For a normal compressor airfoil, under a given incoming Mach number, the total pressure loss as a function of the inlet angle of attack i is obtained as curve ω1 = fω(i), and the lag angle as a function of the inlet angle of attack i is obtained as curve β1 = fβ(i), where fω(i) and fβ(i) represent the angle-of-attack characteristics of the total pressure loss and the angle-of-attack characteristics of the lag angle, respectively. When the aerodynamic characteristics of the normal compressor airfoil are known, the minimum angle of attack point P1, the design angle of attack point P2, and the maximum angle of attack point P3 of the airfoil can be obtained. The corresponding inlet angles of attack are i1, i2 and i3, respectively. For the leading edge chipped compressor blade profile constructed in step one, while keeping the incoming Mach number consistent with the normal compressor blade profile, the total pressure loss and lag angle of the leading edge chipped compressor blade profile are obtained by numerical simulation or experimental testing methods at inlet angles of attack of i1, i2 and i3, respectively. The total pressure loss at inlet angles of attack of i1, i2 and i3 is denoted as ω2P1, ω2P2 and ω2P3, respectively, and the lag angle is denoted as β2P1, β2P2 and β2P3, respectively.

[0026] In yet another specific embodiment of the present invention, step three specifically includes the following steps;

[0027] Step 1: Use quadratic function fitting to obtain the mathematical expressions for the total pressure loss and lag angle of the normal compressor blade as a function of the intake angle of attack;

[0028] For the curves ω1=fω(i) and β1=fβ(i) obtained in step two, which represent the changes in total pressure loss and lag angle of the normal compressor blade as a function of the inlet angle of attack, a quadratic function is used to fit the curves and obtain the mathematical expressions for fω(i) and fβ(i). When the incoming Mach number is constant, any inlet angle of attack i is substituted into the mathematical expressions for fω(i) and fβ(i) to obtain the total pressure loss and lag angle of the blade at that angle of attack.

[0029] Step 2: Use quadratic function fitting to obtain mathematical expressions for the total pressure loss and the lag angle of the compressor blade as a function of the angle of attack;

[0030] The total pressure loss and lag angle of the leading-edge chipped compressor airfoil satisfy a quadratic function relationship with the angle of attack. For the total pressure loss, using data from three points (i1, ω2P1), (i2, ω2P2), and (i3, ω2P3), the quadratic function relationship of the total pressure loss coefficient of the leading-edge chipped compressor airfoil with the inlet angle of attack is calculated: ω2 = fωb(i). For the lag angle, using data from three points (i1, β2P1), (i2, β2P2), and (i3, β2P3), the quadratic function relationship of the lag angle of the leading-edge chipped compressor airfoil with the inlet angle of attack is calculated: β2 = fβb(i). Substituting any inlet angle of attack i into the mathematical expressions of fωb(i) and fβb(i), the total pressure loss and lag angle of the leading-edge chipped compressor airfoil at that inlet angle of attack are obtained.

[0031] Step 3: Subtract the total pressure loss and the lag angle from the normal compressor blade profile and the leading edge chipped compressor blade profile to obtain the changes in total pressure loss and lag angle;

[0032] For any angle of attack i, obtain the total pressure loss and lag angle of the normal compressor blade profile under a given incoming Mach number, denoted as ω1(i) and β1(i), respectively, and the total pressure loss and lag angle of the compressor blade profile with leading edge deflection, denoted as ω2(i) and β2(i), respectively. Subtract the total pressure loss and lag angle of the normal compressor blade profile and the compressor blade profile with leading edge deflection from each other. For any angle of attack i, obtain the change in total pressure loss of the compressor blade profile Δω(i)=ω2(i)-ω1(i) and the change in lag angle Δβ(i)=β2(i)-β1(i) under the influence of leading edge deflection.

[0033] This invention enables rapid calculation of the aerodynamic characteristics of a compressor blade with leading-edge debris at different angles of attack, requiring only the calculation of blade profile losses and lag angles at three angles of attack. This invention simplifies the relationship between compressor blade aerodynamic characteristics and angle of attack into a quadratic function, allowing for rapid prediction of the impact of leading-edge debris on compressor blade aerodynamic characteristics. It has broad application prospects in assessing the aerodynamic impact of blade damage from hard objects on compressors. Attached Figure Description

[0034] Figure 1 It is the leading edge chipped compressor blade geometry;

[0035] Figure 2 It is an aerodynamic characteristic analysis model for compressor blades with leading edge drop. Detailed Implementation

[0036] The invention will now be further described with reference to the accompanying drawings.

[0037] This invention provides a method for analyzing the angle of attack characteristics of a compressor blade profile with leading edge spalling, which specifically includes the following steps.

[0038] I. Constructing the leading-edge chipped compressor blade geometry;

[0039] Based on the geometric characteristics of leading-edge chipping caused by hard objects damaging the blades, the leading-edge curve of the compressor blade profile is regenerated on the existing compressor blade geometry to construct the compressor blade geometry with leading-edge chipping. Specifically, the following steps are included:

[0040] Step 1: Determine the starting positions C1 and C6 of the pressure surface and suction surface when the leading edge blocks fall off;

[0041] like Figure 1 As shown, the suction and pressure surface curves of a normal compressor blade profile converge at the leading edge point A. When a leading edge debris falls off, the starting point of the suction surface curve changes from point A to point C1, and the starting point of the pressure surface curve changes from point A to point C6. Points C1 and C6 are determined based on statistical data of blade damage from hard objects during aero-engine service (this technique is known to those skilled in the art and will not be elaborated further). The distance between point C1 and point A is denoted as DSm, and the distance between point C6 and point A is denoted as DPm. Generally, the ratios of DSm and DPm to the blade chord length are no greater than 12%. This invention, given the known points C1 and C6, uses a cubic B-spline curve to determine the leading edge geometry of the compressor blade profile with debris falling off (a cubic B-spline curve is known to those skilled in the art and will not be elaborated further). Figure 1 As shown, generating the leading edge geometry of the sloping compressor blade requires 6 control points. Points C1 and C6 are known. Using geometric knowledge, the distance between points C1 and C6 can be obtained, denoted as Rt.

[0042] Step 2: Generate control points C2 and C5 for the cubic B-spline curve based on points C1 and C6;

[0043] For point C2, it is located outside the line connecting points C1 and C6, outside point C1. The angle between the line connecting points C2 and C1 and the tangent of the suction surface curve at point C1 is denoted as α1 (α1 ranges from 0° to 180°; in one embodiment of the invention, α1 is 0°). Since the direction of the tangent at any point on the suction surface is known for a given compressor blade profile, based on the above definition, the direction of the line connecting points C1 and C2 is known, and the distance between points C1 and C2 can be taken as k1 times Rt (k1 = 0.1 to 1); in one embodiment of the invention, k1 = 0.2. Given the direction of the line connecting points C1 and C2 and the distance between them, the coordinates of point C2 can be obtained using geometric knowledge based on the coordinates of point C1.

[0044] For point C5, it is located outside the line connecting points C1 and C6, and outside point C6. The angle between the line connecting points C5 and C6 and the tangent to the pressure surface curve at point C6 is denoted as α2 (α2 ranges from 0° to 180°; in one embodiment of the invention, α2 is 0°). Since the direction of the tangent at any point on the pressure surface is known for a given compressor blade profile, based on the above definition, the direction of the line connecting points C5 and C6 is known, and the distance between points C5 and C6 can be taken as k2 times Rt (k2 = 0.1 to 1); in one embodiment of the invention, k2 = 0.2. Given the direction of the line connecting points C5 and C6 and the distance between them, the coordinates of point C5 can be obtained using geometric knowledge based on the coordinates of point C6.

[0045] Step 3: Generate control points C3 and C4 for the cubic B-spline curve based on points C2 and C5;

[0046] After obtaining the coordinates of points C2 and C5, connect them to form a line segment. The line starts at point C2 and ends at point C5. Divide this line segment into three equal parts, obtaining two points on it. One point is 1 / 3 of the line segment length from point C2, denoted as C3b, and the other point is 2 / 3 of the line segment length from point C2, denoted as C4b. Draw perpendicular lines from points C3b and C4b to the line segment, and take points on each of these perpendicular lines, which are C3 and C4 respectively. Obviously, determining points C3 and C4 requires knowing the distances d3 and d4 between points C3 and C4 and points C3b and C4b, as well as the direction of extension of points C3 and C4 in the direction parallel to the x-axis. Let d3 = k3 * Rt and d4 = k4 * Rt, where the values ​​of k3 and k4 range from [-0.2, +0.2]. The specific values ​​are determined using a random number generation method: two numbers are randomly selected from the range [-0.2, +0.2] and assigned to k3 and k4. When the value is negative, the point extends in the negative x-axis direction; when the value is positive, the point extends in the positive x-axis direction. This ensures that the leading-edge cross-sectional shape of the compressor blade has a certain degree of randomness.

[0047] Step 4: Use points C1 to C6 to generate a cubic B-spline curve to obtain the leading edge curve of the scrambled compressor blade profile;

[0048] After determining the coordinates of C1, C2, C3, C4, C5 and C6, they are taken as 6 control points of the leading edge cubic B-spline curve. For a cubic B-spline curve, when 6 control points are used, a knot vector NV containing 10 elements needs to be constructed to generate the B-spline curve, denoted as NV=[n1,n2,n3,n4,n5,n6,n7,n8,n9,n10]. In one embodiment of the present invention, n1=n2=n3=n4=0 and n7=n8=n9=n10=1, which can ensure that the tangent directions of the B-spline curve at points C1 and C6 are consistent with the direction of the connecting line of C1 and C2 and the direction of the connecting line of C5 and C6 respectively. The values of n5 and n6 can be determined within the range of 0 to 1, ensuring 0<n5<n6<1. In one embodiment of the present invention, n5=1 / 3 and n6=2 / 3.

[0049] Through the above steps, the 6 control points and the knot vector NV of the cubic B-spline curve are determined. By utilizing the mathematical knowledge of cubic B-spline curves, the leading edge curve of the leading-edge-chipped compressor airfoil can be obtained, and by splicing the leading edge curve with the suction surface curve and the pressure surface curve, the geometry of the leading-edge-chipped compressor airfoil can be obtained.

[0050] II. Calculating the aerodynamic characteristics of the normal compressor airfoil and the compressor airfoil with leading edge chipping under limited incidence angles;

[0051] For a given compressor airfoil, the incoming flow Mach number for its operation and the available incidence angle range are determined. The incoming flow Mach number is kept unchanged, an inlet flow angle is set, and the compressor airfoil is operated at the design inlet incidence angle, the minimum inlet incidence angle within the available inlet incidence angle range, and the maximum inlet incidence angle within the available incidence angle range respectively. Through experiments and numerical simulation, the loss and deviation angle of the normal compressor airfoil are calculated (the method for calculating the loss and deviation angle of a compressor airfoil is known to those skilled in the art, and will not be repeated herein). For the compressor airfoil with leading edge chipping, with the inlet flow angle kept unchanged, the loss and deviation angle of the compressor airfoil with leading edge chipping are calculated. By comparing the loss and deviation angle of the normal compressor airfoil and the compressor airfoil with leading edge chipping, the changes of the aerodynamic characteristics of the compressor airfoil after leading edge chipping under the design inlet incidence angle, the minimum inlet incidence angle within the available inlet incidence angle range, and the maximum inlet incidence angle within the available inlet incidence angle range are obtained.

[0052] As Figure 2As shown, for a normal compressor blade profile, under a given incoming Mach number, the total pressure loss as a function of the inlet angle of attack i can be obtained as curve ω1=fω(i) using conventional numerical simulation or experimental testing methods, and the lag angle as a function of the inlet angle of attack i can be obtained as curve β1=fβ(i), where fω(i) and fβ(i) represent the angle-of-attack characteristics of the total pressure loss and the angle-of-attack characteristics of the lag angle of the normal blade profile, respectively. Given the known aerodynamic characteristics of a normal compressor blade profile, the minimum angle of attack point P1, the design angle of attack point P2, and the maximum angle of attack point P3 of the blade profile can be obtained, with corresponding inlet angles of attack i1, i2, and i3, respectively. For the leading-edge clipped compressor blade profile constructed in step one, while keeping the incoming Mach number consistent with that of the normal compressor blade profile, the total pressure loss and lag angle of the leading-edge clipped compressor blade profile are obtained using conventional numerical simulation or experimental testing methods at inlet angles of attack i1, i2, and i3, respectively. The total pressure loss at inlet angles of attack i1, i2, and i3 is denoted as ω2P1, ω2P2, and ω2P3, respectively, and the lag angle is denoted as β2P1, β2P2, and β2P3, respectively.

[0053] 3. Correlate the changes in the aerodynamic characteristics of the leading-edge slugging compressor airfoil under a limited angle of attack with the changes in the aerodynamic characteristics of the leading-edge slugging compressor airfoil under other inlet angles of attack;

[0054] For a given compressor blade, the incoming Mach number and available inlet angle of attack range for different blade profiles are generally known. Given a fixed incoming Mach number, the relationship between blade profile loss and lag angle with the inlet angle of attack is also generally known (usually a quadratic function). This invention proposes using a quadratic function curve to fit the relationship between blade profile loss and lag angle with the inlet angle of attack. For a compressor blade with leading-edge spalling, it is similarly assumed that the relationship between blade profile loss and lag angle with the inlet angle of attack also conforms to a quadratic function law. In step two, the blade profile loss and lag angle of the compressor with leading-edge spalling are obtained at three inlet angles of attack. Using these three sets of data, a quadratic function expression for the relationship between leading-edge spalling blade profile loss and lag angle with the inlet angle of attack can be obtained. The method for obtaining this quadratic function expression is known to those skilled in the art and will not be elaborated further. By subtracting the quadratic functions of the leading-edge chipped compressor airfoil loss and lag angle as a function of the inlet angle of attack from the quadratic functions of the normal compressor airfoil loss and lag angle as a function of the angle of attack, the aerodynamic characteristics of the leading-edge chipped compressor airfoil relative to the normal compressor airfoil can be obtained at different inlet angles of attack. Specifically, this includes the following steps:

[0055] Step 1: Use quadratic function fitting to obtain the mathematical expressions for the total pressure loss and lag angle of the normal compressor blade as a function of the intake angle of attack;

[0056] For the curves ω1=fω(i) and β1=fβ(i) obtained in step two, which represent the changes in total pressure loss and lag angle of the normal compressor blade with the inlet angle of attack, this invention uses a quadratic function for fitting to obtain the mathematical expressions for fω(i) and fβ(i) (this mathematical method is well known to those skilled in the art). Therefore, when the incoming Mach number is constant, by substituting any inlet angle of attack i into the mathematical expressions for fω(i) and fβ(i), the total pressure loss and lag angle of the blade at that angle of attack can be obtained.

[0057] Step 2: Use quadratic function fitting to obtain mathematical expressions for the total pressure loss and the lag angle of the compressor blade as a function of the angle of attack;

[0058] Step two yielded the total pressure loss and lag angles of the leading-edge scraggly compressor airfoil at the minimum angle of attack point P1, the design angle of attack point P2, and the maximum angle of attack point P3: ω2P1, ω2P2, ω2P3 and β2P1, β2P2, β2P3. The changes in total pressure loss and lag angle of the leading-edge scraggly compressor airfoil with the angle of attack also satisfy a quadratic function relationship. For the total pressure loss, using the data from the three points (i1, ω2P1), (i2, ω2P2), and (i3, ω2P3), the quadratic function relationship of the total pressure loss coefficient of the leading-edge scraggly compressor airfoil with the inlet angle of attack can be calculated: ω2 = fωb(i); for the lag angle, using the data from the three points (i1, β2P1), (i2, β2P2), and (i3, β2P3), the quadratic function relationship of the lag angle of the leading-edge scraggly compressor airfoil with the inlet angle of attack can be calculated: β2 = fβb(i). After obtaining the expressions for fωb(i) and fβb(i), by substituting any intake angle of attack i into the mathematical expressions for fωb(i) and fβb(i), the total pressure loss and lag angle of the compressor blade at that intake angle can be obtained.

[0059] Step 3: Subtract the total pressure loss and the lag angle from the normal compressor blade profile and the leading edge chipped compressor blade profile to obtain the changes in total pressure loss and lag angle;

[0060] For any angle of attack i, the method of this invention can be used to obtain the total pressure loss and lag angle of the normal compressor blade profile at a given incoming Mach number, denoted as ω1(i) and β1(i), respectively, and the total pressure loss and lag angle of the compressor blade profile with leading-edge deflection, denoted as ω2(i) and β2(i), respectively. By subtracting the total pressure loss and lag angle of the normal compressor blade profile and the compressor blade profile with leading-edge deflection, for any angle of attack i, the changes in the total pressure loss of the compressor blade profile under the influence of leading-edge deflection can be obtained as Δω(i) = ω2(i) - ω1(i) and the changes in the lag angle as Δβ(i) = β2(i) - β1(i).

Claims

1. A method for analyzing the angle-of-attack characteristics of a compressor blade profile with leading-edge spalling, characterized in that, Specifically, it includes the following steps; I. Constructing the leading-edge chipped compressor blade geometry; Based on the geometric features of leading edge chipping caused by hard objects damaging the blades, the leading edge curve of the compressor blade is regenerated on the basis of the existing compressor blade geometry to construct the leading edge chipping compressor blade geometry. II. Calculate the aerodynamic characteristics of the compressor blade profile under normal and limited angle of attack conditions after leading edge decapitation; For a given compressor airfoil, determine its operating inlet Mach number and available angle of attack range; keeping the inlet Mach number constant, set the inlet angle and allow the compressor airfoil to operate at the designed inlet angle of attack, the minimum inlet angle of attack within the available inlet angle of attack range, and the maximum inlet angle of attack within the available inlet angle of attack range; calculate the losses and lag angle of the normal compressor airfoil through experiments and numerical simulations; for a compressor airfoil with leading-edge decapitation, keep the inlet angle constant and calculate the losses and lag angle of the compressor airfoil with leading-edge decapitation; compare the losses and lag angles of the normal compressor airfoil and the compressor airfoil with leading-edge decapitation to obtain the changes in the aerodynamic characteristics of the compressor airfoil after leading-edge decapitation at the designed inlet angle of attack, the minimum inlet angle of attack within the available inlet angle of attack range, and the maximum inlet angle of attack within the available inlet angle of attack range.

3. Correlate the changes in the aerodynamic characteristics of the leading-edge slugging compressor airfoil under a limited angle of attack with the changes in the aerodynamic characteristics of the leading-edge slugging compressor airfoil under other inlet angles of attack; For a compressor airfoil with leading-edge spalling, it is assumed that the airfoil loss and lag angle vary with the inlet angle of attack according to a quadratic function. Step two obtains the airfoil loss and lag angle of the compressor with leading-edge spalling at three inlet angles of attack. Using these three sets of data, the quadratic function relationship between the airfoil loss and lag angle of the compressor with leading-edge spalling and the inlet angle of attack is obtained. The difference between the quadratic function of the airfoil loss and lag angle of the compressor with leading-edge spalling and the quadratic function of the airfoil loss and lag angle of the normal compressor with the angle of attack is obtained to obtain the change of the aerodynamic characteristics of the compressor with leading-edge spalling relative to the normal compressor airfoil at different inlet angles of attack.

2. The method for analyzing the angle of attack characteristics of a compressor blade profile with leading-edge spalling as described in claim 1, characterized in that, Step one specifically includes the following steps; Step 1: Determine the starting positions C1 and C6 of the pressure surface and suction surface when the leading edge blocks fall off; Normally, the suction and pressure surface curves of a compressor blade profile converge at the leading edge point A. When a leading edge delamination occurs, the starting point of the suction surface curve changes from point A to point C1, and the starting point of the pressure surface curve changes from point A to point C6. The distance between point C1 and point A is denoted as DSm, and the distance between point C6 and point A is denoted as DPm. Given points C1 and C6, a cubic B-spline curve is used to determine the leading edge geometry of the delamination compressor blade profile. A total of 6 control points are needed to generate the leading edge geometry of the delamination compressor blade profile. The distance between points C1 and C6 is denoted as Rt. Step 2: Generate control points C2 and C5 for the cubic B-spline curve based on points C1 and C6; Point C2 is located outside the line connecting points C1 and C6, outside point C1. The angle between the line connecting points C2 and C1 and the tangent of the suction surface curve at point C1 is denoted as α1. For a given compressor blade profile, the direction of the line connecting points C1 and C2 is known, and the distance between points C1 and C2 is taken as k1 times Rt. Based on the coordinates of point C1, the coordinates of point C2 are obtained. Point C5 is located outside the line connecting points C1 and C6, and outside point C6. The angle between the line connecting points C5 and C6 and the tangent of the pressure surface curve at point C6 is denoted as α2. For a given compressor blade profile, the direction of the line connecting points C5 and C6 is known, and the distance between points C5 and C6 is taken as k2 times Rt. Given the direction of the line connecting points C5 and C6 and the distance between them, the coordinates of point C5 are obtained based on the coordinates of point C6. Step 3: Generate control points C3 and C4 for the cubic B-spline curve based on points C2 and C5; Connecting points C2 and C5, we obtain a line segment with C2 as the starting point and C5 as the ending point. Divide this line segment into three equal parts, obtaining two points on it. One point is 1 / 3 of the line segment length from C2, denoted as C3b, and the other point is 2 / 3 of the line segment length from C2, denoted as C4b. Draw perpendicular lines from C3b and C4b to the line segment, and take points on these two perpendicular lines, namely C3 and C4. Points C3 and C4 extend in a direction parallel to the x-axis. To determine points C3 and C4, we need to know the distances d3 and d4 between C3 and C4 and C3b and C4b, respectively. Let d3 = k3 * Rt and d4 = k4 * Rt. Within the range of values ​​for k3 and k4, we use a random number generator to determine the values ​​of k3 and k4. When the value is negative, the point extends in the negative x-axis direction; when the value is positive, the point extends in the positive x-axis direction. Step 4: Use points C1 to C6 to generate a cubic B-spline curve to obtain the leading edge curve of the scrambled compressor blade profile; C1, C2, C3, C4, C5, and C6 are designated as six control points for the leading edge cubic B-spline curve. For a cubic B-spline curve using six control points, a node vector NV containing 10 elements needs to be constructed, denoted as NV = [n1, n2, n3, n4, n5, n6, n7, n8, n9, n10]. The values ​​of n5 and n6 are determined within the range of 0 to 1, ensuring that 0... <n5<n6<1; This yields the leading edge curve of the chipped compressor blade profile. By splicing this curve with the suction and pressure surface curves, the geometry of the chipped compressor blade profile is obtained.

3. The method for analyzing the angle of attack characteristics of a compressor blade profile with leading-edge spalling as described in claim 2, characterized in that, The values ​​of α1 and α2 are both in the range of 0° to 180°; k1 = 0.1 to 1, k2 = 0.1 to 1.

4. The method for analyzing the angle of attack characteristics of a compressor blade profile with leading-edge spalling as described in claim 3, characterized in that, The ratios of DSm and DPm to the blade chord length are both no greater than 12%; α1 and α2 are both 0°; k1 = k2 = 0.2; the range of k3 and k4 is [-0.2, +0.2]; let n1 = n2 = n3 = n4 = 0, n7 = n8 = n9 = n10 = 1; let n5 = 1 / 3, n6 = 2 / 3.

5. The method for analyzing the angle of attack characteristics of a compressor blade profile with leading-edge spalling as described in claim 1, characterized in that, Step two specifically includes the following steps; For a normal compressor airfoil, under a given incoming Mach number, the total pressure loss as a function of the inlet angle of attack i is obtained as curve ω1 = fω(i), and the lag angle as a function of the inlet angle of attack i is obtained as curve β1 = fβ(i), where fω(i) and fβ(i) represent the angle-of-attack characteristics of the total pressure loss and the angle-of-attack characteristics of the lag angle, respectively. When the aerodynamic characteristics of the normal compressor airfoil are known, the minimum angle of attack point P1, the design angle of attack point P2, and the maximum angle of attack point P3 of the airfoil can be obtained. The corresponding inlet angles of attack are i1, i2 and i3, respectively. For the leading edge chipped compressor blade profile constructed in step one, while keeping the incoming Mach number consistent with the normal compressor blade profile, the total pressure loss and lag angle of the leading edge chipped compressor blade profile are obtained by numerical simulation or experimental testing methods at inlet angles of attack of i1, i2 and i3, respectively. The total pressure loss at inlet angles of attack of i1, i2 and i3 is denoted as ω2P1, ω2P2 and ω2P3, respectively, and the lag angle is denoted as β2P1, β2P2 and β2P3, respectively.

6. The method for analyzing the angle of attack characteristics of a compressor blade profile with leading-edge spalling as described in claim 5, characterized in that, Step three specifically includes the following steps; Step 1: Use quadratic function fitting to obtain the mathematical expressions for the total pressure loss and lag angle of the normal compressor blade as a function of the intake angle of attack; For the curves ω1=fω(i) and β1=fβ(i) obtained in step two, which represent the changes in total pressure loss and lag angle of the normal compressor blade as a function of the inlet angle of attack, a quadratic function is used to fit the curves and obtain the mathematical expressions for fω(i) and fβ(i). When the incoming Mach number is constant, any inlet angle of attack i is substituted into the mathematical expressions for fω(i) and fβ(i) to obtain the total pressure loss and lag angle of the blade at that angle of attack. Step 2: Use quadratic function fitting to obtain mathematical expressions for the total pressure loss and the lag angle of the compressor blade as a function of the angle of attack; The total pressure loss and lag angle of the leading-edge chipped compressor airfoil satisfy a quadratic function relationship with the angle of attack. For the total pressure loss, using data from three points (i1, ω2P1), (i2, ω2P2), and (i3, ω2P3), the quadratic function relationship of the total pressure loss coefficient of the leading-edge chipped compressor airfoil with the inlet angle of attack is calculated: ω2 = fωb(i). For the lag angle, using data from three points (i1, β2P1), (i2, β2P2), and (i3, β2P3), the quadratic function relationship of the lag angle of the leading-edge chipped compressor airfoil with the inlet angle of attack is calculated: β2 = fβb(i). Substituting any inlet angle of attack i into the mathematical expressions of fωb(i) and fβb(i), the total pressure loss and lag angle of the leading-edge chipped compressor airfoil at that inlet angle of attack are obtained. Step 3: Subtract the total pressure loss and the lag angle from the normal compressor blade profile and the leading edge chipped compressor blade profile to obtain the changes in total pressure loss and lag angle; For any angle of attack i, obtain the total pressure loss and lag angle of the normal compressor blade profile under a given incoming Mach number, denoted as ω1(i) and β1(i), respectively, and the total pressure loss and lag angle of the compressor blade profile with leading edge deflection, denoted as ω2(i) and β2(i), respectively. Subtract the total pressure loss and lag angle of the normal compressor blade profile and the compressor blade profile with leading edge deflection from each other. For any angle of attack i, obtain the change in total pressure loss of the compressor blade profile Δω(i)=ω2(i)-ω1(i) and the change in lag angle Δβ(i)=β2(i)-β1(i) under the influence of leading edge deflection.

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