A method for modifying the double-row blades of a turbine compressor in series

Abstract

The invention discloses a method for serially modifying double-row blades of a turbine compressor. In the method, i flow surfaces are respectively taken along the blade height of the double-row blades of the turbine compressor to obtain geometric coordinates of airfoils on each flow surface; based on the geometric coordinates of the airfoils, the airfoils on the i-th flow surface of the first row of blades and the second row of blades are respectively translated, rotated and dimensionless so that the leading edge points of the respective airfoils are respectively coincident with the origin of the coordinates, and the mid-arc line and thickness distribution of the airfoils on the i-th flow surface are respectively measured; the leading edge points of the airfoils of the first row of blades in the serial arrangement on the i-th flow surface are coincident with the leading edge points of the airfoils of the first row of blades in the double row, and after the positions of the second row of blades in the serial arrangement are determined by calculation, the parameterized geometry of the serial airfoils in the cylindrical coordinate system is obtained; the serial airfoils are sequentially stacked according to the flow surfaces to realize the modification of the double-row blades into the serial blades, the stacking points are all the leading edge points of the airfoils of the first row of blades in the double row of blades on the flow surfaces taken at different blade heights, and the stacking line is the leading edge line of the first row of blades in the double row of blades.

Description

Technical Field

[0001] The invention belongs to the field of turbine machinery pneumatic technology, in particular to a method for serially modifying double-row blades of a turbine compressor. Background Art

[0002] The turbine compressor is one of the core components of aircraft engines, gas turbines and industrial compression systems. Its outlet stator and outlet guide vane have the "heavy task" of airflow deflection and pressure increase: the outlet stator converts the kinetic energy obtained by the airflow from the rotor into pressure energy and allows the airflow to enter the next stage at a suitable angle. The outlet guide vane further converts the remaining kinetic energy into pressure energy to increase the pressure increase of the entire outlet stage. The outlet compressor blades have the characteristics of low aspect ratio, large airflow turning angle, and narrow stable operating range, which seriously affect the performance of the whole machine. At the same time, the axial length of the outlet stage has always been restricted by traditional aerodynamic design methods. How to achieve the shortest possible outlet stage wheelbase while ensuring the outlet stage efficiency and stable working margin is of great significance for further improving the rotor system strength, reducing the shaft system deflection, reducing weight, and saving processing and manufacturing costs of aircraft engines / gas turbines. In view of the characteristics of low aspect ratio, large airflow turning angle and narrow stable operating range of turbine compressor outlet stage blades, researchers have taken a variety of measures, including: changing blade installation angle, compound bend / swept design, casing treatment, non-uniform outlet guide vane design, adsorption blade design, etc. These measures can improve the efficiency and stable working margin of turbine compressors, but fail to take into account the compactness of the structure at the same time.

[0003] As an advanced flow control technology, tandem blades are usually composed of two rows of blades that are staggered circumferentially and separated or overlapped axially. The working principle of tandem blades is to reasonably distribute the aerodynamic load to the front / rear rows of blades, thereby forming a gap flow channel between the two rows of blades. The fluid flowing through the pressure surface of the front row of blades is guided to the suction surface of the rear row of blades through the gap area, blowing away and suppressing the flow separation on the suction surface of the rear row of blades, thereby achieving higher airflow deflection. Therefore, tandem blades have the advantages of large airflow turning angles and a wide range of low-loss stable operating conditions. At the same time, compared with conventional stator-outlet guide vanes, tandem blades can significantly shorten the axial installation dimensions and achieve a compact structure. So far, there has been no precedent for the modification design of tandem blades to replace double-row blades. Therefore, it is very necessary to supplement and develop new aerodynamic layout methods for turbine compressor outlet stages.

[0004] The above information disclosed in this Background section is only for enhancement of understanding of the background of the invention and therefore it may contain information that does not form the prior art that is already known in this country to a person of ordinary skill in the art. Summary of the invention

[0005] In view of the problems existing in the prior art, the present invention proposes a method for modifying the double-row blades of a turbine compressor in series. The object of the present invention is achieved through the following technical scheme. The method for modifying the double-row blades of a turbine compressor in series includes:

[0006] Step 1: Take i flow surfaces along the blade height of the first row of double-row blades of the turbine compressor and obtain the geometric coordinates of the airfoil on each flow surface;

[0007] Step 2: Based on the airfoil geometric coordinates, the airfoil on the i-th flow surface is translated, rotated and dimensionless so that the leading edge point of the airfoil coincides with the origin of the coordinates, and the mid-camber line and thickness distribution of the airfoil on the i-th flow surface are measured;

[0008] Step 3: parameterize the camber line and thickness distribution respectively using NURBS curves, parameterize the thickness distribution respectively using NURBS curves, superimpose the parameterized camber line and thickness distribution to generate an airfoil profile, parameterize the leading edge and trailing edge of the airfoil respectively using NURBS curves, and modify the leading edge and trailing edge respectively to complete the airfoil parameterization of the i-th flow surface of the first row of blades;

[0009] Step 4: Take i flow surfaces along the blade height of the second row of blades in the double row of blades, obtain the airfoil geometric coordinates on each flow surface, and repeat steps 2 to 3 to complete the airfoil parameterization of the i-th flow surface of the second row of blades;

[0010] Step 5: Determine the following parameters based on the assembly dimensions: effective chord length C i，eff is the airfoil chord length C of the first row of blades in the double row of blades i，blade1 With k i times the airfoil chord length C of the second row of blades i，blade2 The sum of C i，eff =C i，blade1 +k i ·C i，blade2 , the axial overlap of the series AO i The length Δm in the m direction of the overlapping part of the first row of blade airfoils and the second row of blade airfoils in the series i，1 The m-direction distance Δm between the leading edge of the first row of blades and the trailing edge of the second row of blades i，2 Ratio, AO i =Δm i，1 / Δm i，2 The m direction is the direction along the meridian streamline on the flow surface, and the axial overlap of the series AO i Determine the position relationship between the trailing edge point of the first row of airfoils and the leading edge point of the second row of airfoils in the m direction, and the chord length ratio CR of the series i The value is the ratio of the chord length of the airfoil of the second row of blades in series to the chord length of the airfoil of the first row of blades, CR i =C i，AA / Ci，FA ;

[0011] Step 6: Determine the trailing edge position of the airfoil of the first row of blades in series and the trailing edge position of the airfoil of the second row of blades in series, based on the effective chord length C i，eff Determine the effective consistency value σ of the tandem blades eff , C i，AA =C i，FA CR i ;

[0012] Step 7: Determine the tandem blade pitch s, the pitch s being the circumferential distance between the leading edge of the airfoil of the first row of blades in the tandem and the leading edge of the airfoil of the first row of blades in the adjacent blade channel, and determine the number of blades z based on the pitch s, z=2πR / s, if the number of blades z is an integer, continue; otherwise, return to step 6 until the number of blades z is an integer;

[0013] Step 8: Determine the circumferential pitch ratio PP of the position relationship between the trailing edge points of the first row of tandem airfoils and the leading edge points of the second row of tandem airfoils along the u direction i t is the circumferential distance between the trailing edge of the first row of blades in the series and the leading edge of the second row of blades in the adjacent series i The ratio of PP to the grating pitch s is t i / s, the u direction is a direction quasi-orthogonal to the m direction on the flow surface, based on the circumferential pitch ratio PP i Obtaining the coordinates of the airfoil of the first row of blades in series and the airfoil of the second row of blades in series;

[0014] Step 9: The coordinates of the airfoil of the first row of blades in the series and the airfoil of the second row of blades in the series are multiplied by the corresponding airfoil chord length, and are rotated and translated so that the leading edge point of the airfoil of the first row of blades in the series coincides with the leading edge point of the airfoil on the i-th flow surface of the first row of blades in the double row of blades, and the flow surface coordinates of the series airfoil are converted into cylindrical coordinates to obtain the parameterized geometry of the series airfoil in the cylindrical coordinate system;

[0015] Step 10: Repeat steps 2 to 10 for the airfoil of the i-th flow surface to obtain the parameterized geometric cylindrical coordinates of the airfoils in series on all flow surfaces; the serial airfoils are stacked in sequence according to the flow surfaces to realize the transformation of the double-row blades into serial blades, the stacking points are all the leading edge points of the airfoils on the flow surfaces taken at different blade heights for the first row of blades in the double-row blades, and the stacking line is the leading edge line of the first row of blades in the double-row blades.

[0016] In the serial modification method of double-row blades of a turbine compressor, in step 1, among the i flow surfaces, i=1 is the flow surface at 0% blade height, i=2 is the flow surface at 25% blade height, i=3 is the flow surface at 50% blade height, i=4 is the flow surface at 75% blade height, and i=5 is the flow surface at 100% blade height.

[0017] In the tandem modification method of double-row blades of a turbine compressor, in step 3, two-segment and 3rd-order NURBS curves are used to parameterize the median camber line and thickness distribution respectively, two-segment and 3rd-order NURBS curves are used to parameterize the thickness distribution, and two-segment and 3rd-order NURBS curves are used to parameterize the leading edge and trailing edge of the airfoil respectively to modify the leading edge and trailing edge of the blade body generated by superimposing the parameterized median camber line and the parameterized thickness distribution to complete the airfoil parameterization of the i-th flow surface of the first row of blades.

[0018] In the method for modifying the double-row blades of a turbine compressor in series, the k i is 1.2 to 1.8, the AO i -0.1~0.1, CR i It is 0.5~2.0.

[0019] In the method for modifying the double-row blades of a turbine compressor in series, the effective chord length C i，eff and axial overlap AO i is the flow surface at 50% blade height, i is 3, and the grating distance s is, z = 2πR / s, R is the radius of the stream surface at 50% of the blade height.

[0020] In the method for modifying the double-row blades of a turbine compressor in series, PP i It is 0.85~0.95.

[0021] In the method for serial modification of double-row blades of a turbine compressor, in step 10, five flow surfaces are stacked along the blade height in the order of i=1 to i=5, the stacking points are all the leading edge points of the airfoil of the first row of blades in the double-row blades on the flow surface at different blade heights, and the stacking line is the leading edge line of the first row of blades in the double-row blades. If there is no geometric interference, the modification of replacing the double-row blades with serial blades is completed; otherwise, return to step 5.

[0022] In the described method for serially modifying double-row blades of a turbine compressor, the double-row blades of the turbine compressor are stator blades.

[0023] In the method for serially modifying double-row blades of a turbine compressor, the double-row blades have two rows with equal blade numbers.

[0024] In the method for serially modifying double-row blades of a turbine compressor, the double-row blades have two rows with unequal numbers of blades.

[0025] Compared with the prior art, the present invention has the following advantages: the method for modifying the tandem arrangement of double-row blades of a turbine compressor described in the present invention is applied to a modified tandem arrangement of double-row stators with equal or unequal blade numbers, which can not only improve the aerodynamic performance of the turbine compressor outlet stage, but also further shorten the outlet stage wheelbase. At the same time, the geometry and configuration parameters of the modified tandem blades can be optimized and adjusted. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] By reading the detailed description of the preferred specific embodiments below, various other advantages and benefits of the present invention will become clear to those of ordinary skill in the art. The drawings in the specification are only for the purpose of illustrating the preferred embodiments and are not considered to be limitations of the present invention. Obviously, the drawings described below are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can also be obtained based on these drawings without creative work. Moreover, the same reference numerals are used to represent the same components throughout the drawings.

[0027] In the attached picture:

[0028] Figure 1 is a schematic diagram of parameterization of the camber line of the i-th flow surface airfoil of a tandem modification method for double-row blades of a turbine compressor according to an embodiment of the present invention;

[0029] Figure 2 is a schematic diagram of parameterization of the camber line of the i-th flow surface airfoil of a tandem modification method for double-row blades of a turbine compressor according to an embodiment of the present invention;

[0030] Figure 3 is a schematic diagram of parameterization of the leading edge of the i-th flow surface airfoil of a tandem modification method for double-row blades of a turbine compressor according to an embodiment of the present invention;

[0031] Figure 4 is a schematic diagram of parameterization of the trailing edge of the i-th flow surface airfoil of a tandem modification method for double-row blades of a turbine compressor according to an embodiment of the present invention;

[0032] Figure 5 1 is a schematic flow chart of a method for modifying a double-row blade of a turbine compressor in series according to an embodiment of the present invention;

[0033] Figure 6 It is a schematic diagram of geometric comparison before and after modification of a tandem modification method of double-row blades of a turbine compressor according to an embodiment of the present invention.

[0034] The present invention is further explained below in conjunction with the accompanying drawings and embodiments. DETAILED DESCRIPTION

[0035] The following will refer to the attached Figures 1 to 6Specific embodiments of the present invention are described in more detail. Although specific embodiments of the present invention are shown in the accompanying drawings, it should be understood that the present invention can be implemented in various forms and should not be limited by the embodiments described herein. On the contrary, these embodiments are provided to enable a more thorough understanding of the present invention and to fully convey the scope of the present invention to those skilled in the art.

[0036] It should be noted that certain words are used in the specification and claims to refer to specific components. Those skilled in the art should understand that technicians may use different nouns to refer to the same component. This specification and claims do not use the difference in nouns as a way to distinguish components, but use the difference in the functions of the components as the criterion for distinction. As mentioned throughout the specification and claims, "including" or "comprising" is an open term, so it should be interpreted as "including but not limited to". The subsequent description of the specification is a preferred embodiment of the present invention, but the description is based on the general principles of the specification and is not intended to limit the scope of the present invention. The scope of protection of the present invention shall be determined by the attached claims.

[0037] To facilitate understanding of the embodiments of the present invention, further explanation will be given below by taking specific embodiments as examples in conjunction with the accompanying drawings, and each of the accompanying drawings does not constitute a limitation on the embodiments of the present invention.

[0038] For a better understanding, Figures 1 to 6 As shown, the method for modifying the double-row blades of a turbine compressor in series includes:

[0039] Step 1: For the first row of blades in the double-row blades, i flow surfaces are taken along the blade height, i=1 is the flow surface at 0% blade height, i=2 is the flow surface at 25% blade height, i=3 is the flow surface at 50% blade height, i=4 is the flow surface at 75% blade height, and i=5 is the flow surface at 100% blade height. The geometric coordinates of the airfoil on each flow surface are obtained.

[0040] Step 2, translate, rotate and dimensionlessly transform the airfoil on the i-th flow surface described in step 1 so that the leading edge point of the airfoil coincides with the coordinate origin (0, 0), and measure the mid-arc line and thickness distribution of the airfoil on the i-th flow surface.

[0041] Step 3, parameterizing the mid-camber line of the airfoil described in step 2 using two-segment, third-order NURBS curves.

[0042] Step 4, parameterize the thickness distribution of the airfoil described in step 2 using a two-segment, third-order NURBS curve.

[0043] Step 5, using the mid-arc line described in step 3 and superimposing the thickness distribution described in step 4 to generate an airfoil profile line.

[0044] Step 6, parameterize the leading edge of the airfoil described in step 2 using two-segment, third-order NURBS curves.

[0045] Step 7, parameterize the trailing edge of the airfoil described in step 2 using a two-segment, third-order NURBS curve.

[0046] Step 8, using step 6 to modify the leading edge of the airfoil described in step 5, using step 7 to modify the trailing edge of the airfoil described in step 5, and completing the airfoil parameterization of the i-th flow surface of the first row of blades described in step 1.

[0047] Step 9: Take the airfoil on the i-th flow surface described in step 1 for the second row of blades in the double row of blades, and repeat steps 2 to 8. Complete the parameterization of the complete airfoil on the i-th flow surface described in step 1 for the second row of blades.

[0048] Step 10: Give the effective chord length C of the tandem airfoil according to the actual assembly dimensions i，eff The effective chord length is defined as the line connecting the leading edge point of the first row of airfoils and the trailing edge point of the second row of airfoils. i，eff Set to the airfoil chord length C of the first row of blades in the double row of blades described in step 1 i，blade1 With k i times the airfoil chord length C of the second row of blades i，blade2 The sum of the second row of leaves is defined as the second row of leaves in step 9: C i，eff =C i，blade1 +k i ·C i，blade2 . k i =1.2~1.8.

[0049] Step 11: Given the axial overlap AO of the serial configuration parameter i The axial overlap is defined as the m-direction length Δm of the overlapping portion of the first row of airfoils and the second row of airfoils in the series. i，1 The m-direction distance Δm between the leading edge point of the first row of airfoils and the trailing edge point of the second row of airfoils i，2 The ratio of AO i =Δm i，1 / Δm i，2 , the m direction is the direction along the meridian streamline on the flow surface. i Take -0.1 to 0.1 to determine the position relationship between the trailing edge points of the first row of airfoils and the leading edge points of the second row of airfoils in the m direction.

[0050] Step 12: Given the string configuration parameter chord length ratio CR i value, the CR i The value is defined as the ratio of the chord length of the second row of airfoils in the series to the chord length of the first row of airfoils, that is, CR i =C i，AA / Ci，FA CR i Take 0.5~2.0.

[0051] Step 13, solve the chord length C of the first row of tandem blades by the following formula: i，FA , determine the position of the trailing edge point of the first row of airfoils in series,

[0052]

[0053] Step 14, solve the chord length C of the second row of tandem blades by the following formula: i，AA , determine the position of the trailing edge point of the second row of tandem airfoils,

[0054] C i，AA =C i，FA CR i .

[0055] Step 15: Give the effective consistency value σ of the tandem blades eff , the effective consistency value is determined by the effective chord length described in step 10.

[0056] Step 16, determine the tandem blade pitch s by the following formula, where the pitch is defined as the circumferential distance between the leading edge point of the first row of the tandem airfoil and the leading edge point of the first row of the tandem airfoil of the adjacent blade path. Wherein, the effective chord length C in step 10 is i，eff and the axial overlap AO described in step 11 i Take the flow surface at 50% blade height described in step 1, i.e. i=3,

[0057]

[0058] Step 17, solve the blade number z by the following formula, where R is the radius of the flow surface at 50% blade height,

[0059] z=2πR / s.

[0060] Step 18: If the number of blades is an integer, continue to step 19; otherwise, return to step 15.

[0061] Step 19, given the circumferential pitch ratio PP i The circumferential pitch ratio is defined as the circumferential distance t between the trailing edge of the first row of tandem airfoils and the leading edge of the second row of tandem blades in the adjacent blade path. i and the ratio of the grating pitch s described in step 16, that is, PP = t i / s.PP i Take 0.85 to 0.95. This determines the positional relationship between the trailing edge points of the first row of airfoils and the leading edge points of the second row of airfoils in the u direction.

[0062] Step 20, obtaining the coordinates of the first row of airfoils in series and the second row of airfoils in series.

[0063] Step 21, multiplying the coordinates of the series airfoils on the i-th flow surface described in step 20 by the chord length of the airfoils, and rotating and translating them so that the leading edge points of the airfoils in the first row of the series coincide with the leading edge points of the airfoils on the i-th flow surface described in step 1.

[0064] Step 22, converting the tandem airfoil stream surface coordinates into cylindrical coordinates to obtain the tandem airfoil parameterized geometry in the cylindrical coordinate system.

[0065] Step 23, repeating steps 2 to 22 for the airfoil of the i-th flow surface described in step 1. If i=5, then step 24; otherwise, return to step 2. Obtain the parametric geometric cylindrical coordinates of the airfoils on all the flow surfaces described in step 1.

[0066] Step 24, stacking the tandem airfoil described in step 23 along the blade height in the order of i=1 to i=5 according to the five flow surfaces described in step 1, the stacking points are all the leading edge points of the airfoil of the first row of blades in the double-row blades described in step 1 on the flow surfaces taken at different blade heights, and the stacking line is the leading edge line of the first row of blades in the double-row blades described in step 1. The double-row blade is converted into a tandem blade.

[0067] Step 25, if there is no geometric interference after parametric modeling, the modified design of replacing double-row blades with tandem blades is completed; otherwise, return to step 10.

[0068] In one embodiment, the method includes,

[0069] Step 1, for the first row of blades in the double row of blades, take i flow surfaces along the blade height, i=1, i.e., the flow surface at 0% of the blade height, i=2, i.e., the flow surface at 25% of the blade height, i=3, i.e., the flow surface at 50% of the blade height, i=4, i.e., the flow surface at 75% of the blade height, and i=5, i.e., the flow surface at 100% of the blade height.

[0070] Step 2: The airfoil of the i-th flow surface described in step 1, the airfoil line coordinate point is (m i ,u i ).

[0071] Step 3: For the airfoil described in step 2, take the leading edge point of the airfoil (m i_le ,u i_le ) as the reference, translate all the coordinates of the airfoil along the horizontal axis of the coordinate system |m i_le | distance, translation along the vertical axis of the coordinate system|u i_le | distance so that the leading edge of the airfoil coincides with the coordinate origin (0, 0) after translation.

[0072] Step 4: Rotate the airfoil obtained in step 3 with the leading edge point (0, 0) as the rotation center so that the chord length of the airfoil is along the m direction. The rotation angle at this time is θ i, the airfoil line coordinates at this time (m i，rot ,u i，rot )for:

[0073] m i，rot =m i ·cosθ-u i sinθ

[0074] u i，rot =u i ·cosθ+m i sinθ

[0075] Step 5: non-dimensionalize the airfoil line coordinates obtained in step 4. The airfoil line coordinates (M i , U i )for:

[0076] M i =m i，rot / C i

[0077] U i =u i，rot / C i

[0078] Step 6, determining the mid-camber coordinates of the airfoil according to the airfoil of the i-th flow surface described in step 5;

[0079] Step 7, determining the thickness distribution coordinates of the airfoil according to the airfoil of the i-th flow surface described in step 5;

[0080] Step 8: construct the relationship between the NURBS curve control points and the geometric coordinates of the airfoil mid-arc line described in step 6. Figure 1 As shown, the mid-camber line is described by two segments of cubic NURBS curves. The curve endpoints 1 and 7 correspond to the leading edge point and the trailing edge point of the airfoil, respectively. The leading edge point is defined as the intersection of the mid-camber line and the leading edge of the airfoil, and the trailing edge point is defined as the intersection of the mid-camber line and the trailing edge of the airfoil. The ordinate of the connection point 4 is the maximum relative deflection. The horizontal axis is the relative distance of the maximum deflection from the leading edge of the airfoil In order to ensure the first-order continuity of the middle arc curve, points 3, 4 and 5 are collinear, and points 2, 3, 5 and 6 move on the corresponding control line segments. To this end, define the parameters:

[0081] λ1=l 12 / l 11′

[0082] λ2=l 34 / l 1′4

[0083] λ3=l 45 / l47′

[0084] λ4=l 67 / l 7′7

[0085] The variation range of λ1~λ4 is (0, 1). Thus, the arc coordinates (a i_NURBS , C i_NURBS ).

[0086] Step 9, construct the relationship between the NURBS curve control points and the airfoil thickness distribution coordinates described in step 7. Figure 2 As shown, the thickness distribution is described by two segments of cubic NURBS curves. The curve endpoints 8 and 14 correspond to the leading edge and trailing edge points of the airfoil respectively; the ordinate of the connection point 11 is the maximum relative thickness. The horizontal axis is the relative distance from the maximum thickness to the leading edge of the airfoil In order to ensure the first-order continuity of the thickness distribution curve, points 10, 11 and 12 are collinear, and points 10, 11 and 12 move on the corresponding control line segments respectively. To this end, define the parameters:

[0087] λ5=l 89 / l 88′

[0088] λ6=l 1011 / l 8′11

[0089] λ7=l 1112 / l 1114′

[0090] λ8=l 1314 / l 14′14

[0091] The variation range of λ5~λ8 is (0, 1). Thus, the parameterized thickness distribution coordinates (e i_NURBS , t i_NURBS ).

[0092] Step 10, generating an airfoil by superimposing the parameterized mid-camber line described in step 8 and the parameterized thickness distribution described in step 9. At this time, the thickness of the airfoil at the leading / trailing edge point is 0, while the thickness of the mid-camber line dividing point below / upstream of the leading / trailing edge point is not 0, and there is a sharp corner at the leading / trailing edge of the airfoil.

[0093] Step 11, constructing the relationship between the NURBS curve control points and the airfoil leading edge coordinates described in step 2. Figure 3 As shown in Figure 1, the leading edge of the airfoil is described by two segments of cubic NURBS curves. Point D corresponds to the leading edge of the airfoil; point S c Corresponding to the connection point between the leading edge and the suction surface, P cCorresponding to the connection point between the leading edge and the pressure surface, line SS c is the extension line of the suction surface, line PP c is the extension line of the pressure surface, and the two are respectively at point S c and P c The line SP passes through the leading edge point D, and its direction is perpendicular to the tangent of the mid-arc line at the leading edge point, ensuring the leading edge curve is first-order continuous at point D. The line SP is connected to the line SS. c Intersecting at point S, line SP and line PP c Intersecting at point P, points S1, S2, P1 and P2 move on the corresponding control line segments. c As the control point, the suction surface edge can be parameterized. Similarly, points D, P1, P2 and P c As a control point, the pressure surface edge can be parameterized. To do this, define the parameters:

[0094]

[0095]

[0096]

[0097]

[0098] and The change interval is (0, 1].

[0099] Step 12, construct the relationship between the NURBS curve control points and the airfoil trailing edge coordinates described in step 2. Figure 4 As shown in Figure 1, the trailing edge of the airfoil is described by two segments of cubic NURBS curves. Point D′ corresponds to the trailing edge of the airfoil; point S c ′ corresponds to the connection point between the trailing edge and the suction surface, P c ′ corresponds to the connection point between the trailing edge and the pressure surface, and the line S′-S c ′ is the extension line of the suction surface, line p′-P c ′ is the extension line of the pressure surface, and the two are respectively at point S c ′ and P c ′ is tangent to the trailing edge, ensuring the first-order continuity of the trailing edge and the blade body at the connection point; the line S′-P′ passes through the trailing edge point D′, and its direction is perpendicular to the tangent of the mid-camber line at the trailing edge point, ensuring the first-order continuity of the trailing edge curve at point D′; the line S′-P′ is connected to the line S′-S c ′ intersects at point S′, and the line S′-P′ intersects the line P′-P c ' intersects at point P', and points S1', S2', P1' and P2' move on the corresponding control line segments.c ′ can be used as a control point to parameterize the trailing edge of the suction surface. Similarly, points D′, P1′, P2′ and P c ′ can be used as a control point to parameterize the trailing edge of the pressure surface. To do this, define the parameters:

[0100]

[0101]

[0102]

[0103]

[0104] and The change interval is (0, 1].

[0105] Step 13, using step 11 to modify the leading edge of the airfoil described in step 10, and using step 12 to modify the trailing edge of the airfoil described in step 10, to generate a complete airfoil, the geometric coordinates of the complete airfoil are (M i_NURBS , U i_NURBS );

[0106] Step 14, repeat steps 1 to 13 for the airfoil of the second row of blades in the double row of blades. The parameterized geometric coordinates of the second row of complete airfoil are (M i_NURBS ′,U i_NURBS ′);

[0107] Step 15, after obtaining the double-row blade profile flow surface coordinates, perform parametric design of the double-row stator modification tandem, as shown in the flow chart Figure 5 As shown. According to the actual assembly dimensions, the effective chord length C of the tandem airfoil is given i，eff The effective chord length is defined as the line connecting the leading edge point of the first row of airfoils and the trailing edge point of the second row of airfoils. i，eff Set to the airfoil chord length C of the first row of blades in step 1 of the prototype double-row blade i，bladel With k i times the airfoil chord length C of the second row of blades i，blade2 The sum of the second row of leaves is defined as the second row of leaves in step 14: C i，eff =C i，blade1 +k i ·C i，blade2 . k i =1.2~2.0.

[0108] Step 16: Given the axial overlap AO of the serial configuration parameter i The axial overlap is defined as the m-direction length Δm of the overlapping portion of the first row of airfoils and the second row of airfoils in the series. i，1The m-direction distance Δm between the leading edge point of the first row of airfoils and the trailing edge point of the second row of airfoils i，2 The ratio of AO i =Δm i，1 / Δm i，2 AO i Take -0.1 to 0.1 to determine the position relationship between the trailing edge points of the first row of airfoils and the leading edge points of the second row of airfoils in the m direction.

[0109] Step 17, given the string configuration parameter chord length ratio CR i value, the CR i The value is defined as the ratio of the chord length of the second row of airfoils in the series to the chord length of the first row of airfoils, that is, CR i =C i，AA / C i，FA CR i Take 0.5~2.0.

[0110] Step 18, solve the chord length C of the first row of tandem blades by the following formula: i，FA , determine the coordinates of the trailing edge points of the first row of airfoils in the series,

[0111]

[0112] Step 19, solve the chord length C of the second row of tandem blades by the following formula: i，AA , determine the position of the trailing edge point of the second row of tandem airfoils,

[0113] C i，AA =C i，FA CR i .

[0114] Step 20, give the effective consistency value σ of the tandem blades eff , the effective consistency value is determined by the effective chord length described in step 15.

[0115] Step 21, determine the tandem blade pitch s by the following formula, wherein the pitch is defined as the circumferential distance between the leading edge point of the first row of the tandem airfoil and the leading edge point of the first row of the tandem airfoil of the adjacent blade path. Wherein, the effective chord length C in step 15 is i，eff and the axial overlap AO described in step 16 i Take the flow surface at 50% blade height described in step 1, i.e. i=3,

[0116]

[0117] Step 22, solve the blade number z by the following formula, where R is the radius of the flow surface at 50% blade height,

[0118] z=2πR / s.

[0119] Step 23, if the number of blades is an integer, proceed to step 24; otherwise, return to step 20.

[0120] Step 24, given the circumferential pitch ratio PP i The circumferential pitch ratio is defined as the circumferential distance t between the trailing edge of the first row of tandem airfoils and the leading edge of the second row of tandem blades in the adjacent blade path. i and the ratio of the grating pitch s described in step 21, that is, PP = t i / s.PP i Take 0.85 to 0.95. This determines the positional relationship between the trailing edge points of the first row of airfoils and the leading edge points of the second row of airfoils in the u direction.

[0121] Step 25, obtain the first row of airfoil coordinates (M i_tan_FA , U i_tan_FA ) and the second row of airfoil coordinates (M i_tan_AA , U i_tan_AA ).

[0122] Step 26, multiplying the coordinates of the first row of airfoils in the series described in step 25 by the chord length of the first row of airfoils,

[0123] m i_tan _ FA =M i_tan_FA ·C i，FA

[0124] u i_tan_FA =U i_tan_FA ·C i，FA

[0125] Multiply the second row of airfoil coordinates in the series described in step 25 by the chord length of the second row of airfoils,

[0126] m i_tan_AA =M i_tan_AA ·C i，AA

[0127] u i_tan_AA =U i_tan_AA ·C i，AA

[0128] Get the coordinates of the first row of airfoils in the series (m i_tan_FA ,u i_tan_FA ) and the second row of airfoil coordinates (m i_tan_AA ,u i_tan_AA ).

[0129] Step 27, rotating the airfoil described in step 26 with the leading edge point of the first row of airfoils in series as the rotation center, and the rotation angle at this time is θ described in step 4 i , the coordinates of the first row of airfoils in the series (m i_tan_FA，rot ,u i_tan_FA，rot)for:

[0130] m i_tan_FA，rot =m i_tan_FA ·cosθ i +u i_tan_FA sinθ i

[0131] v i_tan_FA，rot =m i_tan_FA ·cosθ i -u i_tan_FA sinθ i

[0132] The second row of airfoil coordinates in series (m i_tan_AA，rot ,u i_tan_AA，rot )for:

[0133] m i_tan_AA，rot =m i_tan_AA ·cosθ i +u i_tan_AA sinθ i

[0134] u i_tan_AA，rot =m i_tan_AA ·cosθ i -u i_tan_AA sinθ i

[0135] Step 28: Taking the leading edge point of the first row of airfoils in the series as the reference, translate all coordinates of the series airfoils along the horizontal axis of the coordinate system by -|m i_le | distance, translation along the vertical axis of the coordinate system -|u i_le | distance. Make the leading edge points of the first row of tandem airfoils coincide with the leading edge points of the blades described in step 2 after translation. The coordinates of the tandem airfoils at this time are (m i_tan ,u i_tan ).

[0136] Step 29, the conversion relationship between the known stream surface coordinates and the cylindrical coordinates is:

[0137]

[0138] u i，tan =r i ·θ i

[0139] The radial coordinate r of the airfoil is determined by the stream surface position obtained in step 1 i , can solve the parametric geometric cylindrical coordinates of the tandem airfoil (r i ,θ i , z i ).

[0140] Step 30, repeating steps 1 to 29 for the airfoil of the i-th flow surface described in step 2. If i=5, proceed to step 31; otherwise, return to step 2. Obtain the parametric geometric cylindrical coordinates of the airfoils on all the flow surfaces described in step 1.

[0141] Step 31, stacking the tandem airfoil described in step 30 according to the five flow surfaces described in step 1 in the order of i=1 to i=5 along the blade height, the stacking points are all the leading edge points of the airfoil of the first row of blades in the double-row blades described in step 1 on the flow surfaces taken at different blade heights, and the stacking line is the leading edge line of the first row of blades in the double-row blades described in step 1. The double-row blade is converted into a tandem blade.

[0142] Step 32: If there is no geometric interference after parametric modeling, the modification design of replacing double-row blades with tandem blades is completed. The geometric comparison before and after the modification is as follows: Figure 6 As shown; otherwise, return to step 15.

[0143] Although the embodiments of the present invention are described above in conjunction with the accompanying drawings, the present invention is not limited to the above specific embodiments and application fields, and the above specific embodiments are only illustrative and instructive, rather than restrictive. A person of ordinary skill in the art can also make many forms under the guidance of this specification and without departing from the scope of protection of the claims of the present invention, all of which belong to the protection of the present invention.

Claims

1. A method for modifying the double-row blades of a turbine compressor in series, characterized in that: It includes the following steps, Step 1: Take i flow surfaces along the blade height of the first row of double-row blades of the turbine compressor and obtain the geometric coordinates of the airfoil on each flow surface; Step 2: Based on the airfoil geometric coordinates, the airfoil on the i-th flow surface is translated, rotated and dimensionless, so that the leading edge point of the airfoil coincides with the origin of the coordinates, and the chord length of the airfoil is along the m direction, where the m direction is the direction of the meridian streamline on the flow surface, and the mid-camber line and thickness distribution of the airfoil on the i-th flow surface are measured; Step 3: parameterizing the camber line and thickness distribution respectively using NURBS curves, superimposing the parameterized camber line and thickness distribution to generate an airfoil profile, parameterizing the leading edge and trailing edge of the airfoil respectively using NURBS curves, and then modifying the leading edge and trailing edge of the airfoil to complete the airfoil parameterization of the i-th flow surface of the first row of blades; Step 4: Take the i flow surfaces along the blade height of the second row of blades in the double-row blades respectively, obtain the airfoil geometric coordinates on each flow surface, and repeat steps 2 to 3 to complete the airfoil parameterization of the i-th flow surface of the second row of blades; Step 5: Determine the following parameters based on the assembly dimensions: effective chord length C i，eff is the airfoil chord length C of the first row of blades in the double row of blades i，blade1 With k i times the airfoil chord length C of the second row of blades i，blade2 The sum of C i，eff =C i，blade1 +k i ·C i，blade2 ; Axial overlap of the series AO i The length Δm in the m direction of the overlapping part of the first row of blade airfoils and the second row of blade airfoils in the series i，1 The m-direction distance Δm between the leading edge of the first row of blades and the trailing edge of the second row of blades i，2 Ratio, AO i =Δm i，1 / Δm i，2 , the axial overlap of the series AO i Determine the positional relationship between the trailing edge points of the first row of airfoils in the series and the leading edge points of the second row of airfoils along the m direction; the chord length ratio CR of the series i The value is the ratio of the chord length of the airfoil of the second row of blades in series to the chord length of the airfoil of the first row of blades, CR i =C i，AA / C i，FA ; Step 6: Determine the chord length C of the airfoil of the first row of blades in the tandem i，FA and the chord length C of the airfoil of the second row of blades i，AA , thereby determining the trailing edge point position of the airfoil of the first row of blades in series and the trailing edge point position of the airfoil of the second row of blades in series, based on the effective chord length C i，eff , axial overlap AO i Chord length ratio CR i , C i，AA =C i，FA CR i ; Step 7: Determine the reasonable effective consistency value σ of the tandem blades based on experience eff ; Step 8: Determine the tandem blade pitch s based on the effective chord length C i，eff , effective consistency value σ eff and axial overlap AO i , The grating pitch s is the circumferential distance between the leading edge point of the airfoil of the first row of blades in series and the leading edge point of the airfoil of the first row of blades in series in the adjacent blade channel. The number of blades z is determined based on the grating pitch s, z=2πR / s. If the number of blades z is an integer, continue; otherwise, return to step 7 until the number of blades z is an integer; Step 9: Determine the circumferential pitch ratio PP of the position relationship between the trailing edge points of the first row of tandem airfoils and the leading edge points of the second row of tandem airfoils along the u direction i t is the circumferential distance between the trailing edge of the first row of blades in the series and the leading edge of the second row of blades in the adjacent series i The ratio of PP to the grating pitch s is t i / s, the u direction is a direction quasi-orthogonal to the m direction on the flow surface, based on the circumferential pitch ratio PP i Obtaining the coordinates of the airfoil of the first row of blades in series and the airfoil of the second row of blades in series; Step 10: The coordinates of the airfoil of the first row of blades in the series and the airfoil of the second row of blades in the series are multiplied by the corresponding airfoil chord length, and are rotated and translated so that the leading edge point of the airfoil of the first row of blades in the series coincides with the leading edge point of the airfoil on the i-th flow surface of the first row of blades in the double row of blades, and the flow surface coordinates of the series airfoil are converted into cylindrical coordinates to obtain the parameterized geometry of the series airfoil in the cylindrical coordinate system; Step 11: Repeat steps 2 to 10 for the airfoil on the i-th flow surface to obtain the parameterized geometric cylindrical coordinates of the airfoils in series on all flow surfaces; the serial airfoils are stacked in sequence according to the flow surfaces to realize the transformation of the double-row blades into serial blades, the stacking points are all the leading edge points of the airfoils on the flow surfaces taken at different blade heights for the first row of blades in the double-row blades, and the stacking line is the leading edge line of the first row of blades in the double-row blades.

2. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: Preferably, in step 1, among the i flow surfaces, i=1 is the flow surface at 0% leaf height, i=2 is the flow surface at 25% leaf height, i=3 is the flow surface at 50% leaf height, i=4 is the flow surface at 75% leaf height, and i=5 is the flow surface at 100% leaf height.

3. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: In the step 1, two-segment, third-order NURBS curves are used to parameterize the leading edge and the trailing edge of the airfoil respectively. In the step 3, two-segment, third-order NURBS curves are used to parameterize the center camber line and the thickness distribution respectively. The parameterized center camber line and the thickness distribution are superimposed to generate the airfoil. The parameterized leading edge and trailing edge are used to modify the leading edge and trailing edge of the parameterized airfoil to complete the airfoil parameterization of the i-th flow surface of the first row of blades.

4. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: The k i is 1.2 to 1.8, the AO i -0.1~0.1, CR i It is 0.5~2.

0.

5. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: The effective chord length C i，eff and axial overlap AO i is the flow surface at 50% blade height, i is 3, and the grating distance s is, The blade number z is z=2πR / s, where R is the radius of the flow surface at 50% of the blade height.

6. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: PP i It is 0.85~0.

95.

7. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: In the step 11, the five flow surfaces are stacked along the blade height in the order of i=1 to i=5, the stacking points are all the leading edge points of the airfoil of the first row of blades in the double-row blades on the flow surface at different blade heights, and the stacking line is the leading edge line of the first row of blades in the double-row blades. If there is no geometric interference, the modification of replacing the double-row blades with the tandem blades is completed; Otherwise, return to step 5.

8. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: The double-row blades of the turbine compressor are stator blades.

9. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: The double-row blades have double rows with equal blade numbers.

10. The method for modifying the double-row blades of a turbine compressor in tandem according to claim 1, wherein: The double row of blades has two rows with unequal numbers of blades.

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