CN112307614A - Blade root profile design method for reducing stress of blade root of compressor blade - Google Patents
Blade root profile design method for reducing stress of blade root of compressor blade Download PDFInfo
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- CN112307614A CN112307614A CN202011161313.7A CN202011161313A CN112307614A CN 112307614 A CN112307614 A CN 112307614A CN 202011161313 A CN202011161313 A CN 202011161313A CN 112307614 A CN112307614 A CN 112307614A
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- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
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- G06F—ELECTRIC DIGITAL DATA PROCESSING
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- G06F30/17—Mechanical parametric or variational design
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Abstract
The invention aims to provide a blade root profile design method for reducing the stress of a blade root of a compressor blade, which comprises the following steps: generating a blade back molded line of a blade root blade profile, generating a blade basin molded line of the blade root blade profile, stacking blade body molded lines of other sections according to a stacking rule to generate a three-dimensional blade, determining whether the blade stress index is met or not and the maximum thickness of the given blade basin thickness is not met according to a stress numerical calculation result, and repeating the steps; if the blade stress index is met, final blade coordinates are determined. According to the invention, the blade back profile of the newly designed blade profile is designed according to the advanced blade profile, the speed distribution of the blade back is controlled, the boundary layer separation is inhibited, the aerodynamic performance is good, the blade basin profile design is designed to be locally thickened according to the blade stress calculation result, and the root stress of the blade is reduced. The invention solves the problems of reducing the stress of the blade root and adjusting the frequency of the blade while considering the aerodynamic performance of the blade of the high-load compressor of the ship gas turbine.
Description
Technical Field
The invention relates to a design method of a gas turbine, in particular to a design method of a compressor blade of a marine gas turbine.
Background
The gas turbine technology, particularly the compressor technology, is relatively complex, and particularly the high-load compressor blade has the problems of overhigh blade root stress or improper blade frequency in the design process. For the problem, as shown in fig. 1, a method of integrally thickening a root blade profile is generally adopted in the design of a compressor blade, the leading edge of the blade profile manufactured by the method is thick, the blade profile loss of the blade is increased, and the speed distribution of the blade back can be influenced after the blade back profile line of the blade is changed, so that the development of a boundary layer is not facilitated. How to be the key technology in this paper, has considered the requirement that blade back flow field is not influenced and the demand that the blade root thickens.
Disclosure of Invention
The invention aims to provide a blade root profile design method for reducing the blade root stress of a compressor blade, which is used for thickening a blade root and solving the problem of overhigh blade root stress under the condition of ensuring that the blade back profile of the blade is not changed.
The purpose of the invention is realized as follows:
the invention discloses a blade root profile design method for reducing stress of a blade root of a gas compressor blade of a ship gas turbine, which is characterized by comprising the following steps:
(1) generating a blade back profile of the blade root profile: generating a blade back molded line of a blade profile by adopting a mean camber line loading thickness distribution form, wherein the blade profile mean camber line comprises two or four sections of circular arcs, determining the shape of the mean camber line and an installation angle beta b through iterative calculation according to the modeling parameters of each section of circular arc, an inlet geometric angle beta 1 and an outlet geometric angle beta 2, wherein the modeling parameters are a bend angle of each section of circular arc to a total bend angle theta i/theta t of the blade profile, i is a circular arc serial number, the chord length of each section of circular arc to a total chord length bi/bt of the blade profile, the thickness distribution is generated according to a leading edge radius Rler, a trailing edge radius Rter, a maximum thickness Tmax and a maximum thickness position loc chord length bt, the thickness distribution is generated by connecting two sections of circular arcs to a leading edge circular arc, a maximum thickness circular arc and a trailing edge circular arc, extracting data points from a discrete thickness distribution curve, and loading a blade back thickness distribution data;
(2) generating a blade basin profile of a blade root profile: generating a basin profile of the blade profile by adopting a mean camber line loading thickness distribution form, wherein the mean camber line is used for generating a blade back profile;
(3) stacking blade body molded lines with other sections according to a stacking rule to generate a three-dimensional blade;
(4) determining whether the stress index of the blade is met or not and the maximum thickness of the newly given blade basin thickness is not met according to the stress numerical calculation result, and repeating the steps (2) to (4); if the blade stress index is met, final blade coordinates are determined.
The present invention may further comprise:
1. the lobe-basin molded line generation is as follows:
(a) according to the leading edge radius Rler, the trailing edge radius Rter, the maximum thickness Tmax, the maximum thickness position loc and the chord length bt, two sections of arcs are respectively connected with the leading edge arc, the maximum thickness arc and the trailing edge arc to determine two sections of thickness distribution curves;
(b) respectively disconnecting two sections of thickness distribution curves, wherein a disconnection point of a front section of the thickness distribution curve is a point P, the thickness distribution between a front edge point and the point P is reserved, a disconnection point of a rear section of the thickness distribution curve is a point G, the thickness distribution between a tail edge point and the point G is reserved, the maximum thickness circular arc radius Tmax2 is determined according to a stress calculation result, the maximum thickness point is a point Q, the point P and the point Q are connected by using two sections of cubic polynomial curves, the point Q and the point G are connected, and a PQ section and a QG section of the thickness distribution curve are generated, wherein the constraint condition of the PQ section is as follows: the thickness distribution coordinate of the point P, the slope of the point P, the coordinate of the point Q, the slope Kq of the point Q is equal to 0, and the slope of the connection point P is equal to the slope of the connection point P; the QG segment constraint is: g point thickness distribution coordinates, G point slope Kq and Q point coordinates, wherein Q point slope Kq is 0, and the connection points Q guarantee that the slopes are equal;
(c) generating data point coordinates by a discrete leaf basin thickness distribution curve;
(d) loading a blade basin thickness distribution data point on a mean camber line to generate a blade basin molded line and a coordinate thereof;
the invention has the advantages that: the blade back profile of the blade profile is designed according to the advanced blade profile, the speed distribution of the blade back is controlled, the boundary layer separation is inhibited, the aerodynamic performance is good, the blade basin profile design is designed to be locally thickened according to the blade stress calculation result, and the root stress of the blade is reduced. Compared with the integral thickening mode of the blade profile, the method not only keeps the advanced blade profile of the blade back, but also meets the requirement of reducing the stress of the blade.
Drawings
FIG. 1 shows a root-integrated thickened leaf profile;
FIG. 2 is a schematic representation of the lobe of the present invention;
FIG. 3 is a flow chart of the present invention;
FIG. 4 is a parameterization of a profile camber line;
FIG. 5 is a parameterization of the thickness distribution of the blade back profile;
FIG. 6 is a dispersion of thickness distribution data points for a blade back profile;
FIG. 7 is a parameterization of the thickness distribution of the lobe-basin profile;
FIG. 8 is a dispersion of thickness distribution data points for a bucket line;
FIG. 9 is a dispersion of thickness distribution data points for a lobe-pot line.
Detailed Description
The invention will now be described in more detail by way of example with reference to the accompanying drawings in which:
with reference to fig. 1 to 9, the present invention is mainly characterized in that a root profile design method for local thickening design is adopted, wherein an advanced profile is adopted for a profile back profile of a blade profile, and a thickening design is performed on a local profile on the basis of an advanced profile of a blade basin profile. Fig. 2 shows a design blade profile of the present invention.
The specific process of the invention is as follows:
firstly, generating a blade back profile of a blade root blade profile. And generating a blade back molded line of the blade profile by adopting a mean camber line loading thickness distribution mode. The blade-shaped mean camber line is formed by two or four circular arcs (as shown in fig. 4), the shape of the mean camber line and the installation angle β b are determined through iterative calculation according to the modeling parameters of each circular arc, the inlet geometric angle β 1 and the outlet geometric angle β 2, the modeling parameters are the bending angle of each circular arc to the total blade-shaped bending angle θ i/θ t (i is the serial number of the circular arc), the chord length of each circular arc to the total blade-shaped chord length bi/bt (i is the serial number of the circular arc), and b2/bt is the maximum deflection position. The thickness distribution is determined according to the leading edge radius Rler, the trailing edge radius Rter, the maximum thickness Tmax, the maximum thickness position loc, and the chord length bt. As shown in fig. 5, the thickness distribution is generated by joining the leading edge arc, the maximum thickness arc and the trailing edge arc by two arcs, extracting data points by a discrete thickness distribution curve as shown in fig. 6, and finally, as shown in fig. 9, loading the blade back thickness distribution data points on the mean camber line to generate the blade back profile and the coordinates thereof.
And secondly, generating a blade basin molded line of the blade root blade profile. And generating a lobe basin molded line by adopting a mean camber line loading thickness distribution mode. The camber line used in the generation of the blade back line was used. The generation of the lobe basin molded line is as follows:
(1) and as shown in fig. 7, determining the thickness distribution curves of the AP section and the GB section by using two arcs respectively joining the leading edge arc, the maximum thickness arc and the trailing edge arc according to the leading edge radius Rler, the trailing edge radius Rter, the maximum thickness Tmax, the maximum thickness position loc and the chord length bt.
(2) And breaking the thickness distribution curves at the point P and the point G, determining the maximum thickness circular arc radius Tmax2 according to the stress calculation result, and respectively connecting the point P and the point Q and the point G by using two sections of cubic polynomial curves to generate a thickness distribution curve of the PQ section and the QG section. Wherein the PQ segment constraints are: a P point thickness distribution coordinate (bf1, t1), a P point slope Kp, a Q point coordinate (loc, Tmax2), a Q point slope Kq is 0, and a connecting point P ensures that the slopes are equal; the QG segment constraint is: a G point thickness distribution coordinate (bf2, t2), a G point slope Kq, a Q point coordinate (loc, Tmax2), a Q point slope Kq is 0, and a connecting point Q ensures that the slopes are equal;
(3) as shown in fig. 8, the discrete leaf basin thickness profile generates data point coordinates.
(4) Loading the lobe basin thickness distribution data points on the mean camber line, as shown in fig. 9, generates the lobe basin profile and its coordinates.
Thirdly, as shown in fig. 2, after the blade root profile is generated according to the above method, the blade body profiles with other sections are stacked according to the stacking rule to generate the three-dimensional blade.
Fourthly, determining whether the stress index of the blade is met or not according to the stress numerical calculation result and the maximum thickness of the newly given blade basin thickness is not met, and repeating the contents from the second step to the fourth step; if the blade stress index is met, final blade coordinates are determined.
Claims (2)
1. A blade root profile design method for reducing the stress of a blade root of a compressor blade is characterized by comprising the following steps:
(1) generating a blade back profile of the blade root profile: generating a blade back molded line of a blade profile by adopting a mean camber line loading thickness distribution form, wherein the blade profile mean camber line comprises two or four sections of circular arcs, determining the shape of the mean camber line and an installation angle beta b through iterative calculation according to the modeling parameters of each section of circular arc, an inlet geometric angle beta 1 and an outlet geometric angle beta 2, wherein the modeling parameters are a bend angle of each section of circular arc to a total bend angle theta i/theta t of the blade profile, i is a circular arc serial number, the chord length of each section of circular arc to a total chord length bi/bt of the blade profile, the thickness distribution is generated according to a leading edge radius Rler, a trailing edge radius Rter, a maximum thickness Tmax and a maximum thickness position loc chord length bt, the thickness distribution is generated by connecting two sections of circular arcs to a leading edge circular arc, a maximum thickness circular arc and a trailing edge circular arc, extracting data points from a discrete thickness distribution curve, and loading a blade back thickness distribution data;
(2) generating a blade basin profile of a blade root profile: generating a basin profile of the blade profile by adopting a mean camber line loading thickness distribution form, wherein the mean camber line is used for generating a blade back profile;
(3) stacking blade body molded lines with other sections according to a stacking rule to generate a three-dimensional blade;
(4) determining whether the stress index of the blade is met or not and the maximum thickness of the newly given blade basin thickness is not met according to the stress numerical calculation result, and repeating the steps (2) to (4); if the blade stress index is met, final blade coordinates are determined.
2. The method of blade root profile design for compressor blade root stress reduction as set forth in claim 1, wherein: the lobe-basin molded line generation is as follows:
(a) according to the leading edge radius Rler, the trailing edge radius Rter, the maximum thickness Tmax, the maximum thickness position loc and the chord length bt, two sections of arcs are respectively connected with the leading edge arc, the maximum thickness arc and the trailing edge arc to determine two sections of thickness distribution curves;
(b) respectively disconnecting two sections of thickness distribution curves, wherein a disconnection point of a front section of the thickness distribution curve is a point P, the thickness distribution between a front edge point and the point P is reserved, a disconnection point of a rear section of the thickness distribution curve is a point G, the thickness distribution between a tail edge point and the point G is reserved, the maximum thickness circular arc radius Tmax2 is determined according to a stress calculation result, the maximum thickness point is a point Q, the point P and the point Q are connected by using two sections of cubic polynomial curves, the point Q and the point G are connected, and a PQ section and a QG section of the thickness distribution curve are generated, wherein the constraint condition of the PQ section is as follows: the thickness distribution coordinate of the point P, the slope of the point P, the coordinate of the point Q, the slope Kq of the point Q is equal to 0, and the slope of the connection point P is equal to the slope of the connection point P; the QG segment constraint is: g point thickness distribution coordinates, G point slope Kq and Q point coordinates, wherein Q point slope Kq is 0, and the connection points Q guarantee that the slopes are equal;
(c) generating data point coordinates by a discrete leaf basin thickness distribution curve;
(d) and loading the blade basin thickness distribution data points on the mean camber line to generate a blade basin molded line and coordinates thereof.
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| CN116186945A (en) * | 2023-04-27 | 2023-05-30 | 中国航发四川燃气涡轮研究院 | Method for realizing variable-thickness distribution curve of blade modeling |
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Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN115062398A (en) * | 2022-04-16 | 2022-09-16 | 中国航发沈阳发动机研究所 | Blisk and curve cover amount adjusting method thereof |
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| CN116186945A (en) * | 2023-04-27 | 2023-05-30 | 中国航发四川燃气涡轮研究院 | Method for realizing variable-thickness distribution curve of blade modeling |
| CN116186945B (en) * | 2023-04-27 | 2023-08-18 | 中国航发四川燃气涡轮研究院 | Method for realizing variable-thickness distribution curve of blade modeling |
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