CN117473659A - High Mach number aero-engine fan rotor blade profile optimization construction method - Google Patents

High Mach number aero-engine fan rotor blade profile optimization construction method Download PDF

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Publication number
CN117473659A
CN117473659A CN202311251260.1A CN202311251260A CN117473659A CN 117473659 A CN117473659 A CN 117473659A CN 202311251260 A CN202311251260 A CN 202311251260A CN 117473659 A CN117473659 A CN 117473659A
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CN
China
Prior art keywords
blade
profile
relative
mach number
line
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Pending
Application number
CN202311251260.1A
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Chinese (zh)
Inventor
王姝朗
刘杰
吴松洋
孟德君
尹海宝
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AECC Shenyang Engine Research Institute
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AECC Shenyang Engine Research Institute
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Priority to CN202311251260.1A priority Critical patent/CN117473659A/en
Publication of CN117473659A publication Critical patent/CN117473659A/en
Pending legal-status Critical Current

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    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/17—Mechanical parametric or variational design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • 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
    • G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2113/00—Details relating to the application field
    • G06F2113/08—Fluids
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces

Abstract

The application belongs to the technical field of high Mach number aeroengine fan rotor profile design, and particularly relates to a high Mach number aeroengine fan rotor profile optimization construction method, which comprises the steps of establishing a functional relation of blade roots, blade middle camber lines and blade tip camber lines relative to a bending angle: y "=a 0 +a 1 cos(πx)+b 1 sin(πx)+a 2 cos(2πx)+b 2 sin(2πx)+a 3 cos(3πx)+b 3 sin (3πx) +a4cos4πx+b4sin4πx; x=ζ/C; wherein y' is the relative angle of camber line E [ -1,1]The method comprises the steps of carrying out a first treatment on the surface of the x is the dimensionless coordinate of the direction of the chord line of the mean camber lineThe method comprises the steps of carrying out a first treatment on the surface of the ζ is the direction coordinate of the chord line of the mean camber line; c is the chord length of the mean camber line; a, a 0 、a 1 、a 2 、a 3 、a 4 、b 1 、b 2 、b 3 、b 4 Fitting coefficients; determining the relative bent angle distribution of camber lines of a blade root, a blade middle and a blade tip; symmetrically superposing thicknesses on the blade root, the blade middle and the blade tip camber lines to obtain a blade profile, wherein the blade profile thickness is determined by adopting a quadratic function: p (x) =ax 2 +bx+c; wherein p (x) is the thickness of the leaf profile at x on the mean camber line; a. and b and c are blade profile thickness quadratic function control parameters.

Description

High Mach number aero-engine fan rotor blade profile optimization construction method
Technical Field
The application belongs to the technical field of high Mach number aeroengine fan rotor blade profile design, and particularly relates to a high Mach number aeroengine fan rotor blade profile optimization construction method.
Background
When the incoming flow Mach number of the aeroengine exceeds the sound speed, shock waves are generated in the fan rotor blade profile channel in a return mode, and as the incoming flow Mach number is increased, the blade profile loss is increased in geometric progression, the loss is overlarge, the generated pressure rise is small, the flow capacity is poor, and the practical application value is not achieved.
The present application has been made in view of the existence of the above-mentioned technical drawbacks.
It should be noted that the above disclosure of the background art is only for aiding in understanding the inventive concept and technical solution of the present application, which is not necessarily prior art to the present application, and should not be used for evaluating the novelty and creativity of the present application in the case where no clear evidence indicates that the above content has been disclosed at the filing date of the present application.
Disclosure of Invention
The invention aims to provide a method for optimally constructing a blade profile of a fan rotor of a high Mach number aeroengine, which overcomes or alleviates at least one technical defect existing in the prior art.
The technical scheme of the application is as follows:
a method for optimizing and constructing a blade profile of a fan rotor of a high Mach number aeroengine comprises the following steps:
establishing a functional relation of camber lines of blade roots, blade centers and blade tips relative to a bent angle:
y″＝a 0 +a 1 cos(πx)+b 1 sin(πx)+a 2 cos(2πx)+b 2 sin(2πx)+
a3cos3πx+b3sin3πx+a4cos4πx+b4sin4πx；
x＝ξ/C；
wherein,
y' is the relative angle of camber line, e < -1 >, 1 >;
x is the dimensionless coordinate of the direction of the chord line of the mean camber line;
ζ is the direction coordinate of the chord line of the mean camber line;
c is the chord length of the mean camber line;
a 0 、a 1 、a 2 、a 3 、a 4 、b 1 、b 2 、b 3 、b 4 fitting coefficients;
determining the relative bent angle distribution of camber lines of a blade root, a blade middle and a blade tip;
symmetrically superposing thicknesses on the blade root, the blade middle and the blade tip camber lines to obtain a blade profile, wherein the blade profile thickness is determined by adopting a quadratic function: p (x) =ax 2 +bx+c；
Wherein,
p (x) is the thickness of the profile at x on the mean camber line;
a. and b and c are blade profile thickness quadratic function control parameters.
According to at least one embodiment of the present application, in the method for optimizing and constructing a rotor profile of a fan of a high mach number aeroengine, a defined profile angle is between 5 ° and 55 °, wherein a reference profile angle of a blade root, a blade middle and a blade tip is 50.5 °, 10 ° and 5.5 °.
According to at least one embodiment of the present application, in the method for optimizing and constructing a blade profile of a fan rotor of an aeroengine with high mach number, a relative position of maximum thickness of the defined blade profile is between 0.45 and 0.75, wherein the relative positions of maximum thicknesses of a blade root, a blade middle and a blade tip are 0.45, 0.60 and 0.75.
Drawings
FIG. 1 is a schematic illustration of relative angular distribution of camber lines defining a root, a mid-blade, and a tip camber line provided by an embodiment of the present application;
FIG. 2 is a schematic illustration of determining blade root, mid-blade, and tip profile thickness profiles provided by an embodiment of the present application;
FIG. 3 is a schematic view of a blade profile obtained by the method for optimizing and constructing a blade profile of a fan rotor of a high Mach number aero-engine according to an embodiment of the present application;
fig. 4 is a schematic view of a blade root, in-blade, and tip cross-section of a blade profile provided in an embodiment of the present application.
For the purpose of better illustrating the present embodiments, certain elements of the drawings may be omitted, enlarged or reduced and do not represent the actual product dimensions, and furthermore, the drawings are for illustrative purposes only and are not to be construed as limiting the present patent.
Detailed Description
In order to make the technical solution of the present application and the advantages thereof more apparent, the technical solution of the present application will be more fully described in detail below with reference to the accompanying drawings, it being understood that the specific embodiments described herein are only some of the embodiments of the present application, which are for explanation of the present application, not for limitation of the present application. It should be noted that, for convenience of description, only the portion relevant to the present application is shown in the drawings, and other relevant portions may refer to a general design, and without conflict, the embodiments and technical features in the embodiments may be combined with each other to obtain new embodiments.
Furthermore, unless defined otherwise, technical or scientific terms used in the description of this application should be given the ordinary meaning as understood by one of ordinary skill in the art to which this application belongs. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," "outer," and the like as used in this description are merely used to indicate relative directions or positional relationships, and do not imply that a device or element must have a particular orientation, be configured and operated in a particular orientation, and that the relative positional relationships may be changed when the absolute position of the object being described is changed, and thus should not be construed as limiting the present application. The terms "first," "second," "third," and the like, as used in the description herein, are used for descriptive purposes only and are not to be construed as indicating or implying any particular importance to the various components. The use of the terms "a," "an," or "the" and similar referents in the description of the invention are not to be construed as limited in number to the precise location of at least one. As used in this description, the terms "comprises," "comprising," or the like are intended to cover an element or article that appears before the term and that is listed after the term and its equivalents, without excluding other elements or articles.
Furthermore, unless specifically stated and limited otherwise, the terms "mounted," "connected," and the like in the description herein are to be construed broadly and refer to either a fixed connection, a removable connection, or an integral connection, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can also be communicated with the inside of two elements, and the specific meaning of the two elements can be understood by a person skilled in the art according to specific situations.
The present application is described in further detail below with reference to fig. 1-4.
In order to reduce the vane pattern loss, improve the pressure rise and better control the airflow flow under the supersonic condition, the application provides a method for optimally constructing the vane pattern of the fan rotor of the high Mach number aeroengine, which is based on the parameterized camber line of the camber distribution and builds the vane pattern of the fan rotor by superposing the thickness distribution.
Carrying out dimensionless treatment on the coordinates of the camber lines of the blade root, the blade middle and the blade tip:
x＝ξ/C；
wherein,
x is the dimensionless coordinate of the direction of the chord line of the mean camber line;
ζ is the direction coordinate of the chord line of the mean camber line;
c is the chord length of the mean camber line.
Calculating relative bending angles of camber lines of a blade root, a blade middle and a blade tip:
wherein,
y' is the relative angle of the mean camber line;
alpha (x) is the leaf angle relative to the axial direction at x on the mean camber line;
α 1 calculating a reference blade profile angle for the camber line relative to the bent angle, wherein the reference blade profile angle can correspond to the angle with small blade root, blade middle and blade tip blade profile taking loss, large generated pressure rise and good flow capacity control;
to ensure the practicality of the leaf pattern, a boundary constraint y'. E [ -1,1] is added.
Constructing a functional relation of camber lines of a blade root, a blade middle and a blade tip relative to a bent angle:
fitting the functions in a form of Fourier series expansion, and establishing a functional relation of camber lines of blade roots, blade centers and blade tips relative to the bending angles:
y″＝a 0 +a 1 cos(πx)+b 1 sin(πx)+a 2 cos(2πx)+b 2 sin(2πx)+
a3cos3πx+b3sin3πx+a4cos4πx+b4sin4πx；
wherein,
a 0 、a 1 、a 2 、a 3 、a 4 、b 1 、b 2 、b 3 、b 4 for fitting coefficients, different values are provided for different function fitting results;
the functional relation of the blade root, the blade middle and the blade tip camber lines adopts the expression form of Fourier series, has the characteristics of good continuity, high universality, high degree of freedom of adjustable parameters and the like, and in a specific embodiment, the relative camber line distribution of the blade root, the blade middle and the blade tip camber lines is determined as shown in figure 1.
Symmetrically superposing thicknesses on the camber lines of the blade root, the blade middle and the blade tip to obtain a blade profile:
the leaf thickness is determined by a quadratic function: p (x) =ax 2 +bx+c；
Wherein,
p (x) is the thickness of the profile at x on the mean camber line;
a. b and c are blade thickness quadratic function control parameters;
the blade profile thickness is controlled by quadratic function determination, the continuity is good, the expression is simple, and the peak point is convenient to control, and in a specific embodiment, the blade root, the blade middle and the blade tip blade profile thickness distribution is determined as shown in figure 2.
In a specific embodiment, the incoming flow mach numbers of the blade root, the middle blade and the tip of the fan rotor blade are respectively 0.9, 1.4 and 1.9, for this, it is preferable to define blade profile angles between 5 ° and 55 °, wherein the reference blade profile angles of the blade root, the middle blade and the tip are 50.5 °, 10 ° and 5.5 °, and the maximum thickness relative positions of the defined blade profiles are between 0.45 and 0.75, wherein the relative positions of the maximum thicknesses of the blade root, the middle blade and the tip are 0.45, 0.60 and 0.75, and the obtained blade profile is shown in fig. 3, and the cross section of the blade profile at the blade root, the middle blade and the tip of the blade is shown in fig. 4, so that the method can be well applied to the situation that the incoming flow mach number exceeds 1.5.
In the description, each embodiment is described in a progressive manner, and each embodiment is mainly described by the differences from other embodiments, so that the same similar parts among the embodiments are mutually referred.
Having thus described the technical aspects of the present application with reference to the preferred embodiments illustrated in the accompanying drawings, it should be understood by those skilled in the art that the scope of the present application is not limited to the specific embodiments, and those skilled in the art may make equivalent changes or substitutions to the relevant technical features without departing from the principles of the present application, and those changes or substitutions will now fall within the scope of the present application.

Claims (3)

1. A method for optimizing and constructing a blade profile of a fan rotor of a high Mach number aeroengine is characterized by comprising the following steps:
establishing a functional relation of camber lines of blade roots, blade centers and blade tips relative to a bent angle:
y″＝a 0 +a 1 cos(πx)+b 1 sin(πx)+a 2 cos(2πx)+b 2 sin(2πx)+
a3cos3πx+b3sin3πx+a4cos4πx+b4sin4πx；
x＝ξ/C；
wherein,
y' is the relative angle of camber line, e < -1 >, 1 >;
x is the dimensionless coordinate of the direction of the chord line of the mean camber line;
ζ is the direction coordinate of the chord line of the mean camber line;
c is the chord length of the mean camber line;
a 0 、a 1 、a 2 、a 3 、a 4 、b 1 、b 2 、b 3 、b 4 fitting coefficients;
determining the relative bent angle distribution of camber lines of a blade root, a blade middle and a blade tip;
symmetrically superposing thicknesses on the blade root, the blade middle and the blade tip camber lines to obtain a blade profile, wherein the blade profile thickness is determined by adopting a quadratic function: p (x) =ax 2 +bx+c；
Wherein,
p (x) is the thickness of the profile at x on the mean camber line;
a. and b and c are blade profile thickness quadratic function control parameters.
2. The method for optimized construction of high Mach number aero-engine fan rotor profile as claimed in claim 1, wherein,
the blade profile angle is defined to be between 5 ° and 55 °, wherein the reference blade profile angle of the blade root, the blade center and the blade tip is 50.5 °, 10 ° and 5.5 °.
3. The method for optimized construction of high Mach number aero-engine fan rotor profile as claimed in claim 1, wherein,
the maximum thickness relative position defining the airfoil is between 0.45 and 0.75, wherein the relative positions of the maximum thicknesses of the blade root, the blade center and the blade tip are 0.45, 0.60 and 0.75.
CN202311251260.1A 2023-09-26 2023-09-26 High Mach number aero-engine fan rotor blade profile optimization construction method Pending CN117473659A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202311251260.1A CN117473659A (en) 2023-09-26 2023-09-26 High Mach number aero-engine fan rotor blade profile optimization construction method

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202311251260.1A CN117473659A (en) 2023-09-26 2023-09-26 High Mach number aero-engine fan rotor blade profile optimization construction method

Publications (1)

Publication Number Publication Date
CN117473659A true CN117473659A (en) 2024-01-30

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Family Applications (1)

Application Number Title Priority Date Filing Date
CN202311251260.1A Pending CN117473659A (en) 2023-09-26 2023-09-26 High Mach number aero-engine fan rotor blade profile optimization construction method

Country Status (1)

Country Link
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