CN115270318B - Transonic moving blade modeling method for axial-flow compressor of marine gas turbine - Google Patents
Transonic moving blade modeling method for axial-flow compressor of marine gas turbineInfo
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Abstract
The invention aims to provide a method for modeling a transonic moving blade of an axial-flow compressor of a marine gas turbine, which calculates correlation coefficients through a reliable formula summarized by a large number of experiments, realizes parameterized design of the transonic moving blade of the compressor, can rapidly and efficiently complete a high-performance transonic moving blade modeling scheme, effectively reduces iteration times of optimizing modeling of the transonic moving blade of an inlet of the compressor through three-dimensional CFD calculation, and can save a large amount of design time and shorten design period on the premise of effectively improving pneumatic design precision of the compressor and pneumatic performance level of the compressor. Meanwhile, the invention is not only limited to the axial flow compressor of the marine gas turbine, but also applicable to the pneumatic design process of various industrial axial flow compressors with transonic stages and aero-engine axial flow compressors.
Description
Technical Field
The invention relates to a gas turbine design method, in particular to a compressor movable blade design method.
Background
The performance level of the compressor, which is one of the three core components of the marine gas turbine, directly determines the realization of the technical index of the whole machine. With the continuous development of marine gas turbines, the through-flow capacity and load level of the gas compressor are also continuously improved, and the internal flow of the gas compressor is gradually transited from the traditional subsonic speed to the transonic speed. In particular, in order to meet the requirements of high flow and high pressure ratio of a unit, the inlet stages of the modern high-power marine gas turbine are designed to be transonic stages, and the performance of the transonic stages directly determines the performance index of the whole compressor.
For the transonic stage of the compressor, the flow inside the moving blade is very complex, and strong three-dimensional phenomena such as shock waves, secondary flow, boundary layers and the like exist, and the interaction and influence between the shock waves, the secondary flow, the boundary layers and the like form an unsteady and nonlinear space-time flow field inside the moving blade. The core of the compressor transonic stage with high performance is the flow control of the rotor blades. Therefore, it is necessary to develop a special research work for the compressor transonic stage design, and technological breakthroughs are continuously formed, so that the performance level of the compressor can meet the rapid development of the modern marine gas turbine.
Disclosure of Invention
The invention aims to provide a method for modeling transonic moving blades of an axial-flow compressor of a marine gas turbine, which can improve the transonic performance level of the compressor of the marine gas turbine.
The purpose of the invention is realized in the following way:
The invention discloses a method for modeling transonic moving blades of an axial-flow compressor of a marine gas turbine, which is characterized by comprising the following steps of:
(1) Determining the number of moving blades, and determining the optimal number Z of the blades by taking a transonic moving blade load coefficient as a reference standard according to the through-flow inverse problem design result of the axial-flow compressor;
(2) Determining the chord length, and determining the chord length b of each section of the blade according to the consistency of the blade cascade and the optimal number of the blades;
(3) Calculating an inlet sharp wedge angle of a supersonic part, forming corresponding tangential lines at the tangential positions of a blade back molded line and a blade basin molded line of the supersonic blade profile and an arc of the front edge of the blade profile, wherein the included angle of the tangential lines is the inlet sharp wedge angle xi, and independently calculating the blade profile inlet sharp wedge angle xi for the supersonic part of the transonic moving blade;
(4) Determining attack angle and lag angle, and calculating attack angle i and lag angle delta for each section profile of the transonic moving blade according to the profile type;
(5) Calculating an inlet and outlet geometric angle and a bent angle, obtaining an inlet airflow angle beta 1 and an outlet airflow angle beta 2 of each section of the transonic moving blade according to the through-flow inverse problem design result of the axial-flow compressor, and determining an inlet attack angle i and an outlet lag angle delta through the step (4), so as to determine an inlet geometric angle beta 1k＝β1 +i, an outlet geometric angle beta 2k＝β2 +delta and a bent angle epsilon=beta 2k-β1k of each section blade profile;
(6) Calculating an inlet and outlet wedge angle, and determining an inlet wedge angle χ 1 value of each section blade profile by calculating an equivalent radius R of each section blade profile camber line of the transonic moving blade, so as to obtain an outlet wedge angle χ 2＝ε-χ1;
(7) Calculating the blade mounting angle, and calculating the mounting angle gamma=beta 1k+χ1 of each section blade profile of the transonic moving blade according to the inlet geometric angle beta 1k and the inlet wedge angle χ 1 of each section blade profile;
(8) Selecting an original blade profile, and after determining the modeling design parameters of the blade profiles of all the sections of the transonic moving blade, selecting the original blade profile according to the working conditions of all the sections of the blade to finish the two-dimensional blade profile design of all the sections of the transonic moving blade; on the basis, a center-of-gravity stacking mode is adopted to radially stack the two-dimensional blade profiles of each section to generate a three-dimensional blade, so that a final transonic moving blade modeling scheme is formed.
The invention may further include:
1. In step (1), the optimum blade number Z is calculated as follows:
Wherein, sigma k is the consistency of the blade cascade of the blade top section; The aspect ratio of the blade; For the relative average radius of the blade(s), The ratio of the blade to the hub; y is the chord length ratio of the top part to the root part of the blade, and the value of y is related to the load coefficient psi of the blade:
when the value of the y is more than 0 and less than or equal to 0.2, the value range of the y is 1.20-1.35;
when the value of the y is less than or equal to 0.38 and less than or equal to 0.2, the value range of the y is 1.35-1.50;
the y value is selected in the above manner according to the load factor situation of the transonic rotor blade, so that the optimum number Z of blades is determined.
2. And (3) determining the chord length b of each section of the blade according to the consistency of the blade cascade and the optimal number of the blades in the step (2), wherein the following mode is adopted: firstly, calculating the chord length b k of the blade top according to the consistency sigma k of the blade top section blade cascade, the optimal blade number Z and the outer diameter d k of the transonic moving blade, namelyThen calculating the chord length of the blade root according to the determined chord length ratio y of the top and the root of the bladeAnd finally, calculating the chord lengths of the rest sections according to the linear change of the chord lengths from the top to the root.
3. The supersonic part of the transonic moving blade described in the step (3) is obtained by calculating the blade profile inlet tip wedge angle ζ alone by solving an equation related to the maximum thickness of the blade profile, the equation being as follows:
Wherein C max is the maximum thickness of the leaf shape; b p is the nominal chord length of the airfoil, and has the following relation with the chord length b of the airfoil: b p =1.05b, solving by the equation to obtain the inlet tip wedge angle ζ of the supersonic cascade.
4. The attack angle i and the lag angle delta are calculated according to the type of the blade profile in the step (4), and the following mode is adopted:
for supersonic blade profile:
Inlet angle of attack i: Wherein Deltai is supersonic cascade angle of attack correction, taking 2-4 degrees;
outlet lag angle delta: taking 1-2 degrees;
for subsonic leaf patterns:
Inlet angle of attack i:
Wherein sigma is the consistency of the blade cascade at the corresponding position of each section blade profile; The relative radius of the corresponding position of each section blade profile; m σ＝1.0 is the change rate of attack angle with the angle of attack at the blade-form consistency of 1, is a function of sigma, and m σ＝1.0＝-0.1686σ2 +0.9393 sigma+ 0.2302, if the calculation result is smaller than 0.7, m σ＝1.0 =0.7 is taken; Δβ is the airflow turning angle, Δβ=β 2-β1; The nominal value of the turning angle of the airflow at the blade-shaped consistency of 1 is a function of beta 2, Δi is the angle of attack correction, calculated from the cascade mean radius parameter,Where σ m is the consistency at the mean radius of the cascade, β 1m is the inlet air flow angle at the mean radius of the cascade, Ω m is the degree of reaction at the mean radius of the cascade;
Outlet lag angle delta:
Wherein m is the mean camber line coefficient of the blade profile, Wherein the method comprises the steps ofIs the position of the leaf profile relative to the maximum deflection; is the lag angle correction coefficient, and takes the value and the blade grid hub ratio The following are related:
Wherein a is the ratio of the axial velocity of the blade profile outlet to the axial velocity of the inlet.
The inlet attack angle i and the outlet lag angle delta of each section of the transonic moving blade are calculated in the above mode, and the attack angle and the lag angle of the supersonic part and the subsonic part of the whole blade are smoothly transited along the radial direction.
5. And (3) determining an inlet wedge angle χ 1 value of each section blade profile by calculating the equivalent radius R of each section blade profile camber line of the transonic moving blade in the step (6), wherein the method comprises the following steps:
The equivalent radius R of the camber line of the airfoil is first solved according to the following equation:
Wherein l 1 is 0.2-0.25 for the top of the blade, and gradually and steadily decreases to 0 along with the decrease of the blade profile section relative to the blade height position;
The inlet wedge angle χ 1 value is then solved according to the following equation:
for subsonic cascades, b p = b.
Thereby determining the inlet wedge angle χ 1 value of each section blade profile of the transonic rotor blade.
The invention has the advantages that:
1. the invention realizes the parameterized design of the transonic moving blades of the compressor, can rapidly and efficiently finish the modeling scheme of the transonic moving blades with high performance, effectively improves the pneumatic design precision of the compressor, and improves the pneumatic performance level of the axial-flow compressor of the marine gas turbine.
2. The method effectively reduces the iteration times of optimizing the modeling of the transonic moving blade at the inlet of the compressor through three-dimensional CFD calculation, can save a great amount of design time, and shortens the design period.
3. The invention is not only limited to the axial flow compressor of the marine gas turbine, but also applicable to the pneumatic design process of various industrial axial flow compressors with transonic stages and aero-engine axial flow compressors.
Drawings
FIG. 1 is a flow chart of the present invention;
FIG. 2 is a diagram of a definition of a transonic rotor blade modeling parameter.
Detailed Description
The invention is described in more detail below, by way of example, with reference to the accompanying drawings:
1-2, the invention relates to a method for modeling a transonic moving blade of an axial-flow compressor of a marine gas turbine, which is realized by the following steps:
step one: the number of rotor blades is determined. According to the through-flow inverse problem design result of the axial-flow compressor, determining the optimal blade number Z by taking the transonic-stage moving blade load coefficient as a reference standard;
Step two: the chord length is determined. Determining the chord length b of each section of the blade according to the consistency of the blade cascade and the optimal number of the blades;
Step three: calculating the inlet wedge angle of the supersonic part. The blade back molded line and the blade basin molded line of the supersonic blade profile can form corresponding tangential lines at the tangential position of the circular arc of the front edge of the blade profile, and the included angle of the tangential lines is the inlet sharp wedge angle xi. For the supersonic part of the transonic moving blade, the blade type inlet sharp wedge angle xi is required to be calculated independently;
Step four: the angle of attack and the angle of fall are determined. Calculating attack angle i and lag angle delta for each section profile of the transonic moving blade according to the type of the profile;
Step five: and calculating the geometric angle and the bending angle of the inlet and the outlet. And according to the through-flow inverse problem design result of the axial-flow compressor, obtaining an inlet airflow angle beta 1 and an outlet airflow angle beta 2 of each section of the transonic moving blade. Determining an inlet attack angle i and an outlet lag angle delta through the fourth step, so as to determine an inlet geometric angle beta 1k＝β1 +i and an outlet geometric angle beta 2k＝β2 +delta of each section blade profile, wherein the bending angle epsilon=beta 2k-β1k;
Step six: and calculating the wedge angle of the inlet and the outlet. Determining an inlet wedge angle χ 1 value of each section blade profile by calculating an equivalent radius R of a camber line of each section blade profile of the transonic moving blade, so as to obtain an outlet wedge angle χ 2＝ε-χ1;
Step seven: and calculating the blade mounting angle. According to the inlet geometric angle beta 1k and the inlet wedge angle χ 1 of each section blade profile of the transonic moving blade, the installation angle gamma=beta 1k+χ1 of each section blade profile can be calculated;
Step eight: the original leaf profile is selected. After the modeling design parameters of the blade profiles of all the sections of the transonic moving blade are determined, selecting proper original blade profiles according to the working conditions of all the sections of the blade, and completing the two-dimensional blade profile design of all the sections of the transonic moving blade; on the basis, a center-of-gravity stacking mode is adopted to radially stack the two-dimensional blade profiles of each section to generate a three-dimensional blade, so that a final transonic moving blade modeling scheme is formed.
The optimal blade number Z is determined by taking the transonic moving blade load coefficient as a reference standard in the step one, and the following modes are adopted:
the optimum blade number Z is calculated as follows:
Wherein, sigma k is the consistency of the blade cascade of the blade top section; The aspect ratio of the blade; For the relative average radius of the blade(s), The ratio of the blade to the hub; y is the chord length ratio of the top part to the root part of the blade, and the value of y is related to the load coefficient psi of the blade:
when the value of the y is more than 0 and less than or equal to 0.2, the value range of the y is 1.20-1.35;
When the value of the psi is more than 0.2 and less than or equal to 0.38, the value range of y is 1.35 to 1.50.
The y value is selected in the above manner according to the load factor situation of the transonic rotor blade, so that the optimum number Z of blades is determined.
The chord length b of each section of the blade is determined according to the consistency of the blade cascade and the optimal number of the blades, and the following mode is adopted:
Firstly, calculating the chord length b k of the blade top according to the consistency sigma k of the blade top section blade cascade, the optimal blade number Z and the outer diameter d k of the transonic moving blade, namely Then calculating the root chord length according to the determined chord length ratio y of the blade top and the root according to claim 2And finally, calculating the chord lengths of the rest sections according to the linear change of the chord lengths from the top to the root.
The "supersonic part of transonic moving blade, need to calculate blade profile inlet tip wedge angle ζ alone" described in step three is obtained by solving the equation related to the maximum thickness of blade profile, the equation is as follows:
wherein C max is the maximum thickness of the leaf shape; b p is the nominal chord length of the airfoil, and has the following relation with the chord length b of the airfoil: b p = 1.05b.
The inlet sharp wedge angle zeta of the supersonic blade cascade can be obtained through the equation.
The attack angle i and the lag angle delta are calculated according to the type of the blade profile, and the following mode is adopted:
for supersonic blade profile:
Inlet angle of attack i: Wherein Deltai is the supersonic cascade angle of attack correction, and is usually 2-4 degrees;
outlet lag angle delta: typically 1 deg. to 2 deg..
For subsonic leaf patterns:
Inlet angle of attack i:
Wherein sigma is the consistency of the blade cascade at the corresponding position of each section blade profile; The relative radius of the corresponding position of each section blade profile; m σ＝1.0 is the change rate of attack angle with the angle of attack at the blade-form consistency of 1, is a function of sigma, and m σ＝1.0＝-0.1686σ2 +0.9393 sigma+ 0.2302, if the calculation result is smaller than 0.7, m σ＝1.0 =0.7 is taken; Δβ is the airflow turning angle, Δβ=β 2-β1; The nominal value of the turning angle of the airflow at the blade-shaped consistency of 1 is a function of beta 2, Δi is the angle of attack correction, calculated from the cascade mean radius parameter,Where σ m is the consistency at the mean radius of the cascade, β 1m is the inlet air flow angle at the mean radius of the cascade, and Ω m is the degree of reaction at the mean radius of the cascade.
Outlet lag angle delta:
Wherein m is the mean camber line coefficient of the blade profile, Wherein the method comprises the steps ofIs the position of the leaf profile relative to the maximum deflection; is the lag angle correction coefficient, and takes the value and the blade grid hub ratio The following are related:
Wherein a is the ratio of the axial velocity of the blade profile outlet to the axial velocity of the inlet.
The inlet attack angle i and the outlet lag angle delta of each section of the transonic moving blade are calculated in the above mode, and the attack angle and the lag angle of the supersonic speed part and the subsonic part of the whole blade are ensured to be smoothly transited along the radial direction.
In the sixth step, the "determining the inlet wedge angle χ 1 value of each section blade profile by calculating the equivalent radius R of each section blade profile camber line of the transonic moving blade" adopts the following modes:
The equivalent radius R of the camber line of the airfoil is first solved according to the following equation:
Wherein l 1 is 0.2-0.25 for the top of the blade, and gradually and steadily decreases to 0 along with the decrease of the blade profile section relative to the blade height position.
The inlet wedge angle χ 1 value is then solved according to the following equation:
for subsonic cascades, b p = b.
Thereby determining the inlet wedge angle χ 1 value of each section blade profile of the transonic rotor blade.
Claims (1)
1. A transonic moving blade modeling method of a marine gas turbine axial flow compressor is characterized by comprising the following steps:
(1) Determining the number of moving blades, and determining the optimal number Z of the blades by taking a transonic moving blade load coefficient as a reference standard according to the through-flow inverse problem design result of the axial-flow compressor;
(2) Determining the chord length, and determining the chord length b of each section of the blade according to the consistency of the blade cascade and the optimal number of the blades;
(3) Calculating an inlet sharp wedge angle of a supersonic part, forming corresponding tangential lines at the tangential positions of a blade back molded line and a blade basin molded line of the supersonic blade profile and an arc of the front edge of the blade profile, wherein the included angle of the tangential lines is the inlet sharp wedge angle xi, and independently calculating the blade profile inlet sharp wedge angle xi for the supersonic part of the transonic moving blade;
(4) Determining attack angle and lag angle, and calculating attack angle i and lag angle delta for each section profile of the transonic moving blade according to the profile type;
(5) Calculating an inlet and outlet geometric angle and a bent angle, obtaining an inlet airflow angle beta 1 and an outlet airflow angle beta 2 of each section of the transonic moving blade according to the through-flow inverse problem design result of the axial-flow compressor, and determining an inlet attack angle i and an outlet lag angle delta through the step (4), so as to determine an inlet geometric angle beta 1k＝β1 +i, an outlet geometric angle beta 2k＝β2 +delta and a bent angle epsilon=beta 2k-β1k of each section blade profile;
(6) Calculating an inlet and outlet wedge angle, and determining an inlet wedge angle χ 1 value of each section blade profile by calculating an equivalent radius R of each section blade profile camber line of the transonic moving blade, so as to obtain an outlet wedge angle χ 2＝ε-χ1;
(7) Calculating the blade mounting angle, and calculating the mounting angle gamma=beta 1k+χ1 of each section blade profile of the transonic moving blade according to the inlet geometric angle beta 1k and the inlet wedge angle χ 1 of each section blade profile;
(8) Selecting an original blade profile, and after determining the modeling design parameters of the blade profiles of all the sections of the transonic moving blade, selecting the original blade profile according to the working conditions of all the sections of the blade to finish the two-dimensional blade profile design of all the sections of the transonic moving blade; on the basis, a center-of-gravity stacking mode is adopted to radially stack two-dimensional blade profiles of all sections to generate a three-dimensional blade, so that a final transonic moving blade modeling scheme is formed;
in step (1), the optimum blade number Z is calculated as follows:
Wherein, sigma k is the consistency of the blade cascade of the blade top section; The aspect ratio of the blade; For the relative average radius of the blade(s), The ratio of the blade to the hub; y is the chord length ratio of the top part to the root part of the blade, and the value of y is related to the load coefficient psi of the blade:
when the psi is more than 0 and less than or equal to 0.2, the value range of y is 1.20 to 1.35;
When the psi is more than 0.2 and less than or equal to 0.38, the value range of y is 1.35 to 1.50;
according to the load coefficient condition of the transonic moving blades, selecting y values in the above mode, so as to determine the optimal number Z of the blades;
And (3) determining the chord length b of each section of the blade according to the consistency of the blade cascade and the optimal number of the blades in the step (2), wherein the following mode is adopted: firstly, calculating the chord length b k of the blade top according to the consistency sigma k of the blade top section blade cascade, the optimal blade number Z and the outer diameter d k of the transonic moving blade, namely Then calculating the chord length of the blade root according to the determined chord length ratio y of the top and the root of the bladeFinally, calculating the chord lengths of the rest sections according to the linear change of the chord lengths from the top to the root;
the supersonic part of the transonic moving blade described in the step (3) is obtained by calculating the blade profile inlet tip wedge angle ζ alone by solving an equation related to the maximum thickness of the blade profile, the equation being as follows:
Wherein C max is the maximum thickness of the leaf shape; b p is the nominal chord length of the airfoil, and has the following relation with the chord length b of the airfoil: b p =1.05b, solving by the equation to obtain an inlet sharp wedge angle ζ of the supersonic blade cascade;
the attack angle i and the lag angle delta are calculated according to the type of the blade profile in the step (4), and the following mode is adopted:
for supersonic blade profile:
Inlet angle of attack i: wherein delta i is supersonic cascade attack angle correction quantity, and 2-4 degrees are taken;
outlet lag angle delta: taking 1-2 degrees;
for subsonic leaf patterns:
Inlet angle of attack i:
Wherein sigma is the consistency of the blade cascade at the corresponding position of each section blade profile; The relative radius of the corresponding position of each section blade profile; m σ＝1.0 is the change rate of attack angle with the angle of attack at the blade-form consistency of 1, is a function of sigma, and m σ＝1.0＝-0.1686σ2 +0.9393 sigma+ 0.2302, if the calculation result is smaller than 0.7, m σ＝1.0 =0.7 is taken; Δβ is the airflow turning angle, Δβ=β 2-β1; The nominal value of the turning angle of the airflow at the blade-shaped consistency of 1 is a function of beta 2, Δi is the angle of attack correction, calculated from the cascade mean radius parameter,Where σ m is the consistency at the mean radius of the cascade, β 1m is the inlet air flow angle at the mean radius of the cascade, Ω m is the degree of reaction at the mean radius of the cascade;
Outlet lag angle delta:
Wherein m is the mean camber line coefficient of the blade profile, Wherein the method comprises the steps ofIs the position of the leaf profile relative to the maximum deflection; is the lag angle correction coefficient, and takes the value and the blade grid hub ratio The following are related:
Wherein a is the ratio of the axial speed of the blade-shaped outlet to the axial speed of the inlet;
The inlet attack angle i and the outlet lag angle delta of each section of the transonic moving blade are calculated in the above mode, and the attack angle and the lag angle of the supersonic speed part and the subsonic part of the whole blade are in smooth transition along the radial direction;
And (3) determining an inlet wedge angle χ 1 value of each section blade profile by calculating the equivalent radius R of each section blade profile camber line of the transonic moving blade in the step (6), wherein the method comprises the following steps:
The equivalent radius R of the camber line of the airfoil is first solved according to the following equation:
Wherein l 1 is 0.2-0.25 for the top of the blade, and gradually and steadily decreases to 0 along with the decrease of the blade profile section relative to the blade height position;
The inlet wedge angle χ 1 value is then solved according to the following equation:
For subsonic cascades, b p = b;
Thereby determining the inlet wedge angle χ 1 value of each section blade profile of the transonic rotor blade.
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| CN202210685031.XA CN115270318B (en) | 2022-06-15 | Transonic moving blade modeling method for axial-flow compressor of marine gas turbine |
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| CN202210685031.XA CN115270318B (en) | 2022-06-15 | Transonic moving blade modeling method for axial-flow compressor of marine gas turbine |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102852857A (en) * | 2012-09-28 | 2013-01-02 | 哈尔滨工业大学 | High-load super transonic axial gas compressor aerodynamic design method |
| CN103790639A (en) * | 2013-12-26 | 2014-05-14 | 北京理工大学 | Method for edge strip shape modifying of front edge of end area blade of turbine |
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102852857A (en) * | 2012-09-28 | 2013-01-02 | 哈尔滨工业大学 | High-load super transonic axial gas compressor aerodynamic design method |
| CN103790639A (en) * | 2013-12-26 | 2014-05-14 | 北京理工大学 | Method for edge strip shape modifying of front edge of end area blade of turbine |
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