CN112685855A - Axial flow compressor blade type attack angle and drop relief angle calculation method - Google Patents
Axial flow compressor blade type attack angle and drop relief angle calculation method Download PDFInfo
- Publication number
- CN112685855A CN112685855A CN202011542895.3A CN202011542895A CN112685855A CN 112685855 A CN112685855 A CN 112685855A CN 202011542895 A CN202011542895 A CN 202011542895A CN 112685855 A CN112685855 A CN 112685855A
- Authority
- CN
- China
- Prior art keywords
- angle
- design
- attack
- blade
- bend
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Abstract
The invention aims to provide a method for calculating a blade profile attack angle and drop clearance angle of an axial flow compressor, which is used for finally calculating to obtain values of a designed attack angle and a designed drop clearance angle of a blade profile by calculating a designed attack angle parameter, a drop clearance angle parameter and a blade profile bend angle. The method can quickly and accurately calculate the attack angle and the drop angle of most common blade profiles of the axial flow compressor in a design state, effectively improves the design precision, effectively reduces the times of correcting the attack angle and the drop angle of each stage of blade of the multistage axial flow compressor step by step through three-dimensional calculation, realizes the flow calculation of the modeling parameters of the blades of the compressor, can save a large amount of design iteration time, and shortens the design period. Meanwhile, the method is not limited to the axial flow compressor of the gas turbine, and is also suitable for the pneumatic design process of axial flow compressors and axial flow fans of various industrial axial flow compressors and aviation engines.
Description
Technical Field
The invention relates to a gas turbine calculation method, in particular to a compressor calculation method.
Background
The compressor is one of key core components of the gas turbine, and the performance of the compressor plays a decisive role in realizing the technical indexes of the gas turbine. Throughout the compressor's history of development, all compressor designs are based on pursuit of flow capacity, pressure ratio, efficiency, and surge margin. Nowadays, with the continuous improvement of the performance of a gas turbine, the requirement on the pneumatic performance of a compressor naturally develops into higher pressure ratio, higher efficiency and larger surge margin, and the current pneumatic design system of the compressor is formed based on the requirement.
With the continuous and deep knowledge of the internal flow, the pneumatic design system of the compressor is also developed at a rapid pace, and more technical means are applied to the pneumatic design of the compressor. The essence of the compressor aerodynamic design is the process of converting the aerodynamic performance requirements of the compressor into geometric shapes. The geometric design of the blade is one of the most important key links in the process. At present, a quasi-three-dimensional pneumatic design based on S1 and S2 flow surface theories is used as a core part of a pneumatic design system of a gas compressor at home and abroad at the present stage, and how to convert a quasi-three-dimensional flow design result into a geometric shape of a blade is provided, wherein the key point is the calculation of a blade type attack angle and a drop angle. Therefore, the accuracy of the calculation of the attack angle and the drop clearance angle of the blade profile plays an important role in realizing the pneumatic performance index of the compressor.
Disclosure of Invention
The invention aims to provide a method for calculating the angle of attack and the drop clearance of a blade profile of an axial flow compressor, which solves the problem of calculating the angle of attack and the drop clearance of the blade profile in the pneumatic design of the axial flow compressor.
The purpose of the invention is realized as follows:
the invention discloses a method for calculating a blade type attack angle and drop relief angle of an axial flow compressor, which is characterized by comprising the following steps of:
(1) extracting and calculating input parameters: extraction of the airfoil inlet flow angle β from the S2 through-flow design results1Angle beta to outlet air flow2Meridian velocity C of blade-type inletm1And outlet meridian velocity Cm2(ii) a Extracting the blade profile consistency b/t and the blade profile relative maximum thickness t from the input parameters of the blade profile designmax/b；
(2) Calculating a design attack angle parameter: according to the input parameters in the step (1), the zero bend angle 10% thickness design attack angle (i) is gradually and sequentially calculated0)10Zero bend blade type datum attack angle i0Designing the change rate n of the attack angle along with the bend angle;
(3) calculating design clearance angle parameters: according to the input parameters in the step (1), gradually and sequentially calculating the zero bend 10% thickness design drop angle (delta)0)10Zero bend blade profile datum drop relief angle delta0The rate of change m of the designed falling angle with the bend angle when the blade-shaped consistency is 1b/t＝1Design of the radial velocity correction amount of the falling angle
(4) Calculating the blade profile bend angle:
wherein [ Delta ] beta is [ beta ]2-β1Is a blade-shaped airflow turning angle; mu is a leaf-type consistency index factor, and the calculation formula is as follows:
(5) calculating a design attack angle and a design drop angle: calculating the design attack angle i of the blade profilerefAnd design clearance angle deltaref：
Design angle of attack iref：
iref＝i0+nθ
Design angle of attack deltaref：
The present invention may further comprise:
1. zero bend 10% thickness design angle of attack (i)0)10Zero bend blade type datum attack angle i0The calculation method for designing the angle of attack change rate n along with the bend angle comprises the following steps:
zero bend 10% thickness design angle of attack (i)0)10：
Zero bend blade profile datum angle of attack i0：
Wherein (k)i)profileThe correction coefficient is a reference attack angle blade profile;
the calculation formula is as follows for the relative maximum thickness correction coefficient of the reference attack angle:
designing the change rate n of the attack angle along with the bend angle:
2. zero bend 10% thickness design drop relief angle (delta)0)10Zero bend blade profile datum drop relief angle delta0The rate of change m of the designed falling angle with the bend angle when the blade-shaped consistency is 1b/t＝1Design of the radial velocity correction amount of the falling angle
The calculation method of (2) is as follows:
zero bend 10% thickness design drop relief angle (delta)0)10：
(δ0)10＝0.01β1b/t+[0.74(b/t)1.9+3b/t](β1/90)1.67+1.09b/t
Zero bend angle blade profile datum drop relief angle delta0：
Wherein (k)δ)profileIs used as a reference fall angle blade profile correction coefficient,
the relative maximum thickness correction coefficient of the reference falling angle is calculated according to the following formula:
design falling relief angle change rate m with bend angle when blade profile consistency is 1b/t＝1：
Design drop angle meridian speed correction
The invention has the advantages that:
1. the method for calculating the attack angle and the drop relief angle of the blade profile of the axial flow compressor can quickly and accurately calculate the attack angle and the drop relief angle of most common blade profiles of the axial flow compressor in a design state, effectively improves the design precision and improves the pneumatic performance of the compressor.
2. The method for calculating the angle of attack and the angle of fall of the blade profile of the axial flow compressor effectively reduces the times of correcting the angle of attack and the angle of fall of each stage of blade of the multistage axial flow compressor step by step through three-dimensional calculation, realizes the flow calculation of the blade modeling parameters of the compressor, can save a large amount of design iteration time, and shortens the design period.
3. The method for calculating the blade attack angle and the drop clearance angle of the axial flow compressor blade profile is not limited to the axial flow compressor of the gas turbine, but is also suitable for the pneumatic design process of axial flow compressors and axial flow fans of various industrial axial flow compressors and aviation engines.
Drawings
FIG. 1 is a flow chart of the present invention.
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, the method for calculating the blade profile attack angle and relief angle of the axial flow compressor is implemented by the following steps:
the method comprises the following steps: and calculating the extraction of the input parameters. Including extracting the airfoil inlet flow angle beta from the results of the S2 through-flow design1Angle beta to outlet air flow2Meridian velocity C of blade-type inletm1And outlet meridian velocity Cm2(ii) a Extracting the blade profile consistency b/t and the blade profile relative maximum thickness t from the input parameters of the blade profile designmaxB is the ratio of the total weight of the components to the total weight of the components. Wherein the airflow angles of the blade type inlet and outlet are included angles between the airflow direction and the axial direction of the compressor.
Step two: and (4) designing the calculation of the attack angle parameters. According to the input parameters in the step one, the design attack angle (i) of the zero bend angle and the thickness of 10 percent is gradually and sequentially calculated0)10Zero bend blade type datum attack angle i0And designing the change rate n of the attack angle along with the bend angle. The calculation method is as follows:
zero bend 10% thickness design angle of attack (i)0)10：
Zero bend blade profile datum angle of attack i0：
Wherein (k)i)profileThe correction coefficient of the reference attack angle blade profile is usually in the range of 0.7-1.0, and can be correspondingly selected according to different blade profiles;
the calculation formula is as follows for the relative maximum thickness correction coefficient of the reference attack angle:
designing the change rate n of the attack angle along with the bend angle:
step three: and calculating design fall angle parameters. According to the input parameters in the step one, the design drop angle (delta) of the thickness with the zero bend angle of 10 percent is gradually and sequentially calculated0)10Zero bend blade profile datum drop relief angle delta0The rate of change m of the designed falling angle with the bend angle when the blade-shaped consistency is 1b/t＝1Design of the radial velocity correction amount of the falling angle
The calculation method is as follows:
zero bend 10% thickness design drop relief angle (delta)0)10：
(δ0)10＝0.01β1b/t+[0.74(b/t)1.9+3b/t](β1/90)1.67+1.09b/t
Zero bend angle blade profile datum drop relief angle delta0：
Wherein (k)δ)profileThe value range of the blade profile correction coefficient is usually 0.7-1.0, and the blade profile correction coefficient can be correspondingly selected according to different blade profiles;
the relative maximum thickness correction coefficient of the reference falling angle is calculated according to the following formula:
design falling relief angle change rate m with bend angle when blade profile consistency is 1b/t＝1：
Design drop angle meridian speed correction
Step four: and (4) calculating the leaf profile bending angle. And calculating the blade profile bend angle theta based on the input parameters, the design attack angle parameters and the design drop angle parameters. The calculation method is as follows:
wherein [ Delta ] beta is [ beta ]2-β1Is a blade-shaped airflow turning angle; mu is a leaf-type consistency index factor, and the calculation formula is as follows:
step five: design angle of attack andand calculating the design falling angle. Calculating the design attack angle i of the blade profile on the basis of the blade profile bend angle obtained by calculation and the parameters of the design attack angle and the design clearance anglerefAnd design clearance angle deltaref. The calculation method is as follows:
design angle of attack iref：
iref＝i0+nθ
Design angle of attack deltaref：
The method for calculating the blade attack angle and the drop clearance angle of the axial flow compressor blade profile has universality, is not limited to the axial flow compressor of the gas turbine, and is also suitable for the pneumatic design process of axial flow compressors and aircraft engine axial flow compressors/fans for various industries.
Claims (3)
1. A method for calculating a blade type attack angle and drop relief angle of an axial flow compressor is characterized by comprising the following steps:
(1) extracting and calculating input parameters: extraction of the airfoil inlet flow angle β from the S2 through-flow design results1Angle beta to outlet air flow2Meridian velocity C of blade-type inletm1And outlet meridian velocity Cm2(ii) a Extracting the blade profile consistency b/t and the blade profile relative maximum thickness t from the input parameters of the blade profile designmax/b；
(2) Calculating a design attack angle parameter: according to the input parameters in the step (1), the zero bend angle 10% thickness design attack angle (i) is gradually and sequentially calculated0)10Zero bend blade type datum attack angle i0Designing the change rate n of the attack angle along with the bend angle;
(3) calculating design clearance angle parameters: according to the input parameters in the step (1), gradually and sequentially calculating the zero bend 10% thickness design drop angle (delta)0)10Zero bend blade profile datum drop relief angle delta0The rate of change m of the designed falling angle with the bend angle when the blade-shaped consistency is 1b/t＝1Design of the radial velocity correction amount of the falling angle
(4) Calculating the blade profile bend angle:
wherein [ Delta ] beta is [ beta ]2-β1Is a blade-shaped airflow turning angle; mu is a leaf-type consistency index factor, and the calculation formula is as follows:
(5) calculating a design attack angle and a design drop angle: calculating the design attack angle i of the blade profilerefAnd design clearance angle deltaref：
Design angle of attack iref：
iref＝i0+nθ
Design angle of attack deltaref：
2. The method for calculating the angle of attack and the relief angle of the blade profile of the axial flow compressor as recited in claim 1, wherein: zero bend 10% thickness design angle of attack (i)0)10Zero bend blade type datum attack angle i0The calculation method for designing the angle of attack change rate n along with the bend angle comprises the following steps:
zero bend 10% thickness design angle of attack (i)0)10：
Zero bend blade profile datum angle of attack i0：
Wherein (k)i)profileThe correction coefficient is a reference attack angle blade profile;
the calculation formula is as follows for the relative maximum thickness correction coefficient of the reference attack angle:
designing the change rate n of the attack angle along with the bend angle:
3. the method for calculating the angle of attack and the relief angle of the blade profile of the axial flow compressor as recited in claim 1, wherein: zero bend 10% thickness design drop relief angle (delta)0)10Zero bend blade profile datum drop relief angle delta0The rate of change m of the designed falling angle with the bend angle when the blade-shaped consistency is 1b/t＝1Design of the radial velocity correction amount of the falling angle
The calculation method of (2) is as follows:
zero bend 10% thickness design drop relief angle (delta)0)10：
(δ0)10＝0.01β1b/t+[0.74(b/t)1.9+3b/t](β1/90)1.67+1.09b/t
Zero bend angle blade profile datum drop relief angle delta0：
Wherein (k)δ)profileIs used as a reference fall angle blade profile correction coefficient,
the relative maximum thickness correction coefficient of the reference falling angle is calculated according to the following formula:
design falling relief angle change rate m with bend angle when blade profile consistency is 1b/t＝1：
Design drop angle meridian speed correction
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011542895.3A CN112685855A (en) | 2020-12-22 | 2020-12-22 | Axial flow compressor blade type attack angle and drop relief angle calculation method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011542895.3A CN112685855A (en) | 2020-12-22 | 2020-12-22 | Axial flow compressor blade type attack angle and drop relief angle calculation method |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN112685855A true CN112685855A (en) | 2021-04-20 |
Family
ID=75451353
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011542895.3A Pending CN112685855A (en) | 2020-12-22 | 2020-12-22 | Axial flow compressor blade type attack angle and drop relief angle calculation method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112685855A (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115186400A (en) * | 2022-06-15 | 2022-10-14 | 中国船舶重工集团公司第七0三研究所 | Method for prededesigning blade stall and blockage allowance of axial flow compressor |
| CN115186401A (en) * | 2022-06-15 | 2022-10-14 | 中国船舶重工集团公司第七0三研究所 | Method for determining subsonic cascade modeling key angle parameters of axial flow compressor |
| CN115186398A (en) * | 2022-06-15 | 2022-10-14 | 中国船舶重工集团公司第七0三研究所 | Method for determining key angle parameters of inlet guide vane model of axial flow compressor |
| CN115270318A (en) * | 2022-06-15 | 2022-11-01 | 中国船舶重工集团公司第七0三研究所 | Modeling method for transonic-grade moving blade of axial flow compressor of marine gas turbine |
| CN116561934A (en) * | 2023-07-10 | 2023-08-08 | 陕西空天信息技术有限公司 | Blade performance angle model correction method and device, electronic equipment and storage medium |
| CN116702511A (en) * | 2023-08-01 | 2023-09-05 | 中国航发四川燃气涡轮研究院 | Calculation method of adjustable guide vane lag angle |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106570213A (en) * | 2016-10-11 | 2017-04-19 | 北京航空航天大学 | Design method of variable inlet guide vane, and blade and compressor |
| CN107178526A (en) * | 2017-07-11 | 2017-09-19 | 上海电气燃气轮机有限公司 | Axial Compressor Blade Profile design method and Axial Compressor Blade Profile |
| CN107202029A (en) * | 2017-07-03 | 2017-09-26 | 南京航空航天大学 | A kind of adaptive angle of attack method of axial flow compressor |
| CN107679319A (en) * | 2017-09-29 | 2018-02-09 | 北京航空航天大学 | A kind of Algebra modeling method of circumferential pulsating stress item in through-flow model of turbine |
| CN109779972A (en) * | 2019-01-25 | 2019-05-21 | 杭州汽轮动力集团有限公司 | A kind of compressor two dimension blade profile optimization method for taking into account Low Reynolds number performance |
| US20200042670A1 (en) * | 2016-06-24 | 2020-02-06 | The Boeing Company | Prediction methods and systems for structural repair during heavy maintenance of aircraft |
| CN111027148A (en) * | 2019-10-28 | 2020-04-17 | 西安陕鼓动力股份有限公司 | Loss lag angle model automatic calibration and industrial axial flow compressor performance calculation method |
-
2020
- 2020-12-22 CN CN202011542895.3A patent/CN112685855A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20200042670A1 (en) * | 2016-06-24 | 2020-02-06 | The Boeing Company | Prediction methods and systems for structural repair during heavy maintenance of aircraft |
| CN106570213A (en) * | 2016-10-11 | 2017-04-19 | 北京航空航天大学 | Design method of variable inlet guide vane, and blade and compressor |
| CN107202029A (en) * | 2017-07-03 | 2017-09-26 | 南京航空航天大学 | A kind of adaptive angle of attack method of axial flow compressor |
| CN107178526A (en) * | 2017-07-11 | 2017-09-19 | 上海电气燃气轮机有限公司 | Axial Compressor Blade Profile design method and Axial Compressor Blade Profile |
| CN107679319A (en) * | 2017-09-29 | 2018-02-09 | 北京航空航天大学 | A kind of Algebra modeling method of circumferential pulsating stress item in through-flow model of turbine |
| CN109779972A (en) * | 2019-01-25 | 2019-05-21 | 杭州汽轮动力集团有限公司 | A kind of compressor two dimension blade profile optimization method for taking into account Low Reynolds number performance |
| CN111027148A (en) * | 2019-10-28 | 2020-04-17 | 西安陕鼓动力股份有限公司 | Loss lag angle model automatic calibration and industrial axial flow compressor performance calculation method |
Non-Patent Citations (6)
| Title |
|---|
| BULLOCK, R. O.; JOHNSEN, I. A.: "deviation angle analysis", 《AERODYNAMIC DESIGN OF AXIAL-FLOW COMPRESSORS:NASA SP-36》 * |
| 刘波; 马乃行; 杨小东; 曹志远: "适应较大叶型弯角范围的轴流压气机落后角模型", 《航空动力学报》 * |
| 王剑: "基于流线曲率法的周向弯曲叶片的性能计算", 《中国优秀硕士学位论文全文数据库工程科技II辑》 * |
| 王剑; 欧阳华; 吴亚东; 田杰; 杜朝辉: "跨声速轴流压气机非设计点特性计算", 《汽轮机技术》 * |
| 谷传文;孙丽娟;李晓东;戴韧;范根方;钱海燕: "径流叶片扩压器的气动设计与优化", 《流体机械》 * |
| 龙艳丽: "氦气压气机新型高负荷叶型初步研究", 《中国优秀硕士学位论文全文数据库工程科技II辑》 * |
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115186400A (en) * | 2022-06-15 | 2022-10-14 | 中国船舶重工集团公司第七0三研究所 | Method for prededesigning blade stall and blockage allowance of axial flow compressor |
| CN115186401A (en) * | 2022-06-15 | 2022-10-14 | 中国船舶重工集团公司第七0三研究所 | Method for determining subsonic cascade modeling key angle parameters of axial flow compressor |
| CN115186398A (en) * | 2022-06-15 | 2022-10-14 | 中国船舶重工集团公司第七0三研究所 | Method for determining key angle parameters of inlet guide vane model of axial flow compressor |
| CN115270318A (en) * | 2022-06-15 | 2022-11-01 | 中国船舶重工集团公司第七0三研究所 | Modeling method for transonic-grade moving blade of axial flow compressor of marine gas turbine |
| CN115186400B (en) * | 2022-06-15 | 2024-04-09 | 中国船舶重工集团公司第七0三研究所 | Method for predefining blade stall and blocking allowance of axial flow compressor |
| CN115186398B (en) * | 2022-06-15 | 2024-04-09 | 中国船舶重工集团公司第七0三研究所 | Method for determining key angle parameters of inlet guide vane modeling of axial flow compressor |
| CN116561934A (en) * | 2023-07-10 | 2023-08-08 | 陕西空天信息技术有限公司 | Blade performance angle model correction method and device, electronic equipment and storage medium |
| CN116561934B (en) * | 2023-07-10 | 2023-09-26 | 陕西空天信息技术有限公司 | Blade performance angle model correction method and device, electronic equipment and storage medium |
| CN116702511A (en) * | 2023-08-01 | 2023-09-05 | 中国航发四川燃气涡轮研究院 | Calculation method of adjustable guide vane lag angle |
| CN116702511B (en) * | 2023-08-01 | 2023-10-31 | 中国航发四川燃气涡轮研究院 | Calculation method of adjustable guide vane lag angle |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN112685855A (en) | Axial flow compressor blade type attack angle and drop relief angle calculation method | |
| CN110929357A (en) | Pneumatic design method for high-performance ship gas turbine compressor | |
| CN108229015B (en) | Variable working condition matching design method for high-altitude two-stage turbocharger | |
| US20120051927A1 (en) | Airfoil shape for compressor inlet guide vane | |
| CN109684597B (en) | Method for expanding all-condition characteristics of gas compressor | |
| CN109598081B (en) | Radial-flow turbine pneumatic optimization method based on data dimension reduction and multi-two-dimensional flow surface | |
| CN115186398B (en) | Method for determining key angle parameters of inlet guide vane modeling of axial flow compressor | |
| CN113434965B (en) | Ship gas turbine compressor performance optimization method based on three-dimensional flow field analysis | |
| CN109815624B (en) | Compressor stability boundary judgment method considering total pressure distortion influence of intake air | |
| RU2741172C2 (en) | Improved method of turbine compressor characteristics | |
| CN112685968A (en) | Axial flow compressor pneumatic design method based on space load customization thought | |
| CN114444331A (en) | Stage characteristic matching method of multistage axial flow compressor | |
| CN112699503B (en) | Method for designing inverse problem of S2 of axial flow compressor based on dimensionless load control parameters | |
| CN114186513A (en) | Modeling design method for axial flow compressor blade with reverse S-shaped front edge | |
| CN112685852B (en) | Load customization pneumatic optimization method for axial flow compressor | |
| CN113792502B (en) | Design method for middle stage anti-asthma air release flow under low rotation speed of gas turbine compressor | |
| CN108757516B (en) | Centrifugal fan design optimization method | |
| CN109815590B (en) | Three-dimensional blade modeling method and blade of multistage axial-flow compressor based on end region boundary layer | |
| CN113883093B (en) | Low-reaction-force compressor blade design method, movable blade and compressor | |
| Fei et al. | Application of new empirical models based on mathematical statistics in the through-flow analysis | |
| CN115270343A (en) | Axial flow turbine one-dimensional inverse problem high-reliability pneumatic design method | |
| Li et al. | Development and application of a throughflow method for high-loaded axial flow compressors | |
| CN113569498A (en) | Design method for bent stator blade at end part of axial flow compressor | |
| CN115186401A (en) | Method for determining subsonic cascade modeling key angle parameters of axial flow compressor | |
| CN115270318A (en) | Modeling method for transonic-grade moving blade of axial flow compressor of marine gas turbine |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| WD01 | Invention patent application deemed withdrawn after publication | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20210420 |