CN103953579A - Gas compressor rotor blade with top slit and design method - Google Patents
Gas compressor rotor blade with top slit and design method Download PDFInfo
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- CN103953579A CN103953579A CN201410169055.5A CN201410169055A CN103953579A CN 103953579 A CN103953579 A CN 103953579A CN 201410169055 A CN201410169055 A CN 201410169055A CN 103953579 A CN103953579 A CN 103953579A
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Abstract
The invention discloses a gas compressor rotor blade with a top slit and a design method, and belongs to the technical field of aviation gas turbine engine gas compressors. The gas compressor rotor blade is characterized in that blade arrangement with a slit is adopted in a gas compressor rotor blade top area, that is, on premise that the middle and lower structures of the gas compressor rotor blade are kept unchanged, the top part is slit so as to form a front row of top blades and a rear row of top blades. The gas compressor blade with the blade top slit can be directly applied to aviation gas turbine engines with a high thrust-weight ratio, and high-load large-bent angle design on the top of the gas compressor blade is achieved; on premise that the flowing loss and the axial length of a gas compressor are not increased, the power-increasing pressurizing capability of the gas compressor is greatly improved, and the pneumatic performance of the gas compressor is improved.
Description
Technical field
The present invention relates to a kind of compressor rotor blade and design method thereof with top jag, can effectively control high load compressor rotor blade top suction surface airflow breakaway, realize rotor leaf and rise the large bent angle design of load, under the prerequisite that does not increase gas compressor weight and axial dimension, increase substantially gas compressor and add merit boosting capability, promote gas compressor aeroperformance, be specially adapted to high thrust weight ratio aero gas turbine engine gas compressor.
Background technique
For the overall performance index that makes motor significantly improves, wherein most important target is to improve thrust weight ratio, this just has higher requirement to the design level of gas compressor, higher single-stage pressure ratio and efficiency, on progression still less, realizing higher load just becomes the target that gas compressor development is pursued.The raising of gas compressor single-stage load is shortened axial length for aeroengine, alleviate structure weight, reducing number of components plays an important role, but, the raising of the load of compressor stage mean this air flow stream cross blade deflection angle increase, the large bent angle blade suction surface of high load boundary layer, between shock wave and leaf top zone Complex Flows flow, interact and make compressor blade suction surface easily occur flow separation, cause the aerodynamic stability of gas compressor to reduce, under traditional blade layout, the raising of gas compressor single-stage load is mainly subject to the restriction that boundary layer separates.
In order to promote Capability of Compressor, artificers have been developed the measure that a series of regulation and control compressor rotors top separates, plunder type design unloads rotor top as adopted, or peripheral groove or the axial slit of on the casing of rotor leaf top, processing some, utilize the pumping action of processor box to unload rotor leaf jacking row, control rotor Ye Ding region flow separation thereby improve, but the three-dimensional curved introducing of plunderring design and processor casing structure of rotor blade has increased air compressor structure complexity, weight and manufacturing cost, especially processor casing structure be introduced in the efficiency that can reduce to a certain extent gas compressor when promoting gas compressor nargin.Therefore, need badly and seek a kind ofly can effectively improve compressor rotor leaf top load and not increase air compressor structure complexity and the novel air pressing machine rotor topology layout of manufacturing cost.
Summary of the invention
Technical problem to be solved by this invention is to provide a kind of reasonable in design, can effectively improve compressor rotor leaf top load level, the novel air pressing machine rotor topology layout that does not increase air compressor structure complexity simultaneously and be processed into this.
It is a kind of compressor rotor blade with top jag that the present invention solves the problems of the technologies described above adopted technological scheme, described rotor blade is spatially along being circumferentially arranged in equably on described gas compressor wheel hub, the leaf root part of all jag blades is connected with described gas compressor wheel hub and around compressorshaft rotation, the angle between adjacent two jag blades equates; Described compressor rotor blade is arranged blade in top scope behind the front-seat blade of blade row intermediate open position fork-shaped Cheng Yeding and leaf top, compressor rotor leaf is pushed up between front-seat blade and leaf top rear row blade and pull open certain axial spacing, and respectively towards blade suction surface and blade pressure face tilt, after the front-seat blade of described blade middle and lower part and blade tip and leaf, arrange blade and seamlessly transit.Keeping under the constant prerequisite of compressor rotor blade middle and lower part structure row's blade after its top jag forms leaf to push up front-seat blade and leaf top.
Described in compressor rotor blade of the present invention, compressor rotor blade is single blade in 0%～80% the high scope of leaf, in the high scope of 80%～100% leaf along blade 40%～60% axial jag in axial chord length position, and along circumferentially respectively towards blade suction surface and pressure side inclination several angle, form a leaf and push up row's blade behind front-seat blade and leaf top.The front-seat blade in compressor rotor of the present invention top and rear row's blade are that 1%～2% leaf top chord is long at 100% leaf eminence axial spacing, the front-seat blade in described compressor rotor top and rear row's blade in 100% leaf eminence along circumferentially respectively towards blade suction surface and pressure side θ/8 degree (θ=360/ leaf top blade number) that tilted, ensure that the circumferential distance of arranging between blade behind the front-seat blade in 100% leaf eminence top and top is 1/4 leaf top pitch, described two top jag blades are evenly increased to θ/8 degree at 80%～100% the high scope leaning angle of leaf by 0 degree, it is long that axial spacing is evenly increased to 1%～2% leaf top chord by 0.
Rotor blade of the present invention adopts two circular arcs or many circular arc profiles, and root of blade maximum ga(u)ge is 6%～9% of blade chord length, and blade tip maximum ga(u)ge is 2.5%～4% of blade chord length, and maximum relative thickness is radially hyperbola distribution rule.
According to a further aspect in the invention, the present invention also provides a kind of design method of the compressor rotor blade with top bifurcated, it is characterized in that, first complete common blade design according to traditional compressor design method, rotor blade adopts two circular arcs or many circular arc profiles, its root of blade maximum ga(u)ge is 6%～9% of blade chord length, blade tip maximum ga(u)ge is 2.5%～4% of blade chord length, maximum relative thickness is radially distributed as hyperbola distribution rule, adopts footpath vector product to fold mode in design, then ensure that 0%～80% the high scope blade of leaf is constant, and in the high scope of 80%～100% leaf along blade 40%～60% axial jag in axial chord length position, making to arrange the axial spacing in 100% leaf eminence between blade behind the front-seat blade in compressor rotor top and top is that 1%～2% leaf top chord is long, make to arrange behind the front-seat blade of mechanism of qi rotor top and top simultaneously blade in 100% leaf eminence along circumferentially respectively towards blade suction surface and pressure side cant angle theta/8 degree (θ=360/ leaf top blade number), jag becomes leaf and pushes up row's blade behind front-seat blade and leaf top, two jag blades are evenly increased to θ/8 degree at 80%～100% the high scope leaning angle of leaf by 0 degree, it is long that axial spacing is evenly increased to 1%～2% leaf top chord by 0, after finally making blade middle and lower part and leaf push up front-seat blade and leaf top, row's blade seamlessly transits, and forms the rotor blade structure layout of a kind of top jag.
Compared with traditional compressor, the present invention is ensureing under the constant prerequisite in blade row middle and lower part, and near axial jag leaf top scope is middle along blade row forms after leaf pushes up front-seat blade and leaf top and arranges blade, can effectively reduce on the one hand the thickness of Ye Ding region boundary layer, suppress petiolarea boundary layer and separate; On the other hand, the airspeed of front row blade pressure surface throat is higher, boundary layer to rear row's blade suction surface has acceleration, can effectively suppress to make the airflow breakaway after shock wave, under the prerequisite that does not increase compressor rotor loss, obtain larger air flow deflector and add function power, effectively improving gas compressor aeroperformance; Simultaneously, compare with traditional Tandem Blades To An Aeroengine structure, the compressor rotor blade with top jag of the present invention can effectively shorten gas compressor axial length, alleviate air compressor structure weight and component number, reduced processing cost, significantly improved the integrally-built compactness of motor.
Brief description of the drawings
Fig. 1 is the compressor rotor schematic three dimensional views that the compressor blade with top jag of the present invention is set.
Embodiment
Below in conjunction with embodiment, the present invention is described in further detail, and following examples are explanation of the invention and the present invention is not limited to following examples.
As shown in Figure 1, the present embodiment forms for the outer casing 2 of rotor compressor disk 1, gas compressor and the top jag blade 3 of high thrust weight ratio aero gas turbine engine gas compressor.In order effectively to control compressor rotor tip airflow breakaway, between gas compressor wheel hub and compressor casing, some tops jag blade 3 is circumferentially arranged on edge equably, all jag blades are single blade in 0%～80% the high scope of leaf, push up row's blade 5 behind front-seat blade 4 and leaf top and become leaf along blade row 40%～60% axial chord length position jag in the high scope of 80%～100% leaf, in design, making the axial spacing at the front-seat blade in compressor rotor top and rear row's blade 100% leaf high position place is that 1%～2% leaf top chord is long, simultaneously along circumferentially spending (θ=360/ leaf top blade number) towards blade suction surface and pressure side cant angle theta/8 respectively, making 100% leaf eminence leaf push up the circumferential distance of arranging between blade behind front-seat blade and leaf top is 1/4 leaf top pitch, ensure that two jag blades are evenly increased to θ/8 degree at 80%～100% the high scope leaning angle of leaf by 0 degree, it is long that axial spacing is evenly increased to 1%～2% leaf top chord by 0, and blade middle and lower part and tip segment are seamlessly transitted, form the rotor blade structure layout of a kind of top jag.
In sum, the present invention can be directly used in high thrust weight ratio aero gas turbine engine, by row's blade after axial jag formation leaf pushes up front-seat blade and leaf top along blade row neutral position in leaf top scope, make boundary layer development again on rear row's blade, can allow larger blade loading, suppress the airflow breakaway on heavy camber blade suction surface, under the condition that keeps gas compressor high efficiency, high aerodynamic stability, have than conventional scheme much higher add merit boosting capability.Meanwhile, the compressor blade with top jag of the present invention can effectively shorten gas compressor axial length, alleviate air compressor structure weight and component number, reduced processing cost, significantly improve the integrally-built compactness of motor.
In addition, it should be noted that, the specific embodiment described in this specification, shape, institute's title of being named etc. of its parts and components can be different.All equivalence or simple change of doing according to described structure, feature and the principle of patent design of the present invention, are included in the protection domain of patent of the present invention.Those skilled in the art can make various amendments or supplement or adopt similar mode to substitute described specific embodiment; only otherwise depart from structure of the present invention or surmount this scope as defined in the claims, all should belong to protection scope of the present invention.
Claims (5)
1. one kind has the compressor rotor blade of top jag, be applicable to high thrust weight ratio aero gas turbine engine gas compressor, described compressor rotor blade is spatially along being circumferentially arranged in equably on gas compressor wheel hub, the leaf root part of all compressor rotor blades is connected and rotates around compressorshaft with described gas compressor wheel hub, angle between adjacent two compressor rotor blades equates, it is characterized in that: each described compressor rotor blade middle and lower part is single blade, become along blade row middle part jag in top area after leaf pushes up front-seat blade and leaf top and arrange blade, described leaf pushes up between front-seat blade and leaf top rear row blade and pulls open certain axial spacing, and respectively towards blade suction surface and blade pressure face tilt, after the leaf of the single blade in described compressor rotor blade middle and lower part and tip segment pushes up front-seat blade and leaf top, row's blade seamlessly transits.
2. the compressor rotor blade with top jag according to claim 1, it is characterized in that: described compressor rotor blade is single blade in 0%～80% the high scope of leaf, described compressor rotor blade in the high scope of 80%～100% leaf along blade 40%～60% axial jag in axial chord length position, while, forms a leaf and pushes up row's blade behind front-seat blade and leaf top respectively towards blade suction surface and pressure side inclination several angle circumferentially.
3. according to claim 1 or the compressor rotor blade with top jag claimed in claim 2, it is characterized in that: described leaf pushes up front-seat blade and leaf top rear row blade is that 1%～2% leaf top chord is long at 100% leaf eminence axial spacing, after described leaf pushes up front-seat blade and leaf top, row's blade is circumferentially being spent (θ=360/ leaf top blade number) towards blade suction surface and pressure side cant angle theta/8 respectively in 100% leaf eminence, making 100% leaf eminence leaf push up the circumferential distance of arranging between blade behind front-seat blade and leaf top is 1/4 times of leaf top pitch, after described leaf pushes up front-seat blade and leaf top, the inclination angle of row's blade in 80%～100% the high scope of leaf is evenly increased to θ/8 degree by 0 degree, it is long that axial spacing is evenly increased to 1%～2% leaf top chord by 0.
4. according to the compressor rotor blade that there is leaf and back down fork blade described in claims 1 to 3, it is characterized in that: described compressor rotor blade adopts two circular arcs or many circular arc profiles, root of blade maximum ga(u)ge is 6%～9% of blade chord length, blade tip maximum ga(u)ge is 2.5%～4% of blade chord length, and maximum relative thickness is radially hyperbola distribution rule.
5. according to the design method of the compressor blade with top jag described in claim 1 to 4 any one, it is characterized in that: first complete common blade design according to traditional compressor design method, rotor blade adopts two circular arcs or many circular arc profiles, its root of blade maximum ga(u)ge is 6%～9% of blade chord length, blade tip maximum ga(u)ge is 2.5%～4% of blade chord length, maximum relative thickness radially adopts hyperbola distribution rule, adopts footpath vector product to fold mode in design, then ensure that the high scope of blade row 0%～80% leaf is constant, become to jag along the axial chord length position axis of blade 40%～60% in the high scope of 80%～100% leaf after leaf pushes up front-seat blade and leaf top and arrange blade, in design, ensure that it is that 1%～2% leaf top chord is long in 100% leaf eminence that leaf pushes up axial spacing between front-seat blade and leaf top rear row blade, after making mechanism of qi rotor leaf push up front-seat blade and leaf top, row's blade is circumferentially spent (θ=360/ leaf top blade number) towards blade suction surface and pressure side cant angle theta/8 respectively on 100% leaf eminence edge simultaneously, two jag blades inclination angle in 80%～100% the high scope of leaf is evenly increased to θ/8 degree by 0 degree, it is long that axial spacing is evenly increased to 1%～2% leaf top chord by 0, after finally making the leaf of blade middle and lower part and tip segment push up front-seat blade and leaf top, row's blade seamlessly transits, and forms a kind of top jag rotor blade structure layout.
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| CN201410169055.5A CN103953579B (en) | 2014-04-24 | 2014-04-24 | A kind of compressor rotor blade with top jag and method for designing |
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| CN201410169055.5A CN103953579B (en) | 2014-04-24 | 2014-04-24 | A kind of compressor rotor blade with top jag and method for designing |
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| CN103953579B CN103953579B (en) | 2016-08-31 |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105466689A (en) * | 2014-09-09 | 2016-04-06 | 中国航空工业集团公司沈阳发动机设计研究所 | Installation technology of gas compressor blade surface dynamic pressure measurement device |
| EP3255244A1 (en) * | 2016-05-20 | 2017-12-13 | United Technologies Corporation | Tandem tip blades and corresponding gas turbine engine |
| CN109374252A (en) * | 2018-09-28 | 2019-02-22 | 西北工业大学 | A kind of compressor tandem cascade experimental provision |
| CN109948187A (en) * | 2019-02-17 | 2019-06-28 | 中国科学院工程热物理研究所 | Centrifugal compressor wheel disc laryngeal structure optimum design method based on equal-intensity theory |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2403474A1 (en) * | 1977-09-14 | 1979-04-13 | Sulzer Ag | CENTRIFUGAL PUMP |
| SU1717864A1 (en) * | 1986-03-04 | 1992-03-07 | Всесоюзный Научно-Исследовательский И Проектно-Конструкторский Институт По Оборудованию Для Кондиционирования Воздуха И Вентиляции | Sheet blade of axial-flow fan impeller |
| JPH07279887A (en) * | 1994-04-15 | 1995-10-27 | Ishikawajima Harima Heavy Ind Co Ltd | Axial flow air compressor |
| CN201934377U (en) * | 2010-11-25 | 2011-08-17 | 先锋电器集团有限公司 | Improved centrifugal wind wheel, volute casing and turbine fan assembly |
-
2014
- 2014-04-24 CN CN201410169055.5A patent/CN103953579B/en not_active Expired - Fee Related
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2403474A1 (en) * | 1977-09-14 | 1979-04-13 | Sulzer Ag | CENTRIFUGAL PUMP |
| SU1717864A1 (en) * | 1986-03-04 | 1992-03-07 | Всесоюзный Научно-Исследовательский И Проектно-Конструкторский Институт По Оборудованию Для Кондиционирования Воздуха И Вентиляции | Sheet blade of axial-flow fan impeller |
| JPH07279887A (en) * | 1994-04-15 | 1995-10-27 | Ishikawajima Harima Heavy Ind Co Ltd | Axial flow air compressor |
| CN201934377U (en) * | 2010-11-25 | 2011-08-17 | 先锋电器集团有限公司 | Improved centrifugal wind wheel, volute casing and turbine fan assembly |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105466689A (en) * | 2014-09-09 | 2016-04-06 | 中国航空工业集团公司沈阳发动机设计研究所 | Installation technology of gas compressor blade surface dynamic pressure measurement device |
| EP3255244A1 (en) * | 2016-05-20 | 2017-12-13 | United Technologies Corporation | Tandem tip blades and corresponding gas turbine engine |
| US10151322B2 (en) | 2016-05-20 | 2018-12-11 | United Technologies Corporation | Tandem tip blade |
| CN109374252A (en) * | 2018-09-28 | 2019-02-22 | 西北工业大学 | A kind of compressor tandem cascade experimental provision |
| CN109948187A (en) * | 2019-02-17 | 2019-06-28 | 中国科学院工程热物理研究所 | Centrifugal compressor wheel disc laryngeal structure optimum design method based on equal-intensity theory |
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| CN103953579B (en) | 2016-08-31 |
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Effective date of registration: 20160118 Address after: 610225, Sichuan, Chengdu Shuangliu County Southwest Airlines Economic Development Zone Industrial Concentration Zone Applicant after: Chengdu China Sciences Aircraft Engine Co., Ltd. Address before: 100190 Beijing Haidian District North Fourth Ring Road West, No. 11 A202 Applicant before: Institute of Engineering Thermophysics, Chinese Academy of Sciences |
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Address after: 610299 Sichuan province Chengdu City Road West southwest Shuangliu Airport Economic Development Zone No. 2999 port Applicant after: Chengdu China Sciences Aircraft Engine Co., Ltd. Address before: 610225, Sichuan, Chengdu Shuangliu County Southwest Airlines Economic Development Zone Industrial Concentration Zone Applicant before: Chengdu China Sciences Aircraft Engine Co., Ltd. |
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