CN103806950A - Turbine blade provided with blade tip pressure surface trailing edge cutting structure - Google Patents
Turbine blade provided with blade tip pressure surface trailing edge cutting structure Download PDFInfo
- Publication number
- CN103806950A CN103806950A CN201410026120.9A CN201410026120A CN103806950A CN 103806950 A CN103806950 A CN 103806950A CN 201410026120 A CN201410026120 A CN 201410026120A CN 103806950 A CN103806950 A CN 103806950A
- Authority
- CN
- China
- Prior art keywords
- blade
- trailing edge
- turbine
- turbine blade
- tip pressure
- 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
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a novel turbine blade suitable for aircraft engines. According to the novel turbine blade structure, a blade tip pressure surface trailing edge on the turbine blade is cut. The novel turbine blade is characterized in that cutting of the end of the blade tip pressure surface trailing edge is started from the trailing edge face of the blade, extended to the position away from the trailing edge face of the blade by the distance of 1/4 the chord length of the blade and extended to be in the same plane with a blade end groove face from the end face of a blade top in the blade height direction, a sharp edge formed by cutting of the trailing edge is in arc transition. The novel blade structure is easy to machine and capable of effectively reducing leakage losses of a blade tip, improving the pneumatic performance and improving the blade efficiency.
Description
Technical field
The present invention relates to a kind of turbine blade with blade tip pressure side trailing edge excision structure that is applicable to aeroengine, the blade of this kind of structure can be able to effectively improve the efficiency of aeroengine.
Background technique
Turbine blade is the vitals of gas turbine engine, and the height of turbine efficiency is embodying the performance level of this motor to a great extent.Existing turbine blade comprises blade body, blade body has internal cooling channel, at work, the leaf top top board of gas-turbine blade and casing seal structure coordinate, it is one of principal element of turbine efficiency loss that the high-temperature fuel gas that gap between the two causes is revealed, this is the high temperature due to turbine blade work, high pressure and high-revolving rugged environment, cause the turbine disk, blade and casing are due to centrifugal force, the deformation of the generations such as thermal stress is difficult to realize the best fit design under each operating mode, between turbine blade and seal structure, always there is certain gap, and this gap can cause the leakage of high-temperature fuel gas, cause the loss in efficiency of turbine.Therefore the leakage loss that, how to reduce to greatest extent turbine blade-tip is one of focus of turbine design.The present invention, on the basis of original turbine blade, carries out Partial Resection in blade tip pressure side trailing edge end, and processing is simple, can effectively reduce blade tip and reveal loss, improves aeroperformance, improves blade efficiency.
Summary of the invention
The object of the invention is to further improve the aeroperformance of blade, improve the efficiency of blade.The high-efficiency blade that this kind is applicable to gas turbine engine comprises blade body, blade pressure surface (1), blade trailing edge face (2) and leaf top end (3).The excision exhibition of blade tip pressure side trailing edge end, to from blade trailing edge face (2), extends to 1/4 place for blade chord length apart from blade trailing edge face (2), in the high direction of leaf, extends to end of blade groove surface (4) coplanar from leaf top end (3); The sharp edge that trailing edge excision forms is with arc transition.
The invention has the advantages that: (1) this blade only carries out the excision of blade tip pressure side trailing edge end sections in original blade body, and processing is simple, is convenient to practical application; (2) this kind of structural improvement, has strengthened pressure side mainstream speed, and pressure side pressure is reduced, and has reduced the pressure reduction of pressure side and suction surface, thereby greatly reduces the gap leakage loss of blade tip, improves aeroperformance, improves blade efficiency; (3) in this invention, can greatly weaken blade exit whirlpool, reduce secondary flow loss, improve blade efficiency; (4) pressure side trailing edge is carried out to Partial Resection and reduced blade trailing edge thickness, reduce blade wake passing loss; (5) when main flow combustion gas is through blade tip trailing edge, accelerate, boundary layer attenuation, has reduced the frictional loss in boundary layer.Can find out that from various features of the present invention the blade of this kind of structure is a kind of practicality and the high-efficiency blade that can greatly improve blade efficiency.
Accompanying drawing explanation
Fig. 1 overall structure schematic diagram of the present invention
In figure: 1. blade pressure surface 2. blade trailing edge limit 3. leaf top end 4. end of blade groove surfaces
Embodiment
Below in conjunction with accompanying drawing, the present invention is described in further detail.
Shown in Fig. 1, the present invention is a kind of high-efficiency turbine blade that is applicable to gas turbine engine.The major character that this high-efficiency turbine blade is different from conventional turbine blade is that Partial Resection has been carried out in blade tip pressure side trailing edge end.It is characterized in that: the excision exhibition of blade tip pressure side trailing edge end is to from blade trailing edge face (2), extending to apart from blade trailing edge face (2) is 1/4 place of blade chord length, in the high direction of leaf, extends to end of blade groove surface (4) coplanar from leaf top end (3); The sharp edge that trailing edge excision forms is with arc transition.
High temperature mainstream gas flows along blade pressure surface (1) to blade trailing edge, due to blade tip pressure side trailing edge is carried out to Partial Resection, mainstream gas speed when through this place becomes large, according to bernoulli principle, stream pressure reduces, reduce with blade suction surface pressure reduction, in the case of the geometric gap of blade and ring wall is certain, gap stream weakens, the gas quantity flowing through from gap reduces, reveal loss thereby reduced tip clearance, improved aeroperformance, improved blade efficiency; Blade tip trailing edge is carried out to Partial Resection, blade structure is improved, can effectively weaken the whirlpool in passage, reduce secondary flow loss; Blade tip trailing edge is carried out to Partial Resection and reduced trailing edge segment thickness, reduced to a certain extent the Trailing Edge Loss of blade, improve blade efficiency; This structural improvement makes free-stream acceleration, and boundary layer thickness attenuation has reduced the frictional loss in boundary layer.
Claims (1)
1. one kind is applicable to the Novel turbine blade of aeroengine, this new structure is that the blade tip pressure side trailing edge of turbine blade is excised, it is characterized in that: the excision exhibition of blade tip pressure side trailing edge end is to from blade trailing edge face (2), extending to apart from blade trailing edge face (2) is 1/4 place of blade chord length, in the high direction of leaf, extend to end of blade groove surface (4) coplanarly from leaf top end (3), the sharp edge that trailing edge excision forms is with arc transition.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410026120.9A CN103806950A (en) | 2014-01-20 | 2014-01-20 | Turbine blade provided with blade tip pressure surface trailing edge cutting structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410026120.9A CN103806950A (en) | 2014-01-20 | 2014-01-20 | Turbine blade provided with blade tip pressure surface trailing edge cutting structure |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN103806950A true CN103806950A (en) | 2014-05-21 |
Family
ID=50704278
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201410026120.9A Pending CN103806950A (en) | 2014-01-20 | 2014-01-20 | Turbine blade provided with blade tip pressure surface trailing edge cutting structure |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN103806950A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105673089A (en) * | 2016-03-31 | 2016-06-15 | 中国船舶重工集团公司第七�三研究所 | Crown-free air film cooling rotor blade for turbine of gas turbine |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB872228A (en) * | 1959-03-05 | 1961-07-05 | Gen Dynamics Corp | Improvements in and relating to turbines and like machines |
| US3885886A (en) * | 1972-06-27 | 1975-05-27 | Mtu Muenchen Gmbh | Unshrouded internally cooled turbine blades |
| US4736504A (en) * | 1987-08-12 | 1988-04-12 | The United States Of America As Represented By The Secretary Of The Navy | Alignment method for pressure welded bladed disk |
| CN101158293A (en) * | 2007-11-01 | 2008-04-09 | 北京航空航天大学 | Guiders, method for regulating throat flow area and turbine engine thereof |
| CN101255800A (en) * | 2008-02-28 | 2008-09-03 | 大连海事大学 | Blade tip alula of turbine or steam turbine moving-blade |
| US7645123B1 (en) * | 2006-11-16 | 2010-01-12 | Florida Turbine Technologies, Inc. | Turbine blade with TBC removed from blade tip region |
| US8303254B1 (en) * | 2009-09-14 | 2012-11-06 | Florida Turbine Technologies, Inc. | Turbine blade with tip edge cooling |
-
2014
- 2014-01-20 CN CN201410026120.9A patent/CN103806950A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB872228A (en) * | 1959-03-05 | 1961-07-05 | Gen Dynamics Corp | Improvements in and relating to turbines and like machines |
| US3885886A (en) * | 1972-06-27 | 1975-05-27 | Mtu Muenchen Gmbh | Unshrouded internally cooled turbine blades |
| US4736504A (en) * | 1987-08-12 | 1988-04-12 | The United States Of America As Represented By The Secretary Of The Navy | Alignment method for pressure welded bladed disk |
| US7645123B1 (en) * | 2006-11-16 | 2010-01-12 | Florida Turbine Technologies, Inc. | Turbine blade with TBC removed from blade tip region |
| CN101158293A (en) * | 2007-11-01 | 2008-04-09 | 北京航空航天大学 | Guiders, method for regulating throat flow area and turbine engine thereof |
| CN101255800A (en) * | 2008-02-28 | 2008-09-03 | 大连海事大学 | Blade tip alula of turbine or steam turbine moving-blade |
| US8303254B1 (en) * | 2009-09-14 | 2012-11-06 | Florida Turbine Technologies, Inc. | Turbine blade with tip edge cooling |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105673089A (en) * | 2016-03-31 | 2016-06-15 | 中国船舶重工集团公司第七�三研究所 | Crown-free air film cooling rotor blade for turbine of gas turbine |
| CN105673089B (en) * | 2016-03-31 | 2018-06-29 | 中国船舶重工集团公司第七�三研究所 | A kind of Gas Turbine is without hat gaseous film control rotor blade |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN105179022B (en) | A kind of turbo blade of use leaf top rib wing structure | |
| JP4923073B2 (en) | Transonic wing | |
| US9879544B2 (en) | Turbine rotor blades with improved tip portion cooling holes | |
| US20170167275A1 (en) | Method and system for improving turbine blade performance | |
| CN104405685A (en) | Self-circulation and circumferential groove hybrid treater box for improving performance of air compressor | |
| JP2012052525A (en) | Turbine nozzle with contoured band | |
| EP3047104B1 (en) | Turbomachine with endwall contouring | |
| CN205135721U (en) | Adopt leaf top rib wing structure's turbine blade | |
| JP2010156335A (en) | Method and device concerning contour of improved turbine blade platform | |
| US9885371B2 (en) | Row of aerofoil members | |
| RU2012158322A (en) | TURBINE NOZZLE SHOVEL, TURBINE AND AERODYNAMIC PART OF A TURBINE NOZZLE SHOVEL | |
| JP2012052557A (en) | Airfoil for axial-flow type compression capable of reducing loss in low-reynolds number region | |
| JP2015536410A (en) | Exhaust gas diffuser for gas turbine | |
| JPWO2014203372A1 (en) | Radial inflow axial flow turbine and turbocharger | |
| CN103883361B (en) | Turbo blade | |
| CN102889237B (en) | Blade wheel with large blades and small blades applying front edges with sharp corners and air compressor | |
| JP2013257137A (en) | Turbine engine and aerodynamic element of turbine engine | |
| EP3196409A2 (en) | Turbine compressor vane | |
| Cao et al. | Comparison of flow mechanism of blade sweep between a transonic single rotor and a rotor in stage environment | |
| CN103806950A (en) | Turbine blade provided with blade tip pressure surface trailing edge cutting structure | |
| CN204239331U (en) | One improves Capability of Compressor self-loopa and the hybrid processor box of peripheral groove | |
| CN104929699B (en) | A kind of booster-type comb tooth of obturaging | |
| KR20160074423A (en) | Gas turbine vane | |
| CN103422913A (en) | Turbine with honeycomb inner-wall casing | |
| CN111005770B (en) | Centripetal turbine with splitter blades |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| C02 | Deemed withdrawal of patent application after publication (patent law 2001) | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20140521 |