CN107131009B - A kind of self-locking seal structure of turbomachine and the engine with it - Google Patents
A kind of self-locking seal structure of turbomachine and the engine with it Download PDFInfo
- Publication number
- CN107131009B CN107131009B CN201710344978.3A CN201710344978A CN107131009B CN 107131009 B CN107131009 B CN 107131009B CN 201710344978 A CN201710344978 A CN 201710344978A CN 107131009 B CN107131009 B CN 107131009B
- Authority
- CN
- China
- Prior art keywords
- ring
- honeycomb
- self
- turbine disk
- seal structure
- 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.)
- Active
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
- F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
- F01D11/20—Actively adjusting tip-clearance
- F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
Abstract
The invention discloses a kind of self-locking seal structure of turbomachine and with its engine, ring is obturaged in setting between rotating disk and support ring, it obturages and honeycomb ring is set outside ring, it obturages and is equipped with revolution cavity of the opening towards rotating disk, indent arc structure is set at the low radius of the assembly engagement tooth socket of rotating disk, the tangential extended line of arc structure top end is met in the revolution cavity for obturaging ring.Cooling all the way in air system obturages gas through engaging backlash between blade and rotating disk, into in the revolution cavity for obturaging ring, another way obturages gas under the water conservancy diversion of the indent arc structure of rotating disk, it is tangential to enter in revolution cavity, two-way is cooling to be obturaged after gas converges in revolution cavity, enters mainstream channel from the axial gap between honeycomb ring and rotating disk.By using the structure, the effect of obturaging of axial gap between turbomachine inner rotation-stationary parts can effectively improve, cold air dosage is obturaged in reduction, is prevented combustion gas from flowing backward and is entered inside rotating disc cavities.
Description
Technical field
It is more specifically to adopt the present invention relates to ground gas turbine and high-temperature unit of aircraft engine turbine field
With new structure, improves engine turbine rotor axial gap and obturage effect, reduce secondary air system flow amount, prevent combustion gas from falling
It fills and enters turbine disk chamber, to reduce fuel consumption, improve engine overall efficiency.
Background technique
In ground gas turbine and the hot end turbine part of aero-engine, guider is passed through in the combustion gas of high temperature and pressure
After acceleration, the acting of impulse turbine rotor part.Needs functionally, stationary parts guider and rotatable parts turbine rotor it
Between there are axial gaps.Combustion gas, which is flowed backward, in order to prevent enters turbine disk chamber, needs to obturage these axial gaps.Usually and
Speech, it is bigger to obturage flow, and it is better to obturage effect.But excessive obturage flow, it will excessive cooling air is caused to enter master
Runner brings adverse effect to turbine performance；In addition to this, increased flow capacity is obturaged, it is meant that bigger secondary air system flow
Amount, this is very unfavorable for the performance of engine complete machine.Studies have shown that for using starting for two-stage turbine (turbine)
The air conditioning quantity for being used to obturage is reduced 50% by machine, the efficiency of engine complete machine can be made to improve 0.5%, while reducing by 0.9%
Fuel consumption.
Traditional seal structure is all by the way that cooling air is arranged, and the mode for flowing through axial gap is obturaged.In order to mention
High-performance, it is common practice to which the constantly minimum self-locking flow of obturaging of amendment (prevents high-temperature fuel gas reverse irrigation in mainstream channel from entering whirlpool
Minimum discharge needed for wheel disc chamber) calculation method, or reduce axial gap to improve engine design.But this mode begins
Eventually exist apply drawback, i.e. Theoretical Design is feasible, but in practical operation be easy by various factors interference (such as processing miss
Difference, assembly fault etc.) cause to fail.In order to further enhance the performance of engine, there is an urgent need to a kind of new seal structures, can
With meet obturage require while, reduction obturage cooling air delivery.
Summary of the invention
The present invention is intended to provide a kind of modern needs such as ground gas turbine and aero-engine of improvement carry out axial envelope
Self-locking seal structure inside tight turbomachine.By using the self-locking seal structure, can be effectively improved inside turbomachine
It rotates and obturages effect axial gap between stationary parts, cold air dosage is obturaged in reduction, prevent combustion gas from flowing backward and enter the turbine disk
Chamber improves engine overall efficiency to reduce fuel consumption.
To realize the target, the technical solution adopted by the present invention are as follows:
A kind of self-locking seal structure of turbomachine, the turbomachine are gas turbine or aero-engine, including whirlpool
Wheel disc, turbo blade obturage ring, honeycomb ring and support ring, and ring, honeycomb ring, the support ring of obturaging is static element, the whirlpool
Impeller blade assembly is fixed on the turbine disk, obturages ring described in setting between the turbine disk and adjacent support ring, described
Ring is obturaged to be fixed in the support ring, the honeycomb ring be fixed at it is described obturage in ring or support ring, it is described
The turbine disk and support ring obturage ring composition turbine disk chamber, which is characterized in that
Described obturage has axial gap between ring, honeycomb ring and the turbine disk,
The inner ring surface of the honeycomb ring is abutted with the outer ring surface for obturaging ring, the outer end face of the honeycomb ring and the branch
The preceding installation side of pushing out ring abuts, and the outer ring surface of the honeycomb ring mutually agrees with the trailing edge shape of the turbo blade, and the whirlpool
The trailing edge height of impeller blade is higher than the height of the outer ring surface of the honeycomb ring, and the height of the outer ring surface of the honeycomb ring is not less than institute
State the height on the preceding installation side of support ring；
Described obturage is equipped with revolution cavity of the opening towards the turbine disk, the turbo blade and the turbine disk
Assembly engagement backlash outlet towards the revolution cavity；
At the low radius of the assembly engagement tooth socket of the turbine disk, design has concave arc structure, the arc structure
The tangential extended line of top end meets at described obturage in the revolution cavity of ring；
The turbine disk chamber is located at the inside of the inner ring surface for obturaging ring；
The cooling gas of obturaging of the first via in air system is nibbled by the assembly between the turbo blade and the turbine disk
Close backlash, leak into it is described obturage in the revolution cavity of ring,
The second tunnel in air system is cooling to obturage gas and radially flows the turbine disk is intracavitary, described
Under the guide functions of the indent arc structure of the turbine disk, the revolution cavity of ring is obturaged described in the tangential entrance along the arc structure
In,
Gas is obturaged in the first via cooling and the cooling of the second tunnel obturages gas and converges it in the revolution cavity for obturaging ring
Afterwards, enter mainstream channel from the axial gap between the honeycomb ring and the turbine disk.
More preferably, the ring of obturaging is fixed in the support ring by welding or bolted mode.
More preferably, the cross section for obturaging ring is hollow structure, it is internal exist revolution cavity or other with cavity
The structure of shape, such as ellipse, round, shuttle shape etc..
More preferably, it is assemblied between the turbo blade and the turbine disk on the turbine disk and is assembled by engaging tooth, such as tenon
The revolution cavity that ring is obturaged described in the radial height position face of toothing is engaged in tooth, dovetail toothing, assembly.
More preferably, at the low radius of the blade assembly engagement tooth socket of the turbine disk, design has concave arc structure,
The concave circular arcs structure can optimize to obtain, and the tangential extended line of the top end of the arc structure meets at time for obturaging ring
Turn in cavity, good guide functions can be played.
More preferably, from the air-flow of air system, wherein one engages between cog by the assembly between blade and the turbine disk
Gap leaks into the revolution cavity for obturaging ring, obturages gas for first via cooling, this strand of air flow rate is smaller, and in the turbine disk
Under effect, into having biggish tangential velocity when obturaging ring cavity.
More preferably, from the air-flow of air system, wherein second strand of more air-flow, intracavitary radially from the turbine disk
Flowing outward obturages gas for the cooling of the second tunnel.
More preferably, the second tunnel cooling obturages air-flow under the guide functions of the arc structure of the turbine disk, along turbine disk circular arc
The tangential entrance of structure is obturaged in ring cavity.
More preferably, the first via obturages air-flow and to obturage angle between air-flow be acute angle on the second tunnel, so as to obturage ring
Biggish eddy structure is smoothly formed in cavity, and under the action of the turbine disk, which has comparable tangential speed
Degree.
More preferably, the eddy structure with suitable tangential velocity in ring cavity is being obturaged as the first stage property of the invention
There is the seal structure of self-locking effect, can effectively prevent the leakage of a large amount of cooling tolerances, and prevent combustion gas from flowing backward.
More preferably, air-flow is obturaged by adjusting the first via to obturage the flow relative size of air-flow with the second tunnel and flow to phase
To angle, the whirlpool intensity obturaged in ring can be adjusted, obturage effect so as to adjust first seal structure.
More preferably, it is fixed on and obturages on ring by welding positioned at the honeycomb ring for obturaging ring outer ring.
Selectively, honeycomb ring can also be fixed in support ring by welding.
More preferably, the thickness of honeycomb ring should not be less than 2.5mm, it is ensured that when using welding manner fixed cellular, honeycomb ring
It good can be fixed on object part.And in engine operation process, mill is cut to pieces in case of blade and honeycomb ring
Situation, will not honeycomb ring is worn out and damaged blade, therefore the thickness of honeycomb ring should be by being calculated, it is ensured that meets its function
It can be reliable.
More preferably, the shape of honeycomb ring and the edge shape of working-blade mutually agree with, and are formed and are rotated with blade edge inner ring
Axial direction between part-static element obturages gap.The design mutually agreed with by honeycomb ring with blade edge, it is ensured that air-flow is along leaf
Piece edge tangentially successfully flows into mainstream channel, reduces disturb to mainstream bring as far as possible, reduces aerodynamic loss.
More preferably, which it is adjustable to obturage gap size, and when design is calculated by optimization.Its design principle is,
Under engine operating state, when each part leads to deformation under thermal stress or the effect of other active forces, the gap is as far as possible
Small, however, to ensure that the safe operation of engine, which cannot be zero.
Selectively, for example full and down operating or when excess revolutions under engine operation transitions state perhaps limiting condition, should
Gap can be zero with the short time.Since the material is soft for honeycomb ring, blade can rub with the honeycomb ring short time without damaging
Hurt blade.
More preferably, the reduction formed between honeycomb ring and blade edge, so the gap for being zero, become of the invention another
One of seal structure.
More preferably, in order to guarantee engine operational safety, blade and support ring or other high-intensitive material parts are prevented
Cut mill to pieces, the axial width of the honeycomb ring structure except blade edge should be greater than under various engine operating states,
Rate of travel between blade and honeycomb, it is ensured that under any engine operating condition, blade edge axial position is in bee always
Within the scope of the structure size of nest.
More preferably, in order to reduce aerodynamic loss, according to the position to axial between blade edge, honeycomb ring and support ring
Relationship is set, the maximum radius of downstream direction, three components should be sequentially reduced, it is ensured that air-flow will not encounter in flow process
Protrusion is by biggish disturbance.
More preferably, effect is obturaged for the self-locking seal structure of the optimizing regulation present invention, the quantity of twice seal structure can
It to be increased or be reduced, for example increases radially one and obturages ring, form three seal structures, or cancel and obturage ring,
Only retain honeycomb, effect is obturaged to adjust, to adapt to different gas turbine structure needs.
By using turbomachine axial self-locking seal structure, improve modern ground gas turbine and aero-engine
Effect is obturaged inside the turbomachine that equal needs are axially obturaged, compared with existing structure type, is had the advantage that
1) greatly enhance axial gap between turbomachine inner rotation part-static element and obturage effect, prevent combustion gas
Flow backward, gas consumption is obturaged in reduction, improves engine performance；
2) axially obturage using honeycomb ring structure, by the honeycomb characteristic that the material is soft, can be sent out with moving blade
The raw mill that touches is without damaged blade, so as to reduce axial gap to zero, in the premise for guaranteeing engine safety in operation
Under, it improves obturage effect as far as possible；
3) by using ring design is obturaged, the eddy structure formed by air-flow, which is formed, obturages knot with self-locking effect
Structure, and will not there are service lifes as other parts in engine, and according to the demand obturaged, can also increase and obturage
The quantity of ring, design concept is ingenious, and reliable operation；
4) labyrinth is not present in scheme of the invention, can design and blend with existing engine, be applied to rapidly
Have in engine structure, there is greatly application value and market prospects.
Detailed description of the invention
The schematic diagram of the self-locking seal structure of turbomachine Fig. 1 of the invention.
Specific embodiment
To make the objectives, technical solutions, and advantages of the present invention more comprehensible, the present invention is done below with reference to embodiment
Further to be described in detail, following embodiment is explanation of the invention and the invention is not limited to following embodiments.
Fig. 1 is the application example of the self-locking seal structure of turbomachine of the invention turbine disk back cavity in the gas turbine.Figure
In, 1 is honeycomb ring, and 2 are obturaged ring, and 3 be support ring, this three are static element.By welding or bolted mode,
It obturages ring 2 to be fixed in support ring 3, honeycomb ring 1 is fixed on by welding to be obturaged in ring 2 or support ring 3.4 be whirlpool
Impeller blade, 5 be the turbine disk, this two pieces is rotary part.By the tongue-and-groove 501 being arranged on the turbine disk 5, in turbo blade 4
After upper processing tenon, this two pieces can be assembled together.In order to guarantee the trouble-free operation of engine, rotating member blade 4 with
There are axial gap between static element support ring 3, by the axial gap, the high-temperature high-pressure fuel gas in mainstream channel is (in such as figure
Shown in solid arrow), it may leak by obturaging ring 2, support ring 3, in the turbine disk chamber 7 that the turbine disk 5 forms, threaten
The safe operation of engine.
The generation of such case in order to prevent, design has air system within the engine, and one of effect is to such
It is obturaged in gap.Gas (as indicated by a dashed arrow in the figure) of obturaging with certain pressure and flow is entered by the axial gap
In sprue, combustion gas can be prevented to flow backward.But due to various probabilistic limitations (such as the amount of obturaging can not be accurately pre-
Estimate, the inevitable property of process and assemble error), in order to ensure the safe operation of engine, so that such obturage often was in
The state obturaged seriously restricts the performance boost of engine.
In the present invention, ring 2 is obturaged by being arranged between the turbine disk 5 and adjacent support ring 3, and is obturaging ring 2
Outer ring surface honeycomb ring 1 is set, while abut the outer end face of honeycomb ring 1 with the preceding installation side of support ring 3, honeycomb ring 1 outside
The trailing edge shape of anchor ring and turbo blade 4 is mutually agreed with, and the trailing edge height of turbo blade 4 is higher than the height of the outer ring surface of honeycomb ring 1
Degree, the height of the outer ring surface of honeycomb ring 1 are not less than the height on the preceding installation side of support ring 3；And one is equipped on obturaging ring 2 and is opened
Mouthful towards the turbine disk revolution cavity 6, the axial assembly of turbo blade 4 and the turbine disk 5 engage the outlet direction time of backlash 501
Turn cavity 6；At the low radius of the assembly engagement tooth socket 501 of the turbine disk 5, design has concave arc structure 502, arc structure
The tangential extended line of 502 top ends is met in the revolution cavity 6 for obturaging ring.
Main air has phase by the fit-up gap between turbo blade 4 and the turbine disk 5 along axial direction in air system
When tangential velocity leak into the annular cavity 6 for obturaging ring 2.The secondary air flow of air system passes through 7 edge of turbine disk chamber
It flows radially outward, on the turbine disk at the low radius in 501 inside, is machined with concave arc flow-guiding structure 502, it can be by second
Stock is obturaged cooling air-flow and is imported in the annular cavity for obturaging ring 2 along certain angle, obturages cooling air-flow blending, shape with first strand
At the eddy structure 6 with tangential rotation speed, the eddy structure is highly stable, can it is good prevent combustion gas reverse irrigation enter
It invades.
Mutually agree in the trailing edge shape of the honeycomb 1 for obturaging the outside of ring 2, outer ring surface and blade, while by closing
The gap of calculation optimization is managed, when so that engine being run, the gap between honeycomb and blade is very small, or even in certain engines
Under non-design operating conditions, blade allows to scratch with honeycomb, and gap becomes zero.It, will not since the material is soft for honeycomb
Damage is generated to blade.And by reasonable design, cellular thickness can guarantee will not be worn out by blade, cellular axial direction
Width can guarantee to meet the play to axial between part.Pass through by the air-flow of obturaging in the air system of eddy structure
Gap between honeycomb and blade is approximately parallel to main flow direction and leaks into mainstream under the shape guidance of honeycomb external ring,
The interference to mainstream is reduced as far as possible.
In order to guarantee to reduce the aerodynamic loss of mainstream as far as possible, in axial direction, the height of blade trailing edge is higher than honeycomb
The height of outer ring, the height of honeycomb external ring is not less than the height for installing side before support ring.Thereby guarantee that sprue interior air-flow is flowing
The disturbance of boss will not be encountered during dynamic.
In addition, it should be noted that, the specific embodiments described in this specification, the shape of parts and components are named
Title etc. can be different.The equivalent or simple change that all structure, feature and principles described according to the invention patent design are done, is wrapped
It includes in the scope of protection of the patent of the present invention.Those skilled in the art can be to described specific implementation
Example is done various modifications or additions or is substituted in a similar manner, and without departing from structure of the invention or surmounts this
Range as defined in the claims, is within the scope of protection of the invention.
Claims (13)
1. a kind of self-locking seal structure of turbomachine, the turbomachine is gas turbine, including the turbine disk, turbo blade,
Ring, honeycomb ring and support ring are obturaged, ring, honeycomb ring, the support ring of obturaging is static element, and the turbo blade assembly is fixed
On the turbine disk, ring is obturaged between the turbine disk and adjacent support ring described in setting, it is described to obturage ring fixed setting
In the support ring, the honeycomb ring be fixed at it is described obturage in ring or support ring, the turbine disk and support ring,
It obturages ring and constitutes turbine disk chamber, which is characterized in that
Described obturage has axial gap between ring, honeycomb ring and the turbine disk,
The inner ring surface of the honeycomb ring is abutted with the outer ring surface for obturaging ring, the right side of the honeycomb ring and the support ring
Preceding installation side abut, the outer ring surface of the honeycomb ring mutually agrees with the trailing edge shape of the turbo blade, and the turbine leaf
The trailing edge bottom level of piece is higher than the height of the outer ring surface of the honeycomb ring, and the height of the outer ring surface of the honeycomb ring is not less than institute
State the height on the preceding installation side of support ring；
Described obturage is equipped with revolution cavity of the opening towards the turbine disk, the dress of the turbo blade and the turbine disk
Outlet with engagement backlash is towards the revolution cavity；
At the low radius of the assembly engagement tooth socket of the turbine disk, design has concave arc structure, the arc structure top
The tangential extended line of endpoint meets at described obturage in the revolution cavity of ring；
The turbine disk chamber is located at the inside of the inner ring surface for obturaging ring；
The cooling gas of obturaging of the first via in air system is by the assembly engaging tooth between the turbo blade and the turbine disk
Gap, leak into it is described obturage in the revolution cavity of ring,
The second tunnel in air system is cooling to obturage gas and radially flows the turbine disk is intracavitary, in the turbine
Under the guide functions of the indent arc structure of disk, along the arc structure it is tangential enter described in obturage ring revolution cavity in,
The first via is cooling to obturage gas and the second tunnel is cooling obturages after gas converges in the revolution cavity for obturaging ring, from
Axial gap between the honeycomb ring and the turbine disk enters mainstream channel.
2. self-locking seal structure according to claim 1, which is characterized in that the cross section for obturaging ring is hollow knot
Structure.
3. self-locking seal structure according to claim 1, which is characterized in that pass through between the turbo blade and the turbine disk
Tenon tooth or the assembly of dovetail tooth are fixed, and the revolution cavity of ring is obturaged described in the radial height position face of assembly engagement toothing.
4. self-locking seal structure according to claim 3, which is characterized in that engage tooth socket in the assembly of the turbine disk
At low radius, design has the indent arc structure, and the molded line of the indent arc structure is adjusted by optimization, top end
The tangential extended line of point meets at described obturage in the revolution cavity of ring.
5. self-locking seal structure according to any one of claims 1 to 4, which is characterized in that the second tunnel cooling is obturaged
The flow of gas is greater than the cooling flow for obturaging gas of the first via.
6. self-locking seal structure according to claim 5, which is characterized in that the second tunnel cooling obturages gas in the whirlpool
Under the indent arc structure guidance of wheel disc, ring is obturaged described in a manner of obturaging gas angle at an acute angle with first via cooling
Revolution cavity in converge blending and form eddy structure, form first of seal structure with self-locking effect.
7. self-locking seal structure according to claim 6, which is characterized in that described first obturaging with self-locking effect
Structure, the intensity of whirlpool and obturages effect, obturages gas and the second tunnel cools down and obturages the opposite of gas by adjusting the first via is cooling
The relative angle that flows to of tolerance size and two-way air-flow optimizes adjustment.
8. self-locking seal structure according to claim 1, which is characterized in that the honeycomb ring cross-sectional shape and the whirlpool
The edge of impeller blade agrees with, and the thickness of the honeycomb ring is not less than 2.5mm.
9. self-locking seal structure according to claim 8, which is characterized in that the edge of the turbo blade and honeycomb ring it
Between have axial gap, the axial gap is adjustable, and under engine operating state, which cannot be zero, is sending out
Under motivation transition state or limiting case, which is zero.
10. self-locking seal structure according to claim 9, which is characterized in that the honeycomb ring and turbo blade edge it
Between the axial gap that reduces under engine operating state, form another road seal structure.
11. self-locking seal structure according to claim 9 or 10, which is characterized in that be located at outside the turbo blade edge
The honeycomb ring of side, axial width are greater than the rate of travel between the turbo blade and honeycomb ring under engine different conditions.
12. self-locking seal structure according to claim 1, which is characterized in that the honeycomb ring is fixed by welding manner
Obturage that ring is perhaps described in support ring to be obturaged ring and be fixed on the support ring by welding or bolted mode described
On.
13. a kind of engine includes self-locking seal structure described in above-mentioned any one claim.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710344978.3A CN107131009B (en) | 2017-05-16 | 2017-05-16 | A kind of self-locking seal structure of turbomachine and the engine with it |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710344978.3A CN107131009B (en) | 2017-05-16 | 2017-05-16 | A kind of self-locking seal structure of turbomachine and the engine with it |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN107131009A CN107131009A (en) | 2017-09-05 |
| CN107131009B true CN107131009B (en) | 2019-02-15 |
Family
ID=59732399
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201710344978.3A Active CN107131009B (en) | 2017-05-16 | 2017-05-16 | A kind of self-locking seal structure of turbomachine and the engine with it |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN107131009B (en) |
Families Citing this family (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109356660B (en) * | 2018-12-14 | 2021-11-19 | 中国航发沈阳发动机研究所 | Double-stage high-pressure turbine rotor-stator assembly |
| CN112922681A (en) * | 2021-03-23 | 2021-06-08 | 中国航发沈阳发动机研究所 | Aeroengine rim sealing structure |
| CN113719355B (en) * | 2021-08-24 | 2022-08-12 | 中国航发湖南动力机械研究所 | Sealing structure between breather pipe and bearing seat |
| CN115142907B (en) * | 2022-09-02 | 2022-11-22 | 中国航发沈阳发动机研究所 | Integrated structure of guide vane inner ring of aero-engine |
| CN116399526B (en) * | 2023-06-05 | 2023-09-01 | 中国航发四川燃气涡轮研究院 | Circumferential sealing effect verification device for guide vane of aero-engine |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH05214963A (en) * | 1991-10-30 | 1993-08-24 | General Electric Co <Ge> | Forward seal assembly |
| JP2005530956A (en) * | 2002-06-27 | 2005-10-13 | スネクマ・モトウール | Gas turbine ventilation circuit |
| CN105402028A (en) * | 2015-04-30 | 2016-03-16 | 中国科学院工程热物理研究所 | Rib plate eddy flow control structure, rotation disk cavity system and gas turbine |
| CN106089312A (en) * | 2016-07-29 | 2016-11-09 | 中国科学院工程热物理研究所 | A kind of have the turbine disk cavity configuration obturaging and cooling down deflector |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6481959B1 (en) * | 2001-04-26 | 2002-11-19 | Honeywell International, Inc. | Gas turbine disk cavity ingestion inhibitor |
-
2017
- 2017-05-16 CN CN201710344978.3A patent/CN107131009B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH05214963A (en) * | 1991-10-30 | 1993-08-24 | General Electric Co <Ge> | Forward seal assembly |
| JP2005530956A (en) * | 2002-06-27 | 2005-10-13 | スネクマ・モトウール | Gas turbine ventilation circuit |
| CN105402028A (en) * | 2015-04-30 | 2016-03-16 | 中国科学院工程热物理研究所 | Rib plate eddy flow control structure, rotation disk cavity system and gas turbine |
| CN106089312A (en) * | 2016-07-29 | 2016-11-09 | 中国科学院工程热物理研究所 | A kind of have the turbine disk cavity configuration obturaging and cooling down deflector |
Also Published As
| Publication number | Publication date |
|---|---|
| CN107131009A (en) | 2017-09-05 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN107131009B (en) | A kind of self-locking seal structure of turbomachine and the engine with it | |
| US11193380B2 (en) | Integrated strut-vane | |
| US10415505B2 (en) | Non-axisymmetric fan flow path | |
| US8066471B2 (en) | Annular flow duct for a turbomachine through which a main flow can flow in the axial direction | |
| CN101131096A (en) | Flared tip turbine blade | |
| CN103742450B (en) | A kind of blade diffuser with horse shoe shaped diffusion passage | |
| JP5415532B2 (en) | Air injection into the turbomachine compressor flow path | |
| JP2016109124A (en) | Axial compressor endwall treatment for controlling leakage flow | |
| US8197195B2 (en) | Turbocharger with stepped two-stage vane nozzle | |
| JP5651459B2 (en) | System and apparatus for compressor operation in a turbine engine | |
| JP2012087928A (en) | Turbomachine seal assembly | |
| US10954953B2 (en) | Rotor hub seal | |
| JP2011094616A (en) | Turbine airfoil-sidewall integration | |
| CN110630339A (en) | Turbine disc with disc edge sealing structure | |
| CN107100677A (en) | A kind of nozzle ring assemblies of fixed blade and adjustable vane combination | |
| US20130017059A1 (en) | Hole for rotating component cooling system | |
| US20160003082A1 (en) | Contoured blade outer air seal for a gas turbine engine | |
| JPH09317696A (en) | Stator blade structure of axial flow compressor | |
| US10900414B2 (en) | Fan assembly having flow recirculation circuit with guide vanes | |
| CN210289847U (en) | Turbine and exhaust gas turbocharger | |
| US10962021B2 (en) | Non-axisymmetric impeller hub flowpath | |
| JPS5951104A (en) | Internal structure of turbine stage | |
| CN206668332U (en) | A kind of nozzle ring assemblies of fixed blade and adjustable vane combination | |
| CN104718361B (en) | Turbocharger, turbine nozzle and boats and ships | |
| US11499475B2 (en) | Fan assembly having flow recirculation circuit with rotating airfoils |
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 | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |