CN107448240A - Core engine and turbogenerator - Google Patents
Core engine and turbogenerator Download PDFInfo
- Publication number
- CN107448240A CN107448240A CN201610374373.4A CN201610374373A CN107448240A CN 107448240 A CN107448240 A CN 107448240A CN 201610374373 A CN201610374373 A CN 201610374373A CN 107448240 A CN107448240 A CN 107448240A
- Authority
- CN
- China
- Prior art keywords
- turbine
- final stage
- compressor
- cooling
- compressed gas
- 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.)
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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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
Abstract
The present invention relates to a kind of core engine and turbogenerator, wherein, turbogenerator includes：Compressor, combustion chamber and turbine, turbine includes turbine one-level movable vane, turbine one-level movable vane has the preceding cooling chamber close to combustion chamber and the rear cooling chamber away from combustion chamber, preceding cooling chamber and rear cooling chamber are set independently of one another, core engine also includes the first cooling flowing path and the second cooling flowing path, the final stage compressed gas of compressor final stage can be incorporated into preceding cooling chamber by the first cooling flowing path, and the non-final stage compressed gas of the non-final stage of compressor can be incorporated into rear cooling chamber by the second cooling flowing path.Core engine of the present invention is cooled down by distinguishing bleed from two diverse locations of compressor for turbine one-level movable vane, final stage compressed gas and non-final stage compressed gas are provided commonly for the cooling of turbine one-level movable vane, total flow needed for turbine one-level movable vane cold air is reduced on the whole, engine thermal efficiency is improved, reduces oil consumption rate.
Description
Technical field
The present invention relates to turbogenerator technical field, more particularly to a kind of core engine and turbogenerator.
Background technology
In order to improve the civil aviation fanjet thermal efficiency, oil consumption rate is reduced, inlet temperature constantly improves before turbine,
The problem of bringing therewith is that cold air also rises needed for high-temperature component, especially dual stage turbo.
At present, dual stage turbo first order movable vane cold air is used from high-pressure compressor final stage bleed, is entered after preswirl nozzle
First order movable vane piece, after the abundant heat convection of first order movable vane piece inner cooling path, finally from first order movable vane piece ante-chamber and tail
The hole of edge is split seam and is discharged into sprue.
Following two technological deficiencies at least be present in existing tradition supply and blade cooling design method：
1st, not being avoided that from high-pressure compressor final stage bleed, the bleed of high-pressure compressor final stage reduces engine thermal efficiency；
2nd, higher from the cold air temperature of high-pressure compressor final stage bleed, required air conditioning quantity is big, reduce further engine
The thermal efficiency.
The content of the invention
To overcome above technological deficiency, present invention solves the technical problem that it is to provide a kind of core engine and turbogenerator,
Engine thermal efficiency can be improved.
In order to solve the above technical problems, the invention provides a kind of core engine, it includes：Compressor, combustion chamber and whirlpool
Wheel, turbine include turbine one-level movable vane, and turbine one-level movable vane has close to the preceding cooling chamber of combustion chamber and away from behind combustion chamber
Cooling chamber, preceding cooling chamber and rear cooling chamber are set independently of one another, and core engine also includes the first cooling flowing path and the second cooling flowing path,
The final stage compressed gas of compressor final stage can be incorporated into preceding cooling chamber by the first cooling flowing path, and the second cooling flowing path will can press
The non-final stage compressed gas of the non-final stage of mechanism of qi is incorporated into rear cooling chamber.
Further, turbine also includes the turbine two level movable vane with cooling chamber, and the second cooling flowing path can be by one
The non-final stage compressed gas divided is incorporated into rear cooling chamber, and the non-final stage compressed gas of another part is incorporated into cooling chamber.
Further, it is logical to be provided with first communicated respectively with preceding cooling chamber and rear cooling chamber for the blade root of turbine one-level movable vane
Final stage compressed gas is incorporated into preceding cooling chamber, the second cooling stream by road and second channel, the first cooling flowing path by first passage
Non- final stage compressed gas is incorporated into rear cooling chamber by road by second channel.
Further, the tenon portion of turbine one-level movable vane, which is provided with, is used to isolate final stage compressed gas and non-final stage compressed gas
Boss dividing plate.
Further, the second cooling flowing path is partially disposed in the core passage of core engine.
Further, compressor is high-pressure compressor, and turbine is high-pressure turbine, and core passage includes compressor and low-pressure shaft
The compressor core passage of formation, the passage that the high-pressure shaft of core engine and low-pressure shaft are formed and turbine turbine one-level disk and
The turbine core passage that low-pressure shaft is formed, the second cooling flowing path include disk chamber, the compressor disc between the non-final stage disk of compressor
Heart passage, passage and turbine core passage, the drum barrel of compressor are provided with air admission hole, and non-final stage compressed gas is via air admission hole
Into disk chamber.
Further, disk intracavitary, which is provided with, subtracts whirlpool device, and the non-final stage compressed gas into disk intracavitary is entered by subtracting whirlpool device
Enter compressor core passage.
Further, the first cooling flowing path is partially disposed in the combustion chamber inner ring in combustion chamber, and core engine also includes
Preswirl nozzle, receiver hole and the impeller of booster being arranged between combustion chamber and preceding cooling chamber, final stage compressed gas by preswirl nozzle,
Receiver hole and impeller of booster enter preceding cooling chamber.
Further, it is provided with supercharging device on rear end face of the turbine one-level disk away from combustion chamber of turbine.
Present invention also offers a kind of turbogenerator, and it includes above-mentioned core engine.
Thus, used based on above-mentioned technical proposal, core engine of the present invention by distinguishing bleed from two diverse locations of compressor
Cooled down in turbine one-level movable vane, because turbine one-level movable vane cold air is typically discharged from blade inlet edge and trailing edge, needed for leading edge exhaust
Cold air supply gas pressure is big, and by setting the first cooling flowing path, cooling gas is quoted from compressor final stage；And pressure needed for trailing edge exhaust
It is small, by setting non-final stage of second cooling flowing path by cooling gas quoted from compressor.Final stage compressed gas and the compression of non-final stage
Gas is provided commonly for the cooling of turbine one-level movable vane, reduces the amount of air entrainment of the final stage compressed gas of compressor final stage；And from calming the anger
The temperature of the non-final stage compressed gas of the non-final stage of machine is relatively low, and flow is small, thus reduces turbine one-level movable vane cold air institute on the whole
Total flow is needed, so as to improve engine thermal efficiency, reduces oil consumption rate.Turbogenerator provided by the invention correspondingly also has upper
State beneficial effect.
Brief description of the drawings
Accompanying drawing described herein is used for providing a further understanding of the present invention, forms the part of the application, this hair
Bright schematic description and description is only used for explaining the present invention, does not form inappropriate limitation of the present invention.In the accompanying drawings：
Fig. 1 is the overall structure diagram of core engine embodiment of the present invention；
Fig. 2 is the structural representation of the turbine portion of core engine embodiment of the present invention；
Fig. 3 is the structural representation of the blade root of turbine one-level movable vane in core engine of the present invention；
Fig. 4 is the cross-sectional view of turbine one-level movable vane blade root in core engine of the present invention.
Embodiment
Below by drawings and examples, technical scheme is described in further detail.
The embodiment of the present invention is for the ease of the design to the present invention, the technical problem solved, forms skill
The technical characteristic of art scheme and the technique effect brought have further description.It should be noted that for these embodiment party
The explanation of formula does not form limitation of the invention.In addition, the technical characteristic being related in following embodiments of the present invention is only
Conflict can is not formed each other to be mutually combined.
Final stage bleed cooling for current existing turbogenerator can reduce engine thermal efficiency, and the present invention devises
A kind of core engine, the core engine are moved by distinguishing bleed from two diverse locations of compressor final stage and non-final stage for turbine one-level
Leaf cools down, and final stage compressed gas and non-final stage compressed gas are provided commonly for the cooling of turbine one-level movable vane, reduce compressor final stage
Final stage compressed gas amount of air entrainment；And the temperature of the non-final stage compressed gas from the non-final stage of compressor is relatively low, flow is small, thus
Total flow needed for turbine one-level movable vane cold air is reduced on the whole, so as to improve engine thermal efficiency, reduces oil consumption rate.
In the schematical embodiment of core engine one of the present invention, as shown in Figure 1 to 4, core engine includes：Compressor 1,
Combustion chamber 2 and turbine, turbine include turbine one-level movable vane 3, and turbine one-level movable vane 3 has the preceding cooling chamber close to combustion chamber 2
21 and the rear cooling chamber 20 away from combustion chamber 2, preceding cooling chamber 21 and rear cooling chamber 20 set independently of one another, and air enters compressor
1 and by compressor 1 compress after, burnt into combustion chamber 2, from combustion chamber 2 come out combustion gas drive turbine one-level movable vane 3 again.
Core engine also includes the first cooling flowing path and the second cooling flowing path, and the first cooling flowing path can be by the final stage pressure of the final stage of compressor 1
Contracting gas B is incorporated into preceding cooling chamber 21, and the second cooling flowing path can introduce the non-final stage compressed gas A of 1 non-final stage of compressor
To rear cooling chamber 20.
In the schematical embodiment, because turbine one-level movable vane cold air is typically discharged from blade inlet edge and trailing edge, lead to
Cross and the first cooling flowing path and the second cooling flowing path are set, in the presence of the centrifugation pump work of turbine one-level disk 11, the first cooling stream
Cooling chamber 21 before the final stage compressed gas B of compressor final stage is incorporated into by road, is finally discharged into combustion gas from the leading edge of turbine one-level movable vane 3
Mainstream channel；And pressure needed for trailing edge exhaust is small, the second cooling flowing path draws the non-final stage compressed gas A of 1 non-final stage of compressor
Enter to rear cooling chamber 20, finally combustion gas mainstream channel is discharged into from the trailing edge of turbine one-level movable vane 3, subsequently into back segment region 5.Due to
Preceding cooling chamber 21 and rear cooling chamber 20 are set independently of one another, final stage compressed gas and non-final stage pressure into turbine one-level movable vane 3
Contracting gas will not mutual play.Final stage compressed gas B and non-final stage compressed gas A is provided commonly for turbine one-level movable vane 3 and cooled down, drop
The low final stage compressed gas B of the final stage of compressor 1 amount of air entrainment；And the temperature of the non-final stage compressed gas A from 1 non-final stage of compressor
Spend relatively low, flow is small, thus reduces total flow needed for the cold air of turbine one-level movable vane 3 on the whole, so as to improve engine thermal effect
Rate, reduce oil consumption rate.
Wherein, turbine one-level movable vane 3 be provided with the leading edge Cooling Holes that are communicated respectively with preceding cooling chamber 21 and rear cooling chamber 20 and
Trailing edge Cooling Holes, leading edge Cooling Holes and trailing edge Cooling Holes preferably enter preceding cold in the structure type for splitting seam, final stage compressed gas B
But the runner for entering turbine by leading edge Cooling Holes is passed through after chamber 21, non-final stage compressed gas A is passed through by trailing edge after entering rear cooling chamber 20
Cooling Holes enter the runner of turbine.
It should be noted that non-final stage compressed gas A bleed position can be compressor 1 intergrade or
Compressor it is preceding what, it is not limited to which level, as long as supply gas pressure is enough, draw as much as possible from the forward level position of compressor
Gas.
As the improvement to above-described embodiment, as depicted in figs. 1 and 2, turbine also includes the turbine with cooling chamber 22
The non-final stage compressed gas A of a part, i.e., non-final stage compressed gas Aa can be incorporated into by two level movable vane 4, the second cooling flowing path
Cooling chamber 20 afterwards, and the non-final stage compressed gas A of another part, i.e., non-final stage compressed gas Ab are incorporated into cooling chamber 22,
So non-final stage compressed gas can also cool down to turbine two level movable vane 4, be set compared to turbine two level movable vane 4 is individually for
Cooling flowing path simplifies structure, improves stream utilization rate, meets existing light-weight design needs.Certainly, non-final stage compression
Gas A can also be served only for cooling down turbine one-level movable vane 3, and to turbine two level movable vane 4 cooling can from compressor 1 its
His position bleed, it is numerous to list herein.
Specifically, in a preferred embodiment, as shown in figure 4, the blade root 23 of turbine one-level movable vane 3 be provided with point
The first passage 25,26 and second channel 27,28 not communicated with preceding cooling chamber 21 and rear cooling chamber 20, the first cooling flowing path passes through
Cooling chamber 21 before final stage compressed gas B is incorporated into by first passage 25,26, the second cooling flowing path will by second channel 27,28
Non- final stage compressed gas A is incorporated into rear cooling chamber 20.By setting first passage 25,26 and second channel 27,28, final stage compression
Gas B and non-final stage compressed gas A carries out cold from the bottom of the blade root 23 of turbine one-level movable vane 3 into turbine one-level movable vane 3
But, compared to the mode from engine stator part air inlet, then from disk chamber supply, reliability and stability is more preferable, and cooling effectiveness is more
It is high.It is further preferred that in order to prevent from entering the final stage compressed gas B of turbine one-level movable vane 3 and non-final stage pressure from blade root bottom
The mutual plays of contracting gas A and influence to cool down and be vented, as shown in figure 4, the tenon portion of turbine one-level movable vane 3 be provided be used for isolate end
Level compressed gas B and non-final stage compressed gas A boss dividing plate 24.
For the set-up mode of the second cooling flowing path, in a preferred embodiment, the second cooling flowing path is partly set
Put in the core passage of core engine, can so avoid using core engine exterior line bleed, without setting Puncture Tube
Road, save space-consuming.
In one particularly preferred embodiment, as shown in figure 1, compressor 1 is high-pressure compressor, turbine is high pressure whirlpool
Wheel, back segment region 5 is high and low pressure turbine changeover portion.Core passage includes the compressor core that compressor 1 is formed with low-pressure shaft 7 and led to
The passage 15 and the turbine one-level disk 11 and low-pressure shaft 7 of turbine that road 14, the high-pressure shaft 6 of core engine and low-pressure shaft 7 are formed are formed
Turbine core passage 16, disk chamber 13, compressor core passage between the non-final stage disk of the second cooling flowing path including compressor 1
14th, passage 15 and turbine core passage 16, the drum barrel 8 of compressor 1 are provided with air admission hole, and non-final stage compressed gas A is via entering
Stomata enters disk chamber 13.By taking core engine structure as shown in Figure 1 as an example, opening position of the drum barrel 8 between compressor disc 9 and 10 is set
Air admission hole is put, the disk chamber 13 that non-final stage compressed gas A enters between non-final stage disk compressor disc 9 and 10 from air admission hole, passes through pressure
What compressor core passage 14, passage 15 and core passage 16 after mechanism of qi, reach turbine one-level disk 11 and two level disk 12
The twin-stage high-pressure turbine disk chamber 17 and 18 of formation, it is preferable that on rear end face of the turbine one-level disk 11 away from combustion chamber 2 of turbine
Provided with supercharging device, that is to say, that supercharging device is set in high-pressure turbine disk chamber 17, to enter the non-final stage pressure of rear cooling chamber 20
Contracting gas Aa provides preferably centrifugation pump work.
As illustrated in fig. 1 and 2, a part of non-final stage compressed gas Aa is under the centrifugation pump work of turbine one-level disk 11, into whirlpool
Cooling chamber 20 after wheel one-level movable vane, is finally discharged into combustion gas mainstream channel from the trailing edge of turbine one-level movable vane 3；The non-final stage of another part
Compressed gas Ab, under the centrifugation pump work of turbine two level disk 12, into the cooling chamber 22 of turbine two level movable vane 4, it is ultimately discharged into
Combustion gas mainstream channel.
It is provided with a further preferred embodiment, in disk chamber 13 and subtracts whirlpool device, into the non-final stage in disk chamber 13
Compressed gas A enters compressor core passage 14 by subtracting whirlpool device.Non- final stage compressed gas A can be reduced by subtracting whirlpool device
The pressure loss in one cooling flowing path, ensure cooling effectiveness and flow, it is preferably that tubular type subtracts whirlpool device to subtract whirlpool device, also may be used certainly
Subtract whirlpool device accordingly not designed according to actual parameter.
For the set-up mode of the first cooling flowing path, in a preferred embodiment, as illustrated in fig. 1 and 2, first cools down
Circuit portion it is arranged in the combustion chamber inner ring 19 in combustion chamber 2, core engine also includes being arranged on combustion chamber 2 and preceding cooling chamber
Preswirl nozzle, receiver hole and impeller of booster between 21, final stage compressed gas B are entered by preswirl nozzle, receiver hole and impeller of booster
Cooling chamber 21 before entering, so as to enter preceding cooling chamber 21 with ensureing the loss of final stage compressed gas B no pressures, have higher reliable steady
It is qualitative.
Present invention also offers a kind of turbogenerator, and it includes above-mentioned core engine.Because core engine of the present invention can
Engine thermal efficiency is improved, correspondingly, turbogenerator of the present invention also has above-mentioned advantageous effects, no longer superfluous herein
State.
Above in association with embodiment be described in detail for embodiments of the present invention, but the present invention is not limited to be retouched
The embodiment stated.For a person skilled in the art, in the case where not departing from the principle of the present invention and connotation
These embodiments are carried out with a variety of changes, modification, equivalence replacement and modification to still fall within protection scope of the present invention.
Claims (10)
- A kind of 1. core engine, it is characterised in that including：Compressor (1), combustion chamber (2) and turbine, the turbine include turbine One-level movable vane (3), the turbine one-level movable vane (3) have close to the preceding cooling chamber (21) of the combustion chamber (2) and away from described The rear cooling chamber (20) of combustion chamber (2), the preceding cooling chamber (21) and the rear cooling chamber (20) are set independently of one another, the core Scheming also includes the first cooling flowing path and the second cooling flowing path, and first cooling flowing path can be by the compressor (1) final stage Final stage compressed gas (B) be incorporated into the preceding cooling chamber (21), second cooling flowing path can be non-by the compressor (1) The non-final stage compressed gas (A) of final stage is incorporated into the rear cooling chamber (20).
- 2. core engine according to claim 1, it is characterised in that the turbine also includes the whirlpool with cooling chamber (22) Take turns two level movable vane (4), second cooling flowing path can by a part the non-final stage compressed gas (A) be incorporated into it is described after Cooling chamber (20), and the non-final stage compressed gas (A) of another part is incorporated into the cooling chamber (22).
- 3. core engine according to claim 1, it is characterised in that the blade root (23) of the turbine one-level movable vane (3) is set Have respectively with the preceding cooling chamber (21) and it is described after the first passage (25,26) that communicates of cooling chamber (20) and second channel (27, 28), first cooling flowing path by the first passage (25,26) by the final stage compressed gas (B) be incorporated into it is described before Cooling chamber (21), second cooling flowing path are introduced the non-final stage compressed gas (A) by the second channel (27,28) To the rear cooling chamber (20).
- 4. core engine according to claim 3, it is characterised in that the tenon portion of the turbine one-level movable vane (3), which is provided with, to be used In the boss dividing plate (24) for isolating the final stage compressed gas (B) and the non-final stage compressed gas (A).
- 5. core engine according to claim 1, it is characterised in that second cooling flowing path is partially disposed at the core In the core passage of scheming.
- 6. core engine according to claim 5, it is characterised in that the compressor (1) is high-pressure compressor, the turbine For high-pressure turbine, the core passage include the compressor core passage (14) that the compressor (1) formed with low-pressure shaft (7), The passage (15) and the turbine one-level disk of the turbine that the high-pressure shaft (6) and the low-pressure shaft (7) of the core engine are formed (11) and the turbine core passage (16) that is formed of the low-pressure shaft (7), second cooling flowing path include the compressor (1) Disk chamber (13), the compressor core passage (14), the passage (15) and the turbine core between non-final stage disk lead to Road (16), the drum barrel (8) of the compressor (1) are provided with air admission hole, and the non-final stage compressed gas (A) is via the air admission hole Into the disk chamber (13).
- 7. core engine according to claim 6, it is characterised in that be provided with the disk chamber (13) and subtract whirlpool device, into institute The non-final stage compressed gas (A) stated in disk chamber (13) enters the compressor core passage by the whirlpool device that subtracts (14)。
- 8. core engine according to claim 1, it is characterised in that first cooling flowing path is partially disposed at the combustion Burn in the combustion chamber inner ring (19) in room (2), the core engine also includes being arranged on the combustion chamber (2) and the preceding cooling chamber (21) preswirl nozzle, receiver hole and impeller of booster between, the final stage compressed gas (B) by the preswirl nozzle, described connect Batter and the impeller of booster enter the preceding cooling chamber (21).
- 9. core engine according to claim 5, it is characterised in that remote described in the turbine one-level disk (11) of the turbine The rear end face of combustion chamber (2) is provided with supercharging device.
- 10. a kind of turbogenerator, it is characterised in that including the core engine described in any one of claim 1~9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610374373.4A CN107448240A (en) | 2016-05-31 | 2016-05-31 | Core engine and turbogenerator |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610374373.4A CN107448240A (en) | 2016-05-31 | 2016-05-31 | Core engine and turbogenerator |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN107448240A true CN107448240A (en) | 2017-12-08 |
Family
ID=60485131
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201610374373.4A Pending CN107448240A (en) | 2016-05-31 | 2016-05-31 | Core engine and turbogenerator |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN107448240A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110359971A (en) * | 2018-03-26 | 2019-10-22 | 中国航发商用航空发动机有限责任公司 | Aero-turbine movable vane cools down air supply system |
| CN113027609A (en) * | 2021-03-23 | 2021-06-25 | 天津鱼羊文化传播有限公司 | Turbofan engine |
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|---|---|---|---|---|
| US20030079478A1 (en) * | 2001-10-26 | 2003-05-01 | Giuseppe Romani | High pressure turbine blade cooling scoop |
| US20070137221A1 (en) * | 2005-10-21 | 2007-06-21 | Snecma | Device for ventilating turbine disks in a gas turbine engine |
| US20080112794A1 (en) * | 2006-11-10 | 2008-05-15 | General Electric Company | Compound nozzle cooled engine |
| US20080112791A1 (en) * | 2006-11-10 | 2008-05-15 | General Electric Company | Compound turbine cooled engine |
| US20100284799A1 (en) * | 2009-05-07 | 2010-11-11 | Ian David Wilson | Method and apparatus for turbine engines |
-
2016
- 2016-05-31 CN CN201610374373.4A patent/CN107448240A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20030079478A1 (en) * | 2001-10-26 | 2003-05-01 | Giuseppe Romani | High pressure turbine blade cooling scoop |
| US20070137221A1 (en) * | 2005-10-21 | 2007-06-21 | Snecma | Device for ventilating turbine disks in a gas turbine engine |
| US20080112794A1 (en) * | 2006-11-10 | 2008-05-15 | General Electric Company | Compound nozzle cooled engine |
| US20080112791A1 (en) * | 2006-11-10 | 2008-05-15 | General Electric Company | Compound turbine cooled engine |
| US20100284799A1 (en) * | 2009-05-07 | 2010-11-11 | Ian David Wilson | Method and apparatus for turbine engines |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110359971A (en) * | 2018-03-26 | 2019-10-22 | 中国航发商用航空发动机有限责任公司 | Aero-turbine movable vane cools down air supply system |
| CN110359971B (en) * | 2018-03-26 | 2022-03-25 | 中国航发商用航空发动机有限责任公司 | Aircraft engine turbine bucket cooling air supply system |
| CN113027609A (en) * | 2021-03-23 | 2021-06-25 | 天津鱼羊文化传播有限公司 | Turbofan engine |
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Application publication date: 20171208 |