CN1204005A - Jet engine - Google Patents
Jet engine Download PDFInfo
- Publication number
- CN1204005A CN1204005A CN 98115075 CN98115075A CN1204005A CN 1204005 A CN1204005 A CN 1204005A CN 98115075 CN98115075 CN 98115075 CN 98115075 A CN98115075 A CN 98115075A CN 1204005 A CN1204005 A CN 1204005A
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- firing chamber
- jet engine
- downstream
- turbo machine
- jet
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Abstract
In a jet engine having a fan (9) of conventional type, a second combustion chamber (5) is provided downstream of the turbine (4), and a second turbine (6) acts downstream of this second combustion chamber (5). The hot gases (16) partly expanded in the first turbine (4) are brought up to temperature again in the second combustion chamber (5), whereby, at a low absolute hot-gas temperature, more output is then available in the engine.
Description
The present invention relates to a kind of jet engine.
Modern jet engine has single shaft or multiaxis is arranged.Usually, gas turbine group is divided into a high-pressure section and a low-pressure section.High-pressure section comprises a high pressure compressor, a combustion chamber unit and a high pressure turbine.The compressor of high-pressure section and turbo machine are fixed on one and are total on the hollow shaft.One low pressure compressor reaches, if a bypass flow motor then is the upstream that a propeller is disposed at high pressure compressor.And then one low-pressure turbine is arranged in after the high pressure turbine.Propeller, low pressure compressor and low-pressure turbine have formed the low-pressure section of motor.This low-pressure section can be arranged in the public altogether rotating shaft with high-pressure section.But in the high-performance enginer in the present age, high-pressure section has independently second rotating shaft usually, and sometimes even have one the 3rd rotating shaft, these axles are with the coaxial installation of high-pressure shaft and rotate in the high pressure rotating shaft.Like this, just can realize different rotational velocity between low pressure parts and the high voltage component.
A common feature of the structure of all these forms is that they only have an independent firing chamber in the downstream of compressor bank.
The highest permission combustion temperature in this combustion chamber unit is subjected to technologic restriction.This combustion temperature was lasting in recent years improves.In order to reach pollutant effulent level that law allows and the combustion technology that adopted in the combustion chamber unit is combined with material with cooling technology on the gas turbine, this combination has limited this temperature.Therefore, correspondingly control predetermined mass flow rate at the entry of combustion chamber place by engine combustion process, the energy that fuel can provide is subjected to the restriction of the highest permission combustion temperature.As a result, for having the motor that a given mass flow is distributed, its total output power also is restricted.
Therefore, one object of the present invention is exactly, in the jet engine of described form at a first brush, by proposing some novel methods, expand above-mentioned restriction, in the hope of be issued to higher power output in the same process level to output power by the technology decision.
A major advantage of the present invention at first is motor of the present invention, under the identical prerequisite of output power, compares with prior art state motor, and fuel consumption rate significantly reduces, and therefore not only efficient has improved, and the unit pollutant effulent has also obviously reduced.
Another major advantage of the present invention is relevant with the ratio of bypass flow and inner stream flow: the fuel consumption rate of high-engine will be low more more for this ratio.For this reason, must utilize the output power of propeller as far as possible, this output power is converted into thrust according to the bypass flow ratio.
Other advantages of the present invention mainly are: the invention provides one second firing chamber, the hot gas of demi-inflation is heated in this second firing chamber again in first turbo machine, therefore under the highest identical hot air temperature, motor of the present invention can provide higher output power.
The present invention seeks to by adopting following useful and suitable improvement to reach, they comprise:
Propeller is equipped with adjustable vane, and is multilevel hierarchy; Belong at least one compressor of described compressor unit, described first turbo machine and described second turbine arrangement in a public rotating shaft; The mobile machinery of all fluids of this jet engine is arranged in two rotating shafts that operation links to each other each other at least; Being operatively connected of rotating shaft can be reached by at least one speed change gear; Second Combustion chamber design be one from the ignition combustion chamber; Upstream arrangement in second firing chamber has a diffuser; Second firing chamber is equipped with vortex generating component; The side surface that goes out to flow of vortex generating component has a sudden change on cross section; One main fuel can be sprayed into described second firing chamber in described vortex generating component downstream; Main fuel is imported by the jet pipe of some, and these jet pipes are arranged on the peripheral direction of described second firing chamber; First firing chamber can be operated with some Premix burners; First and second Combustion chamber design are annular combustion chamber; First and/or second firing chamber comprises some independent cylinder combustion spaces around described rotating shaft layout; The expansion of the hot gas in first turbo machine is reduced to minimum, and the hot gas of demi-inflation flows into second firing chamber in downstream with the temperature higher than the spontaneous ignition temperature of the main fuel that sprays into described second firing chamber; At least in the transient load scope with in the partial load scope, second firing chamber has ancillary method or auxiliary device, is guaranteed by this measure or device burning.
By with reference to below in conjunction with the description of accompanying drawing, the many advantages that can more completely understand the present invention and be had like a dream just, a wherein independent width of cloth accompanying drawing shows the sectional elevation of the air breathing engine among the bypass flow embodiment.
In the accompanying drawings, all are all saved to directly understanding the unnecessary member of the present invention.Arrow shows the flow direction of medium gas among the figure.
Existing the characteristics of motor are that it only has an independent rotating shaft 7 as shown in the figure referring to accompanying drawing, and air inlet 12 is compressed in a propeller 9, a low pressure compressor 1 and high pressure compressor 2 subsequently in this motor.Afterwards, compressed air 13 flows into first firing chamber 3, reaches the maximum temperature that allows from technological standpoint by supplying with appropriate liquid fuel 14 in this firing chamber.Hot gas 15 demi-inflation in first turbo machine 4 that in this firing chamber 13, generates, and output power passed to rotating shaft 7.The hot gas 16 of this demi-inflation leaves this first turbo machine 4 with a lower temperature, flows into second firing chamber 5 afterwards, and these hot gas 16 are heated to maximum temperature for the second time in this firing chamber 5.Also will further set forth the combustion technology in first and second firing chambers below.After the gas in second firing chamber 5 was heated to maximum temperature once more, they promptly expanded in second turbo machine 6 in downstream, and correspondingly output power are passed to rotating shaft 7.The final gas 17 that expands also is exhaust, discharges motor by a thrust nozzle (not being shown specifically in the accompanying drawings).The high pressure compressor 2 that the part of shaft power is connected in the rotating shaft 7 consumes.Therefore in view of the rotating speed of this rotating shaft is too high for propeller 9,, this rotating speed is decreased to the degree that propeller can be controlled by machinery reliably with a speed change gear 8 for consideration to the latter.The air mass flow 12 that enters motor is compressed by propeller 9 in the first order; Then this mass flow is divided into an inner stream and a bypass flow.Expand in bypass flow afterwards the is through-flow aside nozzle 10, the gross thrust of motor is played main contribution.
Bypass flow is big more with the ratio of inner stream, and the fuel consumption rate of motor will be low more.Given this, improve the available output power of propeller 9 as much as possible, this point is very important, to continue to realize a high bypass flow ratio.
Must finish the compression of final pressure all the time in the upstream of first firing chamber, in motor, form maximum pressure simultaneously.
Mind that the reason that adopts single shaft to arrange is: if a multi shaft engine, to such an extent as to must consume in order to drive the high pressure turbine of high pressure compressor on high-pressure shaft that bigger a part of output power of obtaining pressure in the high pressure turbine downstream is low to be not enough to another part exported hot gas and to infeed and make it further be heated to maximum temperature in second firing chamber from gas flow, thereby significantly reduce this output power in the low-pressure turbine.To be that the efficient of this unit is too low can not allow the people accept to the result of this structure.
On the contrary, if the whole gas turbine group of this motor all is arranged in the independent rotating shaft 7, low-pressure turbine 6 will help to drive high pressure compressor 2 so, thereby it is more reasonable that output power is distributed.Unnecessary shaft power offers the slower propeller of motion speed 9 through speed change gear 8.The number of the bearing of rotating shaft 7 can be done suitably to change by different situations.
For the variation of power demand being made reaction faster, propeller 9 also can be equipped adjustable flabellum.Have again,, in some special applications, propeller 9 is designed to multistage perhaps properly, be converted into thrust with a high efficient with unnecessary output power with turbo machine if bypass flow is smaller.
First firing chamber 3 is preferably loop configuration, in this case, this firing chamber can also comprise some axially, fiducial axis to or the independent burning space of 3 screw arrangement around the shaft.In a side of head, annular combustion chamber 3 herein as a benchmark, has some burners (not being shown specifically among the figure), and these burners are distributed in periphery and go up and produce hot gas.In fact herein can also adopt diffusion burner.In order to reduce the pollutant effulent of from this firing unit, discharging, especially NO
xEffulent preferably provides a kind of layout as the described Premix burner of European patent EP-B-0321809, constitutes an integral part of the present invention's description from the subject matter of this publication.
With regard to the layout of the Premix burner on annular combustion chamber 3 peripheral directions, if desired, this layout can be different with the conventional structure of identical burner on being distributed in periphery.On the contrary, can adopt the Premix burner of different size.This layout is preferably carried out as follows, promptly a less Premix burner that best structure is identical is being set in each casing between two big Premix burners.Must exercise pressurized air that the size of the big Premix burner of main burner functions also promptly discharges from compressor unit according to the combustion air that flows through big or small premix burner determines pro rata one by one with the little Premix burner as the pilot burner of this firing chamber.Pilot burner conduct independently Premix burner is worked in the whole load range of this firing chamber, and it is constant that in fact its air coefficient keeps.Because these pilot burner can turn round with perfect mixture in whole load range, so NO
xEven effulent is also very low under sub load.According to this configuration, the circulation line in annular combustion chamber 3 front areas is very near the eddy current center of pilot burner, so in fact only may light a fire with these pilot burner.At starting period, the amount of fuel of supplying with through pilot burner increases gradually up to this pilot burner is started, also promptly till whole amount of fuel are all available.
Under the maximum load of motor, master burner is also all started.In view of the hot eddy current center of " little " of igniting by pilot burner be in that master burner produces the colder eddy current center of " greatly " between this structure be proved to be very unstable, even if therefore when master burner is worked in the sub load scope with a weak mixture, also be that hot eddy current that pilot burner produces penetrates under the situation in the little eddy current of master burner at once, also can reach good perfect combustion, remove NO
xLow outer CO of effulent and UHC effulent are also lower.
The hot gas 15 of discharging from annular combustion chamber 3 enters and is right after this combustion chamber placement first turbo machine 4 in downstream, and the thermal expansion of 4 pairs of hot gas of this first turbo machine is variable.
The structure influence of this first turbo machine 4 temperature of demi-inflation hot gas 16, thereby determined that with the used fuel oil in second firing chamber second firing chamber whether can be by from ignition procedures work.
Here, this second firing chamber 5 has the axial or fiducial axis of a continuous circular shape basically to cylindrical form.
For this structure by looping firing chamber, an independent combustion space 5, some fuel oil jet pipes 18 are arranged on the peripheral direction of this circular cylinder, and these fuel oil jet pipes preferably are connected with each other by a loop (not shown).When this second firing chamber 5 was designed to from Spark ignition type, itself did not have burner: take place by the mode of lighting certainly in this burning from the demi-inflation hot gas 16 of turbo machine 4.If used fuel oil is a liquid fuel, 800 ℃ temperature is for based on just enough from the work of ignition way so.
Be designed under the situation of Spark ignition type in the firing chamber,, the most important thing is that flame front will keep local stable in order to ensure functional reliability and high efficiency.For this reason, be provided with the vortex generating component 11 of some in firing chamber 5, these members are preferably disposed on the inner and outer wall on the peripheral direction and preferably are arranged in the upstream of fuel oil jet pipe 18 vertically.The work of these so-called vortex generators 11 is to produce the eddy current that continues, and causes a recirculating zone according to these eddy current of European patent EP-B0321809.Based on axial arranged and consideration entire length, this firing chamber 5 is designed to a kind of high-velocity combustion chamber, and its mean velocity surpasses 60m/s, so vortex generator 11 necessary appropriate design are so that it is consistent with air-flow.In influent stream one side, these vortex generators 11 preferably are made of a tetrahedron shape with the surface that tilts with respect to influent stream.This vortex generator 11 can place on the outer surface of firing chamber 5 also can its internal surface on, as shown in the figure, also can place on these two positions.Have, from seeing the embodiment as shown in the figure, the inclined surface between the inside and outside vortex generator 11 preferably becomes mirror-image arrangement again, like this, experiences the expansion that produces eddy current in the cross section of the air-flow of the firing chamber, downstream 5 of described position in fuel injection district 18.Certainly, vortex generator 11 also can axially be provided with toward each other.The side surface that goes out to flow of vortex generator 11 is radially forming basically, therefore begins to form a recirculation zone from this position.If selected in firing chamber 5 according to from ignition way work, so just must guarantee can both be real this from lighting in all load ranges, therefore ancillary method and auxiliary device are provided, if former thereby no longer guarantee when lighting, to make the intervention firing mode owing to a certain reason or other.
Short firing chamber 5 entire length combine with the lasting effect from the vortex generator of lighting 11 of the steady flame and assurance that is used for burning are fast and effeciently carried out, and make fuel oil the shortest in the time maintenance of scorching hot flame front zone stop simultaneously.The effect relation that can go up directly measurement from burning that brings thus is to NO
xEffulent, this NO
xTo such an extent as to effulent be reduced to minimum they no longer be a problem.In addition, this initial conditions makes that the position of burning can clearly be limited, and the cooling that this clearly qualification is reflected in firing chamber 5 members obtains optimization.Afterwards, as mentioned above, the hot gas that produces in the firing chamber 5 enters second turbo machine 6 that is arranged in the downstream.
A little diffuser (not shown) if can be set in the upstream of second firing chamber 5 then can benefit to the efficient that improves motor.Thereby the loss of total pressure of whole system is reduced.Diffuser structure figure with reference to common can prove, promptly uses the diffuser of a minimum length, still can reach high kinetic pressure recovery rate.
At last, second annular combustion chamber 5 can further be cut apart as follows at peripheral direction, makes the independent cooperation combustion space of some keep the generation of hot gas.These can produce hot gas or say from the burning aspect according to optimized mode generation hot gas each other with relying in the 7 independent combustion spaces that are provided with around the shaft in a local autonomous mode.
Clearly, can modifications and variations of the present invention are according to above-mentioned given content.Therefore be appreciated that except that by the specifically described content, can in appended claims scope required for protection, implement the present invention.
Icon list
1 low pressure compressor
2 high pressure compressors
3 first firing chambers
4 first turbo machines, high pressure turbine
5 second firing chambers
6 second turbo machines, low-pressure turbine
7 rotating shafts
8 speed change gears
9 propeller, bypass flow compressor
10 bypass flow nozzles
11 vortex generators
12 air inlets, air mass flow
13 pressurized air
14 fuel oils, fuel delivery
15 hot gas
16 demi-inflation hot gas
17 final expansion hot gas, exhaust
18 fuel oils, fuel oil jet pipe
Claims (17)
1. jet engine, mainly comprise at least one propeller (9), at least one is arranged in the compressor unit (1,2) in this propeller downstream, a firing chamber (3) that is arranged in this compressor unit downstream, with a turbo machine (4) that is arranged in this downstream, firing chamber, it is characterized in that: also be furnished with one second firing chamber (5) in the downstream of described turbo machine (4), also be furnished with one second turbo machine (6) in the downstream of this second firing chamber (5).
2. jet engine as claimed in claim 1 is characterized in that: described propeller (9) is equipped with adjustable vane.
3. jet engine as claimed in claim 1 is characterized in that: described propeller (9) is for being multilevel hierarchy.
4. jet engine as claimed in claim 1 is characterized in that: at least one compressor (2), described first turbo machine (4) and described second turbo machine (6) that belong to described compressor unit are arranged in the public rotating shaft (7).
5. jet engine as claimed in claim 1 is characterized in that: the mobile machineries of all fluids of this jet engine (9,1,2,4,6) are arranged in two rotating shafts that operation links to each other each other at least.
6. jet engine as claimed in claim 5 is characterized in that: described being operatively connected of described rotating shaft can be reached by at least one speed change gear (8).
7. jet engine as claimed in claim 1 is characterized in that: described second firing chamber (5) is designed to one from the ignition combustion chamber.
8. jet engine as claimed in claim 1 is characterized in that: the upstream arrangement in described second firing chamber (5) has a diffuser.
9. jet engine as claimed in claim 1 is characterized in that: described second firing chamber is equipped with vortex generating component (11).
10. jet engine as claimed in claim 9 is characterized in that: the side surface that goes out to flow of described vortex generating component (11) has a sudden change on cross section.
11. as claim 1 and 9 described jet engines, it is characterized in that: a main fuel (18) can be sprayed in described second firing chamber (5) in described vortex generating component (11) downstream.
12. jet engine as claimed in claim 11 is characterized in that: described main fuel (18) is imported by the jet pipe of some, and these jet pipes are arranged on the peripheral direction of described second firing chamber (5).
13. jet engine as claimed in claim 1 is characterized in that: the more available Premix burner operations in described first firing chamber (3).
14. jet engine as claimed in claim 1 is characterized in that: described first and second firing chambers (3,5) are designed to annular combustion chamber.
15. jet engine as claimed in claim 1 is characterized in that: described first and/or second firing chamber (3,5) comprises some independent cylinder combustion spaces around described rotating shaft (7) layout.
16. the method for operation of jet engine as claimed in claim 1, it is characterized in that: the expansion of the hot gas (15) in described first turbo machine (4) is reduced to minimum, and the hot gas of demi-inflation (16) flows into second firing chamber (5) in downstream with the temperature higher than the spontaneous ignition temperature of the main fuel that sprays into described second firing chamber (5).
17. method as claimed in claim 16 is characterized in that: at least in the transient load scope with in the partial load scope, described second firing chamber (5) has ancillary method or auxiliary device, is guaranteed by this measure or device burning.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 98115075 CN1204005A (en) | 1997-06-26 | 1998-06-25 | Jet engine |
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| DE19726975.3 | 1997-06-26 | ||
| CN 98115075 CN1204005A (en) | 1997-06-26 | 1998-06-25 | Jet engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN1204005A true CN1204005A (en) | 1999-01-06 |
Family
ID=5224408
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN 98115075 Pending CN1204005A (en) | 1997-06-26 | 1998-06-25 | Jet engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN1204005A (en) |
Cited By (16)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101297156B (en) * | 2005-09-05 | 2010-10-20 | 西门子公司 | Burner arrangement for a combustion chamber, associated combustion chamber and method for combusting fuel |
| CN102619642A (en) * | 2010-10-19 | 2012-08-01 | 靳北彪 | Efficient turbojet engine |
| CN103047012A (en) * | 2013-01-11 | 2013-04-17 | 中国兵器工业集团第七0研究所 | Ejection intercooling gas turbine |
| CN103742295A (en) * | 2014-01-15 | 2014-04-23 | 苟仲武 | Turbojet engine and method for mixing liquid gas during operation of turbojet engine |
| CN104968893A (en) * | 2012-10-23 | 2015-10-07 | 通用电气公司 | Unducted thrust producing system architecture |
| CN105864100A (en) * | 2014-11-21 | 2016-08-17 | 通用电气公司 | Turbine engine assembly and method of manufacturing thereof |
| CN106560606A (en) * | 2015-10-06 | 2017-04-12 | 熵零股份有限公司 | Propfan engine |
| CN107725190A (en) * | 2017-09-26 | 2018-02-23 | 南京航空航天大学 | A kind of ultra-compact combustion chamber of change geometry of adjustable boundary burning |
| CN109826706A (en) * | 2019-01-24 | 2019-05-31 | 四川旭虹光电科技有限公司 | A kind of pair of oxygen-containing gas carries out the device of heating pressurized treatment |
| US11300003B2 (en) | 2012-10-23 | 2022-04-12 | General Electric Company | Unducted thrust producing system |
| CN114341479A (en) * | 2019-08-30 | 2022-04-12 | 赛峰航空器发动机 | Convergent-divergent flap pair for a variable-geometry turbojet nozzle, the flaps of which each comprise a cooling air circulation duct |
| US11391298B2 (en) | 2015-10-07 | 2022-07-19 | General Electric Company | Engine having variable pitch outlet guide vanes |
| US11492918B1 (en) | 2021-09-03 | 2022-11-08 | General Electric Company | Gas turbine engine with third stream |
| US11680530B1 (en) | 2022-04-27 | 2023-06-20 | General Electric Company | Heat exchanger capacity for one or more heat exchangers associated with a power gearbox of a turbofan engine |
| US11834954B2 (en) | 2022-04-11 | 2023-12-05 | General Electric Company | Gas turbine engine with third stream |
| US11834995B2 (en) | 2022-03-29 | 2023-12-05 | General Electric Company | Air-to-air heat exchanger potential in gas turbine engines |
-
1998
- 1998-06-25 CN CN 98115075 patent/CN1204005A/en active Pending
Cited By (25)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101297156B (en) * | 2005-09-05 | 2010-10-20 | 西门子公司 | Burner arrangement for a combustion chamber, associated combustion chamber and method for combusting fuel |
| CN102619642A (en) * | 2010-10-19 | 2012-08-01 | 靳北彪 | Efficient turbojet engine |
| US10202865B2 (en) | 2012-10-23 | 2019-02-12 | General Electric Company | Unducted thrust producing system |
| US11300003B2 (en) | 2012-10-23 | 2022-04-12 | General Electric Company | Unducted thrust producing system |
| US10907495B2 (en) | 2012-10-23 | 2021-02-02 | General Electric Company | Unducted thrust producing system |
| CN104968893A (en) * | 2012-10-23 | 2015-10-07 | 通用电气公司 | Unducted thrust producing system architecture |
| US10704410B2 (en) | 2012-10-23 | 2020-07-07 | General Electric Company | Unducted thrust producing system architecture |
| US10669881B2 (en) | 2012-10-23 | 2020-06-02 | General Electric Company | Vane assembly for an unducted thrust producing system |
| CN103047012B (en) * | 2013-01-11 | 2017-02-08 | 中国兵器工业集团第七0研究所 | Ejection intercooling gas turbine |
| CN103047012A (en) * | 2013-01-11 | 2013-04-17 | 中国兵器工业集团第七0研究所 | Ejection intercooling gas turbine |
| CN103742295A (en) * | 2014-01-15 | 2014-04-23 | 苟仲武 | Turbojet engine and method for mixing liquid gas during operation of turbojet engine |
| CN105864100A (en) * | 2014-11-21 | 2016-08-17 | 通用电气公司 | Turbine engine assembly and method of manufacturing thereof |
| CN106560606A (en) * | 2015-10-06 | 2017-04-12 | 熵零股份有限公司 | Propfan engine |
| US11391298B2 (en) | 2015-10-07 | 2022-07-19 | General Electric Company | Engine having variable pitch outlet guide vanes |
| US11585354B2 (en) | 2015-10-07 | 2023-02-21 | General Electric Company | Engine having variable pitch outlet guide vanes |
| CN107725190B (en) * | 2017-09-26 | 2019-10-15 | 南京航空航天大学 | A kind of ultra-compact combustion chamber of change geometry of adjustable boundary burning |
| CN107725190A (en) * | 2017-09-26 | 2018-02-23 | 南京航空航天大学 | A kind of ultra-compact combustion chamber of change geometry of adjustable boundary burning |
| CN109826706A (en) * | 2019-01-24 | 2019-05-31 | 四川旭虹光电科技有限公司 | A kind of pair of oxygen-containing gas carries out the device of heating pressurized treatment |
| CN114341479A (en) * | 2019-08-30 | 2022-04-12 | 赛峰航空器发动机 | Convergent-divergent flap pair for a variable-geometry turbojet nozzle, the flaps of which each comprise a cooling air circulation duct |
| CN114341479B (en) * | 2019-08-30 | 2023-06-16 | 赛峰航空器发动机 | Convergence-divergence flap pairs for variable geometry turbojet nozzles, the flaps each comprising a cooling air circulation duct |
| US11492918B1 (en) | 2021-09-03 | 2022-11-08 | General Electric Company | Gas turbine engine with third stream |
| US11859516B2 (en) | 2021-09-03 | 2024-01-02 | General Electric Company | Gas turbine engine with third stream |
| US11834995B2 (en) | 2022-03-29 | 2023-12-05 | General Electric Company | Air-to-air heat exchanger potential in gas turbine engines |
| US11834954B2 (en) | 2022-04-11 | 2023-12-05 | General Electric Company | Gas turbine engine with third stream |
| US11680530B1 (en) | 2022-04-27 | 2023-06-20 | General Electric Company | Heat exchanger capacity for one or more heat exchangers associated with a power gearbox of a turbofan engine |
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