GB2476855A - Acoustic liner and heat exchanger for gas turbine inlet duct - Google Patents
Acoustic liner and heat exchanger for gas turbine inlet duct Download PDFInfo
- Publication number
- GB2476855A GB2476855A GB201016602A GB201016602A GB2476855A GB 2476855 A GB2476855 A GB 2476855A GB 201016602 A GB201016602 A GB 201016602A GB 201016602 A GB201016602 A GB 201016602A GB 2476855 A GB2476855 A GB 2476855A
- Authority
- GB
- United Kingdom
- Prior art keywords
- layer
- acoustic attenuation
- heat exchanger
- gas turbine
- air
- 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.)
- Granted
Links
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
- F02C7/045—Air intakes for gas-turbine plants or jet-propulsion plants having provisions for noise suppression
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
- F02C7/047—Heating to prevent icing
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/08—Heating air supply before combustion, e.g. by exhaust gases
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
- B64D2033/0206—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes comprising noise reduction means, e.g. acoustic liners
Abstract
A liner 1 for an air flow passage of a gas turbine engine comprises an acoustic attenuation layer 3 forming an air-washed surface of the flow passage and a heat exchanger layer 2 which extends as a backing to the acoustic attenuation layer. The acoustic attenuation layer comprises perforated sheet 4 overlying a cellular honeycomb structure 5 extending upwardly from a backing sheet 6. The backing sheet forms a boundary between the heat exchange layer and the acoustic attenuation layer. In the heat exchange layer a corrugated sheet is sandwiched between the backing sheet and a base sheet. Side-by-side passageways are formed by the peaks and troughs of the corrugated sheet for the passage of a heated fluid. The heat exchanger layer is configured to carry a heated fluid flow and to transport heat from the fluid flow to the acoustic attenuation layer from where heat is transferred to air flowing through the air flow passage.
Description
AIR FLOW PASSAGE LINER
The present invention relates to a liner for an air flow passage of a gas turbine engine such as the bypass duct of a turbofan gas turbine engine.
The intake and bypass ducts of modern gas turbine engines are lined with acoustic attenuation panels that absorb sound energy and reduce the level of emitted noise.
The panels usually work by resonating to the sound energy produced by the engine and dissipating the energy as heat into the air. A conventional acoustic attenuation panel has a perforated sheet which overlays a cellular honeycomb structure.
As well as producing noise, a gas turbine engine generates considerable heat.
Cooling systems are needed to keep the components of the engine at operational temperatures. Specialised lubricants and oils are used to cool the components and are usually pumped around the engine in a recirculatory system. The lubricant leaving the hottest parts of the engine must be cooled to prevent overheating and degradation. For this purpose it is known to use a heat exchanger to transfer heat from the lubricant to either the fuel or to air passing through the engine. For example, a surface-air cooled, oil cooler (SA000) can be mounted on an inner surface of a fan bypass duct. Such a cooler typically has a fin and plate construction with air fins both at inner and outer sides of the cooler and with the oil passing through a central plate in a cross-flow pattern. However, such coolers, while effective at removing heat from the oil, can disturb the flow of air through the duct, which can cause efficiency losses.
Thus, in a first aspect of the invention there is provided a liner for an air flow passage of a gas turbine engine, the liner comprising: an acoustic attenuation layer which forms an air-washed surface of the flow passage, and a heat exchanger layer which extends as a backing to the acoustic attenuation layer, the heat exchanger layer being configured to carry a heated fluid flow and to transport heat from the fluid flow to the acoustic attenuation layer from where the heat is transferred to air flowing through the air flow passage.
In such a liner, heat can be transferred from the heated fluid flow to the air flow passage without compromising the acoustic attenuation properties of the acoustic attenuation layer. Indeed, the heat exchanger layer may also have acoustic attenuation properties. Furthermore, since the heat exchanger layer is beneath the acoustic attenuation layer, the heat exchanger layer does not disturb the air flow through the air flow passage, which can produce efficiency gains. Even though the heat exchanger layer extends as a backing to the acoustic attenuation layer, sufficient heat can be dissipated via the acoustic attenuation layer to adequately cool the heated fluid.
Although the heat transferred per unit area of the liner may be less than that of, for example, a conventional SA000, as the liner typically produces much less drag than a conventional SA000, the liner may compensate by covering a much larger surface area. Indeed, the surface area required by the liner may be less than or equal to the area which is in any event conventionally covered by acoustic attenuation panels.
The liner of the invention may have any one or, to the extent that they are compatible, any combination of the following optional features.
Typically, the liner is for lining an air flow passage of a turbofan gas turbine engine, such as a bypass duct of a turbofan gas turbine engine.
Typically, the heat exchanger layer of the liner is configured to carry a heated oil flow.
The heat exchanger layer may have a plurality of fins in contact with the heated fluid flow, the fins conducting heat from the heated fluid flow to the acoustic attenuation layer. The fins provide additional heat exchange surfaces that aid in the transfer of heat from the heated fluid flow to the acoustic attenuation layer. The fins typically define passageways therebetween for the heated fluid flow. For example, the fins may conveniently be formed by a corrugated sheet structure, although other fin arrangements may be adopted.
The fins may be attached to a backing sheet which is at the boundary between the heat exchanger layer and the acoustic attenuation layer. In this way, good thermal contact between the heat-exchanger layer and the acoustic attenuation layer can be achieved. The fins may be sandwiched between the backing sheet and a base sheet of the heat exchanger layer. This can provide a structure that has the strength to tolerate the potentially high pressure of the heated fluid flow.
Typically, the acoustic attenuation layer has a cellular honeycomb structure. The acoustic attenuation layer may further have a perforated sheet which overlays the cellular honeycomb structure to form the air-washed surface.
Advantageously, the combination of a heat exchanger layer having a plurality of fins and a heat exchanger layer having a cellular honeycomb structure may provide sufficient strength such that the liner can carry significant loads, allowing it to be an integral, load-bearing part of the engine.
To improve the conductive heat transfer through an acoustic attenuation layer having a cellular honeycomb structure, higher thermal conductivity materials and/or higher wall thicknesses may be used for the honeycomb structure than are typically used in conventional acoustic attenuation panels.
In a second aspect of the invention there is provided a gas turbine engine having an air flow passage lined with one or more liners according to the first aspect (the liners optionally having any one, or to the extent that they are compatible, any combination of the optional features of the first aspect). The engine may be a turbofan gas turbine engine and the lined air flow passage may be a bypass duct.
An embodiment of the invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 shows schematically a longitudinal cross-section through a ducted fan gas turbine engine; and Figure 2 shows a part cut-away perspective schematic view of a liner according to an embodiment of the invention.
With reference to Figure 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second airflow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted.
The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
Figure 2 shows a liner 1 having a heat exchanger layer 2 that extends as a backing to an acoustic attenuation layer 3. The acoustic attenuation layer 3 comprises a perforated sheet 4 overlying a cellular honeycomb structure 5. A backing sheet 6 forms the boundary between the heat exchanger layer 2 and the acoustic attenuation layer 3, the walls of honeycomb structure 5 extending upwardly from the backing sheet. In the heat exchanger layer 2, a corrugated sheet 7 is sandwiched between the backing sheet 6 and a base sheet 8. Side-by-side passageways 9 for a flow of heated fluid are formed by the peaks and troughs of the corrugated sheet 7.
In use, the liner lines the air intake 11 and/or the bypass duct 22, for example, of the turbofan gas turbine engine. The perforated sheet 4 is air-washed by air flowing through the passage of the bypass duct. Heated oil flows through the passageways 9 in the heat exchanger layer 2. The sandwich structure of the heat exchanger layer has sufficient strength to tolerate the potentially high pressure of the heated oil flow.
The undulations of the corrugated sheet 7 act as a plurality of fins extending between the backing sheet 6 and a base sheet 8. Heat is transferred from the heated oil flow to the fins and thence to the backing sheet 6, or directly to the backing sheet 6, and from there through the acoustic attenuation layer to be dissipated in the air flowing through the bypass duct. Thus the liner is able to simultaneously reduce the level of noise emitted from the engine and cool the oil flowing through the liner. The air-flow through the bypass duct is minimally disturbed because there are no fins protruding into the air flow.
While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.
Claims (9)
- CLAIMS1 A liner for an air flow passage of a gas turbine engine, the liner comprising: an acoustic attenuation layer which forms an air-washed surface of the flow passage, and a heat exchanger layer which extends as a backing to the acoustic attenuation layer, the heat exchanger layer being configured to carry a heated fluid flow and to transport heat from the fluid flow to the acoustic attenuation layer from where the heat is transferred to air flowing through the air flow passage.
- 2 A liner according to claim 1 which is for lining the bypass duct of a turbofan gas turbine engine.
- 3 A liner according to claim 1 wherein the heat exchanger layer is configured to carry a heated oil flow.
- 4 A liner according to claim 1, wherein the heat exchanger layer has a plurality of fins in contact with the heated fluid flow, the fins conducting heat from the fluid flow to the acoustic attenuation layer.
- A liner according to claim 4, wherein the fins are formed by a corrugated sheet structure.
- 6 A liner according to claim 4, wherein the fins are attached to a backing sheet which is at the boundary between the heat exchanger layer and the acoustic attenuation layer.
- 7 A liner according to claim 6, wherein the fins are sandwiched between the backing sheet and a base sheet of the heat exchanger layer.
- 8 A liner according to claim 1, wherein the acoustic attenuation layer has a cellular honeycomb structure.
- 9 A liner according to claim 1, wherein the acoustic attenuation layer further has a perforated sheet which overlays the cellular honeycomb structure to form the air-washed surface.A gas turbine engine having an air flow passage lined with one or more liners according to claim 1.11 a gas turbine engine according to claim 9 which is a turbofan gas turbine engine and the lined air flow passage is a bypass duct.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB201016602A GB2476855B (en) | 2009-11-27 | 2010-10-04 | Acoustic liner and heat exchanger for gas turbine inlet duct |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GBGB0920741.6A GB0920741D0 (en) | 2009-11-27 | 2009-11-27 | Air flow passage liner |
GB201016602A GB2476855B (en) | 2009-11-27 | 2010-10-04 | Acoustic liner and heat exchanger for gas turbine inlet duct |
Publications (3)
Publication Number | Publication Date |
---|---|
GB201016602D0 GB201016602D0 (en) | 2010-11-17 |
GB2476855A true GB2476855A (en) | 2011-07-13 |
GB2476855B GB2476855B (en) | 2012-05-02 |
Family
ID=43243407
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
GB201016602A Expired - Fee Related GB2476855B (en) | 2009-11-27 | 2010-10-04 | Acoustic liner and heat exchanger for gas turbine inlet duct |
Country Status (1)
Country | Link |
---|---|
GB (1) | GB2476855B (en) |
Cited By (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2519501A (en) * | 2013-08-01 | 2015-04-29 | Rolls Royce Plc | Acoustic liner |
FR3041704A1 (en) * | 2015-09-29 | 2017-03-31 | Snecma | THERMAL EXCHANGE AND NOISE REDUCTION PANEL FOR A PROPULSIVE ASSEMBLY |
GB2547049A (en) * | 2016-02-08 | 2017-08-09 | Gkn Aerospace Services Ltd | Integrated heater |
FR3051019A1 (en) * | 2016-05-03 | 2017-11-10 | Airbus Operations Sas | STRUCTURE PROVIDING ACOUSTIC WAVE ATTENUATION AND THERMAL EXCHANGE |
EP3483413A1 (en) * | 2017-11-14 | 2019-05-15 | The Boeing Company | Sound-attenuating heat exchangers and methods of utilizing the same |
US10619570B2 (en) | 2017-11-14 | 2020-04-14 | The Boeing Company | Dendritic heat exchangers and methods of utilizing the same |
US11143170B2 (en) | 2019-06-28 | 2021-10-12 | The Boeing Company | Shape memory alloy lifting tubes and shape memory alloy actuators including the same |
US11168584B2 (en) | 2019-06-28 | 2021-11-09 | The Boeing Company | Thermal management system using shape memory alloy actuator |
US11525438B2 (en) | 2019-06-28 | 2022-12-13 | The Boeing Company | Shape memory alloy actuators and thermal management systems including the same |
US20230167770A1 (en) * | 2021-10-13 | 2023-06-01 | Airbus Sas | Propulsion assembly for an aircraft |
Citations (5)
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US5743488A (en) * | 1994-12-05 | 1998-04-28 | Short Brothers Plc | Aerodynamic low drag structure |
GB2410769A (en) * | 2004-02-05 | 2005-08-10 | Rolls Royce Plc | Engine cooling |
US20060042225A1 (en) * | 2004-08-27 | 2006-03-02 | Pratt & Whitney Canada Corp. | Bypass duct fluid cooler |
EP1860301A2 (en) * | 2006-05-26 | 2007-11-28 | United Technologies Corporation | Micro-perforated acoustic liner |
EP2026325A2 (en) * | 2007-08-15 | 2009-02-18 | Rohr, Inc. | Linear acoustic liner |
Family Cites Families (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9938931B2 (en) * | 2008-12-23 | 2018-04-10 | General Electric Company | Combined surface cooler and acoustic absorber for turbomachines |
-
2010
- 2010-10-04 GB GB201016602A patent/GB2476855B/en not_active Expired - Fee Related
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5743488A (en) * | 1994-12-05 | 1998-04-28 | Short Brothers Plc | Aerodynamic low drag structure |
GB2410769A (en) * | 2004-02-05 | 2005-08-10 | Rolls Royce Plc | Engine cooling |
US20060042225A1 (en) * | 2004-08-27 | 2006-03-02 | Pratt & Whitney Canada Corp. | Bypass duct fluid cooler |
EP1860301A2 (en) * | 2006-05-26 | 2007-11-28 | United Technologies Corporation | Micro-perforated acoustic liner |
EP2026325A2 (en) * | 2007-08-15 | 2009-02-18 | Rohr, Inc. | Linear acoustic liner |
Cited By (17)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2519501A (en) * | 2013-08-01 | 2015-04-29 | Rolls Royce Plc | Acoustic liner |
FR3041704A1 (en) * | 2015-09-29 | 2017-03-31 | Snecma | THERMAL EXCHANGE AND NOISE REDUCTION PANEL FOR A PROPULSIVE ASSEMBLY |
US10794246B2 (en) | 2015-09-29 | 2020-10-06 | Safran Aircraft Engines | Heat-exchange and noise-reduction panel for a propulsion assembly |
GB2547049A (en) * | 2016-02-08 | 2017-08-09 | Gkn Aerospace Services Ltd | Integrated heater |
US11338933B2 (en) | 2016-02-08 | 2022-05-24 | Gkn Aerospace Services Limited | Acoustic honeycomb panel with integrated electrical heater |
GB2547049B (en) * | 2016-02-08 | 2019-12-25 | Gkn Aerospace Services Ltd | Integrated heater |
FR3051019A1 (en) * | 2016-05-03 | 2017-11-10 | Airbus Operations Sas | STRUCTURE PROVIDING ACOUSTIC WAVE ATTENUATION AND THERMAL EXCHANGE |
US10480412B2 (en) | 2016-05-03 | 2019-11-19 | Airbus Operations (S.A.S.) | Structure ensuring attenuation of acoustic waves and thermal exchange |
US10619570B2 (en) | 2017-11-14 | 2020-04-14 | The Boeing Company | Dendritic heat exchangers and methods of utilizing the same |
CN109779761A (en) * | 2017-11-14 | 2019-05-21 | 波音公司 | Noise-decaying heat exchanger and the method for utilizing it |
US11060480B2 (en) | 2017-11-14 | 2021-07-13 | The Boeing Company | Sound-attenuating heat exchangers and methods of utilizing the same |
EP3483413A1 (en) * | 2017-11-14 | 2019-05-15 | The Boeing Company | Sound-attenuating heat exchangers and methods of utilizing the same |
US11143170B2 (en) | 2019-06-28 | 2021-10-12 | The Boeing Company | Shape memory alloy lifting tubes and shape memory alloy actuators including the same |
US11168584B2 (en) | 2019-06-28 | 2021-11-09 | The Boeing Company | Thermal management system using shape memory alloy actuator |
US11525438B2 (en) | 2019-06-28 | 2022-12-13 | The Boeing Company | Shape memory alloy actuators and thermal management systems including the same |
US20230167770A1 (en) * | 2021-10-13 | 2023-06-01 | Airbus Sas | Propulsion assembly for an aircraft |
US11867122B2 (en) * | 2021-10-13 | 2024-01-09 | Airbus Sas | Propulsion assembly for an aircraft |
Also Published As
Publication number | Publication date |
---|---|
GB201016602D0 (en) | 2010-11-17 |
GB2476855B (en) | 2012-05-02 |
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Legal Events
Date | Code | Title | Description |
---|---|---|---|
PCNP | Patent ceased through non-payment of renewal fee |
Effective date: 20211004 |