GB2119859A - Exhaust mixer for bypass gas turbine aeroengine - Google Patents
Exhaust mixer for bypass gas turbine aeroengine Download PDFInfo
- Publication number
- GB2119859A GB2119859A GB08213147A GB8213147A GB2119859A GB 2119859 A GB2119859 A GB 2119859A GB 08213147 A GB08213147 A GB 08213147A GB 8213147 A GB8213147 A GB 8213147A GB 2119859 A GB2119859 A GB 2119859A
- Authority
- GB
- United Kingdom
- Prior art keywords
- stream
- bypass
- exhaust
- turbine
- turbine exhaust
- 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.)
- Withdrawn
Links
- 238000000034 method Methods 0.000 claims abstract description 8
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 5
- 230000000977 initiatory effect Effects 0.000 claims description 4
- 238000002347 injection Methods 0.000 abstract description 3
- 239000007924 injection Substances 0.000 abstract description 3
- 239000007789 gas Substances 0.000 description 9
- 239000000725 suspension Substances 0.000 description 6
- 230000001141 propulsive effect Effects 0.000 description 5
- 230000001154 acute effect Effects 0.000 description 2
- 239000000446 fuel Substances 0.000 description 2
- 238000002955 isolation Methods 0.000 description 2
- 230000000712 assembly Effects 0.000 description 1
- 238000000429 assembly Methods 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/38—Introducing air inside the jet
- F02K1/386—Introducing air inside the jet mixing devices in the jet pipe, e.g. for mixing primary and secondary flow
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Supercharger (AREA)
Abstract
The mixer 11 includes an outer mixing nozzle 23 for aiding mixing between the turbine exhaust stream 19 and a major bypass stream portion 17a, an inner mixing nozzle 25 for aiding mixing between the turbine exhaust stream and a minor bypass stream portion 17b, and air supply ducts 31 which take the minor bypass stream portion 17b from the bypass stream 17 at locations upstream of the outer mixing nozzle 23 and convey it to the interior of inner mixing nozzle 25. Amongst other advantages, the injection of bypass air 17b into the centre of turbine exhaust stream 19 reduces the velocity differential between the major bypass stream portion 17a and the turbine exhaust stream, so reducing noise generated during the mixing process. Air supply ducts 31 are conveniently fairings for struts acting to support the turbine 16 within the aeroengine. <IMAGE>
Description
SPECIFICATION
Exhaust mixer for bypass gas turbine aeroengine
The present invention relates to exhaust mixers for bypass gas turbine aeroengines in which the turbine exhaust gases and the by-pass air are combined with each other before exit from a propulsion nozzle.
It is known to mix the turbine exhaust stream with the bypass stream in a bypass gas turbine engine or "turbofan", using so-called "multilobed" and "multi-chuted" types of exhaust mixer. Such mixers improve the propulsion efficiency of this type of aeroengine by projecting portions of the two streams into each other and increasing the area of contact between them, resulting in a transfer of thermal energy from the hot turbine exhaust stream to the cooler bypass stream.
Mixers also give improved noise characteristics for the engine by reducing the amount of "jet noise" generated by the propulsive jet and ensuring that mixing between the turbine exhaust stream and the bypass stream occurs within the engine, where mixing noise can be absorbed as it arises by means of suitable acoustic linings.
It is important to note that even small improvements in the mixing efficiency can reduce noise and also significantly improve the propulsive efficiency of a mixed flow bypass gas turbine engine, thereby allowing lower specific fuel consumption.
A main objective of mixer design is therefore to maximise the contribution of the mixer to mixing efficiency and noise reduction.
According to the present invention, there is provided an exhaust mixer for a bypass gas turbine aeroengine in which the exhaust mixer combines the bypass stream with the turbine exhaust stream before passage of the combined stream through a final propulsion nozzle of the engine, the exhaust mixer including:
means for dividing the bypass stream into a major portion and a minor portion;
means for combining the major portion of the bypass stream with the turbine exhaust stream; and
means for combining the minor portion of the bypass stream with the turbine exhaust stream such that the process of combination of the minor portion of the bypass stream with the turbine exhaust stream is initiated radially inwardly of, and independently of, the initiation of the process of combination of said major-portion of the bypass stream with the turbine exhaust stream.
The means for dividing the bypass stream into a major portion and a minor portion may comprise a plurality of duct means arranged to take said minor portion from the bypass stream upstream of the means for combining the major portion of the bypass stream with the turbine exhaust stream, and to deliver said minor portion to the means for combining said minor portion with the turbine stream. Conveniently, the duct means comprise fairings for strut members which support the turbine within the aeroengine.
The means for combining the major portion of the bypass stream with the turbine exhaust stream preferably comprises a mixing nozzle whose inner surface is contacted by the turbine exhaust stream and whose outer surface is contacted by the major portion of the bypass stream, whilst the means for combining the minor portion of the bypass stream with the turbine exhaust stream preferably comprises a mixing nozzle whose outer surface is contacted by the turbine exhaust stream and whose inner surface is contacted by the minor portion of the bypass stream. Either mixing nozzle may be of the convoluted or multi-lobed type.
Further aspects of the invention will be apparent from the accompanying description and claims.
An embodiment of the invention will now be described by way of example only with reference to the accompanying drawings in which:
Figure 1 shows a partly "broken-away" side view in diagrammatic form of a turbofan aeroengine fitted with an exhaust mixer in accordance with the invention;
Figure 2 shows an enlarged part-sectioned side view of part of the "broken-away" area of Figure 1;
Figure 3 is a view on arrow A in Figure 2, showing the configuration of the exhaust mixer as seen in rear elevation; and
Figure 4 is a perspective view of the exhaust mixer nozzles as seen in isolation from other structure.
The drawings are not to scale.
Referring first to Figure 1, a low bypass ratio gas turbine aeroengine 1 has: an engine core 3; a bypass duct 5 defined between the engine core 3 and the outer engine casing 7; and an exhaust system 9, which includes the exhaust mixer 1 a jet pipe 12 and a final propulsive exhaust nozzle 1 3. The bypass duct 5 is supplied with bypass air from low pressure compressor or fan 15, which also supplies core 3, the fan 1 5 being driven from low pressure turbine 1 6 in core 3. The fan air stream 1 7 and the turbine exhaust stream 19 are combined by the mixer 11 and mixing continues in the jet pipe 12 of exhaust system 9 before exit of the combined stream to atmosphere through propulsion nozzle 13.In order to absorb mixing noise as it arises, jet pipe 12 is provided with acoustic linings 20 as known.
It will be seen from Figure 1 that the exhaust mixer 11 includes an outer mixing nozzle 23 (shown only schematically) for aiding mixing between the turbine exhaust stream 1 9 and a major portion 1 7a of the bypass stream 17, an inner mixing nozzle 25 for aiding mixing between the turbine exhaust stream 1 9 and a minor portion 1 7b of the bypass stream 17, and air supply ducts 31 for taking the minor portion 1 7b of the bypass stream from the bypass stream at locations upstream of the outer mixing nozzle 23 and conveying it to the interior of the inner mixing nozzle 23 and conveying it to the interior of the inner mixing nozzle 25.The bypass stream 17 is thus divided into the major and minor portions 1 7a and 1 7b respectively by means of the air supply ducts 31, and the process of combination of the minor portion 1 7b of the bypass stream with the turbine exhaust stream, which is facilitated by inner mixing nozzle 25, is initiated radially inwardly of, and independently of the initiation of the process of combination of the
major portion 1 7a of the bypass stream with the turbine exhaust stream, which is facilitated by outer mixing nozzle 23.
Mixing nozzle 23 is attached to the rear end of outer core casing 27 of engine core 3 and is contacted on its outer surfaces by the major portion 1 7a of the bypass stream and on its inner surfaces by turbine exhaust stream 1 9. Mixing nozzle 25 is fixed to the rear end of inner casing 29 of engine core 3, its outer surfaces being in contact with the turbine exhaust stream 19, but its inner surfaces being in contact with the minor portion 1 7b of the bypass stream.As the inner
nozzle 25 feeds bypass air into the centre of the turbine exhaust stream 19, the average velocity of
at least the central regions of the turbine exhaust stream is thereby reduced; in the present embodiment this combination of the minor bypass stream portion 1 7b with turbine exhaust stream 1 9 begins nearly simultaneousiy with that of major bypass portion 1 7a with stream 19, but this could easily be emprically modified within the scope of the invention to give optimum mixing noise reduction by altering somewhat the relative axial positions of the respective trailing edges of the inner and outer mixing nozzles.The reduction in velocity differential between the main body of the bypass stream and the turbine exhaust stream reduces the amount of noise generated by the mixing process and therefore reduces the amount of noise emanating from the propulsion nozzle 1 3 and the propulsive jet.
Figures 2 to 4 show the exhaust mixer 11 and associated structure in more detail. Figure 2 is in fact a composite view, being composed of the view on section B-B in Figure 3 so far as concerns the air supply ducts 31, and the view on section C-C in Figure 3 so far as concerns the nozzles 23 and 25 and the outer engine casing 7.
In fact, Figure 3 does not show the engine nacelle and outer engine casing 7 or the jet pipe 12, for reasons of clarity, whilst Figure 4 shows only the two mixing nozzles 23, 25 in isolation from other structure and as seen from a positipn forward of flanges 24, 26 and offset from the centreline 32 of the mixer 11.
Dealing first with outer mixer nozzle 23, this has a convoluted or corrugated appearance when seen in end elevation, comprising-ten convexoutward bulges or "lobes" 33 which are equiangularly spaced from each other around the nozzle, and which flare outwardly and rearwardly from their points of origin at the end of frustoconical surface 35. Lobes 33 are separated from each other by concavities or troughs 37, 39 between the lobes, these troughs being directed rearwardly and inwardly towards the centreline 32
of the exhaust mixer 11.Lobes 33 act to channel
portions of the turbine exhaust stream 1 9 outwards into the surrounding bypass air stream 1 7a whilst the troughs 37, 39 channel portions of
the bypass air stream inwards into the interior of
the turbine exhaust stream, mixing occurring by
virtue of forced interpenetration of the two
streams.
It will be noted that troughs 37 are shorter in
radial and axial extent than troughs 39, the effect
as seen in rear view in Figure 3 being that of
alternate long and short inwardly projecting
"fingers". As seen more clearly in Figure 2, the
long and short troughs are formed by scarfing
portions of the nozzle forwardly or rearwardly of a
diametral plane 41 containing the radially
outermost tips of the lobes 33.Thus, for example,
short trough 37' is a'result of the trailing
(rearmost) edge of the nozzle in the sector defined
between radial lines D and E (Figure 3) being
coincident with an imaginary forwardly inclined
conical surface 43 with its apex at 44 which
makes an acute angle of Owith plane 41, whilst
long trough 39' is a result of the trailing edge of
the nozzle between radial lines F and D being
coincident with an imaginary rearwardly inclined
conical surface 45 with its apex at 46 which makes an acute angle of Q1 with plane 41, Angles O and QI may be equal to each other, but are not
necessarily so.
This configuration of mixing nozzle 23 has the
advantage that mixing between the two streams 1 7a and 1 9 is initiated progressiveiy over a rear portion of the axial extent of the nozzle.
The inner mixing nozzle 25 may also be said to
have a convoluted or corrugated appearance when
seen in end elevation, through its shape may
be specifically likened to that of a five-pointed
star. It comprises five lobes 47 which
progressively flare directly out of what is basically
a frusto-conical surface, the lobes 47 again being
equi-angularly spaced from each other around the
nozzle and defining troughs 49 between them.
Troughs 49 act to channel turbine exhaust gases
inwards, whilst lobes 47 act to channel by-pass air
outward, mixing again occurring by forced
interpenetration of streams 1 7b and 1 9.
The ducts 31, which supply bypass air to inner
nozzle 25, form fairings for bearing support struts
51 which radiate from bearing housing 53 in the
manner of spokes from a hub and whose outer
ends are attached to suspension ring 53 which
circumscribes the outer open ends 63 of air supply
ducts 31. Suspension ring 53 is bolted to the
outer core casing 27 of the engine core through
flange feature 55 and is connected to suspension
links 57 by means of bolts through lugs 59 on ring
53. Suspension links 57 are in turn connected to
the outer engine casing 7 and nacelle through
bolted mounting assemblies 61. Suspension links
57, suspension ring 53 and struts 51 act together
in known manner in order to maintain
concentricity between outer engine casing 7,
outer core casing 27, and bearing housing 53
despite differential expansion due to high engine core temperatures.One advantage associated with the invention is that the flow of bypass air from by-pass duct 5, through air supply ducts 31 via their outer and inner open ends 63 and 65 respectively, and then through nozzle 25, acts to cool ducts 31, struts 51, nozzle 25 and the rear end of inner casing 29, thereby helping to reduce differential expansion and also extend the useful iives of the components.
A further expected advantage of the invention arises from the fact that the injection of bypass air into the centre of the turbine exhaust stream, allied with the lobed shape of the inner mixing nozzle 25, should reduce the circumferentially directed ("whirl") component of velocity of the turbine exhaust stream imparted by turbine 1 6 and consequently should also reduce the whirl component of velocity of the mixed stream in the jet-pipe 1 2. Reduction of whirl velocity in propulsive jets is desirable because it reduces the specific fuel consumption of the engine by increasing the effective thrust.
An attractive feature of the invention is the fact that existing engines can be modified to incorporate it, the inner mixing nozzle 25 being a replacement for the inner exhaust cone located there in prior art engines, and air supply ducts 31 being already present in the existing engines as fairings for struts 51 but being blanked off at their radially outer and radially inner ends 63, 65 respectively.
Although the invention has been specifically described in relation to a low bypass ratio gas turbine engine, it could also be incorporated in a high bypass ratio engine of the type in which the core and bypass streams are mixed interna!ly of the engine before passage through a final propulsion nozzle.
The exhaust mixer described above has inner and outer mixing nozzles of the "convoluted", multi-lobed type. However, it is possible to utilise other means known per sue, for forced-mixing of the bypass and turbine exhaust streams whilst remaining within the scope of the invention. For example, the outer mixing nozzle 23 could be replaced with an annular member incorporating apertures or ducts such that the apertures or ducts would divide the bypass stream into a plurality of jets of bypass air which would penetrate inwards into the turbine exhaust stream. Similarly, the inner mixing nozzle could be replaced with means such as a muiti-ducted nozzle as known, which would divide the diverted minor portion of the bypass stream into a plurality of jets of bypass air which would also penetrate into the turbine exhaust stream. It might even prove possible, within the scope of the invention, to replace either or both of the convoluted mixing nozzles 23 and 25 by circular nozzles (perhaps provided with turbulence producing features on their trailing edges) as the means for initiating corflbination 9f the streams, thus saving weight and manufacturing cost yet still retaining advantages associated with injection of a portion of the bypass flow into the central region of the turbine exhaust flow before passage of the combined stream through the propulsion nozzle.
Claims (11)
1. An exhaust mixer for a bypass gas turbine aeroengine in which the exhaust mixer combines the bypass stream with the turbine exhaust stream before passage of the combined stream through a final propulsion nozzle of the engine, the exhaust mixer including:
means for dividing the bypass stream into a major portion and a minor portion;
means for combining the major portion of the bypass stream with the turbine exhaust stream; and
means for combining the minor portion of the bypass stream with the turbine exhaust stream such that the process of combination of the minor portion of the bypass stream with the turbine exhaust stream is initiated radially inwardly of, and independently of, the initiation of the process of combination of said major portion of the bypass stream with the turbine exhaust system.
2. An exhaust mixer according to claim 1 in which the means for dividing the bypass stream into a major portion and a minor portion comprises a plurality of duct means arranged to take said minor portion from the bypass stream upstream of the means for combining the major portion of the bypass stream with the turbine exhaust stream, and to deliver said minor portion to the means for combining said minor portion with the turbine exhaust stream.
3. An exhaust mixer according to claim 2 in which the duct means comprise fairings for strut members which support the turbine within the aeroengine.
4. An exhaust mixer according to any one of
claims 1 to 3 in which the means for combining
the major portion of the bypass stream with the
turbine exhaust stream comprises a mixing nozzle
whose inner surface is contacted by the turbine
exhaust stream and whose outer surface is
contacted by the major portion of the bypass
stream.
5. An exhaust mixer according to any one of
claims 1 to 4 in which the means for combining
the minor portion of the bypass stream with the
turbine exhaust stream comprises a mixing nozzle
whose outer surface is contacted by the turbine
exhaust stream and whose inner surface is
contacted by the minor portion of the bypass
stream.
6. An exhaust mixer according to claim 4 or
claim 5 in which the mixing nozzle is of the
convoluted or multi-lobed type.
7. An exhaust mixer for a bypass gas turbine - aernengine in which the turbine exhaust stream
and the bypass air stream are combined by the
exhaust mixer before exit through a final
propulsion nozzle, the exhaust mixer including an
outer mixing nozzle for aiding mixing between the
turbine exhaust stream and a major portion of the
bypass stream, an inner mixing nozzle for aiding
mixing between the turbine exhaust stream and a minor portion of the bypass stream, and duct means for taking said minor portion of the bypass stream from the bypass stream upstream of the outer mixing nozzle and conveying said minor portion to the inner mixing nozzle.
8. An exhaust mixer according to claim 6 in which at least one of the mixing nozzles is of the convoluted or multi-lobed type.
9. An exhaust mixer according to claim 7 or claim 8 in which the duct means comprise fairings for strut members which support the turbine within the aeroengine.
10. An exhaust mixer substantially as described in this specification with reference to and as illustrated in the accompanying drawings.
11. A bypass gas turbine aeroengine fitted with an exhaust mixer according to any one of claims 1 to 10.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08213147A GB2119859A (en) | 1982-05-06 | 1982-05-06 | Exhaust mixer for bypass gas turbine aeroengine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08213147A GB2119859A (en) | 1982-05-06 | 1982-05-06 | Exhaust mixer for bypass gas turbine aeroengine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| GB2119859A true GB2119859A (en) | 1983-11-23 |
Family
ID=10530204
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB08213147A Withdrawn GB2119859A (en) | 1982-05-06 | 1982-05-06 | Exhaust mixer for bypass gas turbine aeroengine |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2119859A (en) |
Cited By (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2207468A (en) * | 1987-06-01 | 1989-02-01 | Secr Defence | Vortex silencing in gas turbine engines |
| FR2657399A1 (en) * | 1990-01-25 | 1991-07-26 | Gen Electric | MIXER ARRANGEMENT FOR A DOUBLE FLOW GAS TURBINE ENGINE AND MOTOR THUS OBTAINED. |
| WO1998059162A1 (en) * | 1997-06-24 | 1998-12-30 | Sikorsky Aircraft Corporation | Multi-stage mixer/ejector for suppressing infrared radiation |
| WO1998059163A1 (en) * | 1997-06-24 | 1998-12-30 | Sikorsky Aircraft Corporation | Exhaust nozzle for suppressing infrared radiation |
| EP0913568A3 (en) * | 1997-10-30 | 2000-07-26 | Stage III Technologies L.C. | Lobed mixer/ejector nozzle |
| WO2000053915A1 (en) * | 1999-03-05 | 2000-09-14 | Rolls-Royce Deutschland Gmbh | Bloom mixer for a turbofan engine |
| US6786038B2 (en) * | 2002-02-22 | 2004-09-07 | The Nordam Group, Inc. | Duplex mixer exhaust nozzle |
| GB2447291A (en) * | 2007-03-08 | 2008-09-10 | John Edward Randell | A turbojet engine having a bypass flow through the engine core |
| US7543452B2 (en) * | 2005-08-10 | 2009-06-09 | United Technologies Corporation | Serrated nozzle trailing edge for exhaust noise suppression |
| FR2952403A1 (en) * | 2009-11-12 | 2011-05-13 | Snecma | ANNULAR METAL CONNECTION STRUCTURE FOR AIRCRAFT TURBOMACHINE |
| US8371806B2 (en) * | 2007-10-03 | 2013-02-12 | United Technologies Corporation | Gas turbine engine having core auxiliary duct passage |
| WO2014007907A3 (en) * | 2012-04-27 | 2014-04-17 | General Electric Company | Variable immersion lobe mixer for turbofan jet engine exhauts and method of fabricating the same |
| US10738649B2 (en) | 2017-08-03 | 2020-08-11 | Rolls-Royce Corporation | Reinforced oxide-oxide ceramic matrix composite (CMC) component and method of making a reinforced oxide-oxide CMC component |
| CN115614177A (en) * | 2022-08-29 | 2023-01-17 | 中国航发四川燃气涡轮研究院 | Full-shielding mixing integrated casing |
| CN115949530A (en) * | 2023-03-09 | 2023-04-11 | 中国航发四川燃气涡轮研究院 | Stealthy device of dysmorphism spray tube |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB837501A (en) * | 1957-08-07 | 1960-06-15 | Rolls Royce | Improved jet noise suppressor nozzle |
| GB1442740A (en) * | 1970-12-10 | 1976-07-14 | Gen Motors Corp | Gas turbine exhaust duct |
| GB1506034A (en) * | 1974-06-11 | 1978-04-05 | United Technologies Corp | Turbofan engine with augmented combustion chamber |
| US4214441A (en) * | 1978-09-12 | 1980-07-29 | The United States Of America As Represented By The Secretary Of The Navy | Infrared suppressor device |
| GB2038948A (en) * | 1978-12-21 | 1980-07-30 | Secr Defence | Gas Turbine By-pass Jet Engines |
| GB1601807A (en) * | 1970-07-24 | 1981-11-04 | Gen Motors Corp | Ducted fan engine exhaust mixer |
-
1982
- 1982-05-06 GB GB08213147A patent/GB2119859A/en not_active Withdrawn
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB837501A (en) * | 1957-08-07 | 1960-06-15 | Rolls Royce | Improved jet noise suppressor nozzle |
| GB1601807A (en) * | 1970-07-24 | 1981-11-04 | Gen Motors Corp | Ducted fan engine exhaust mixer |
| GB1442740A (en) * | 1970-12-10 | 1976-07-14 | Gen Motors Corp | Gas turbine exhaust duct |
| GB1506034A (en) * | 1974-06-11 | 1978-04-05 | United Technologies Corp | Turbofan engine with augmented combustion chamber |
| US4214441A (en) * | 1978-09-12 | 1980-07-29 | The United States Of America As Represented By The Secretary Of The Navy | Infrared suppressor device |
| GB2038948A (en) * | 1978-12-21 | 1980-07-30 | Secr Defence | Gas Turbine By-pass Jet Engines |
Cited By (24)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2207468A (en) * | 1987-06-01 | 1989-02-01 | Secr Defence | Vortex silencing in gas turbine engines |
| FR2657399A1 (en) * | 1990-01-25 | 1991-07-26 | Gen Electric | MIXER ARRANGEMENT FOR A DOUBLE FLOW GAS TURBINE ENGINE AND MOTOR THUS OBTAINED. |
| WO1998059162A1 (en) * | 1997-06-24 | 1998-12-30 | Sikorsky Aircraft Corporation | Multi-stage mixer/ejector for suppressing infrared radiation |
| WO1998059163A1 (en) * | 1997-06-24 | 1998-12-30 | Sikorsky Aircraft Corporation | Exhaust nozzle for suppressing infrared radiation |
| US5992140A (en) * | 1997-06-24 | 1999-11-30 | Sikorsky Aircraft Corporation | Exhaust nozzle for suppressing infrared radiation |
| US6016651A (en) * | 1997-06-24 | 2000-01-25 | Sikorsky Aircraft Corporation | Multi-stage mixer/ejector for suppressing infrared radiation |
| EP0913568A3 (en) * | 1997-10-30 | 2000-07-26 | Stage III Technologies L.C. | Lobed mixer/ejector nozzle |
| WO2000053915A1 (en) * | 1999-03-05 | 2000-09-14 | Rolls-Royce Deutschland Gmbh | Bloom mixer for a turbofan engine |
| US6578355B1 (en) | 1999-03-05 | 2003-06-17 | Rolls-Royce Deutschland Ltd & Co Kg | Bloom mixer for a turbofan engine |
| US6786038B2 (en) * | 2002-02-22 | 2004-09-07 | The Nordam Group, Inc. | Duplex mixer exhaust nozzle |
| US7543452B2 (en) * | 2005-08-10 | 2009-06-09 | United Technologies Corporation | Serrated nozzle trailing edge for exhaust noise suppression |
| GB2447291A (en) * | 2007-03-08 | 2008-09-10 | John Edward Randell | A turbojet engine having a bypass flow through the engine core |
| US8371806B2 (en) * | 2007-10-03 | 2013-02-12 | United Technologies Corporation | Gas turbine engine having core auxiliary duct passage |
| FR2952403A1 (en) * | 2009-11-12 | 2011-05-13 | Snecma | ANNULAR METAL CONNECTION STRUCTURE FOR AIRCRAFT TURBOMACHINE |
| CN102597429A (en) * | 2009-11-12 | 2012-07-18 | 斯奈克玛 | Metal annular connecting structure for an aircraft turbine engine |
| WO2011058041A1 (en) * | 2009-11-12 | 2011-05-19 | Snecma | Metal annular connecting structure for an aircraft turbine engine |
| CN102597429B (en) * | 2009-11-12 | 2014-11-19 | 斯奈克玛 | Metal annular connecting structure for an aircraft turbine engine |
| US9169743B2 (en) | 2009-11-12 | 2015-10-27 | Snecma | Metallic annular connection structure for aircraft turbomachine |
| WO2014007907A3 (en) * | 2012-04-27 | 2014-04-17 | General Electric Company | Variable immersion lobe mixer for turbofan jet engine exhauts and method of fabricating the same |
| US9995245B2 (en) | 2012-04-27 | 2018-06-12 | General Electric Company | Variable immersion lobe mixer for turbofan jet engine exhaust and method of fabricating the same |
| US10738649B2 (en) | 2017-08-03 | 2020-08-11 | Rolls-Royce Corporation | Reinforced oxide-oxide ceramic matrix composite (CMC) component and method of making a reinforced oxide-oxide CMC component |
| CN115614177A (en) * | 2022-08-29 | 2023-01-17 | 中国航发四川燃气涡轮研究院 | Full-shielding mixing integrated casing |
| CN115614177B (en) * | 2022-08-29 | 2024-04-16 | 中国航发四川燃气涡轮研究院 | Full shielding blending integrated casing |
| CN115949530A (en) * | 2023-03-09 | 2023-04-11 | 中国航发四川燃气涡轮研究院 | Stealthy device of dysmorphism spray tube |
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