GB2189844A - Gas turbine engines - Google Patents
Gas turbine engines Download PDFInfo
- Publication number
- GB2189844A GB2189844A GB08610559A GB8610559A GB2189844A GB 2189844 A GB2189844 A GB 2189844A GB 08610559 A GB08610559 A GB 08610559A GB 8610559 A GB8610559 A GB 8610559A GB 2189844 A GB2189844 A GB 2189844A
- Authority
- GB
- United Kingdom
- Prior art keywords
- turbine
- rotor
- aerofoil
- compressor
- circumferentially arranged
- 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- Withdrawn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
- 239000000446 fuel Substances 0.000 claims description 20
- 238000011144 upstream manufacturing Methods 0.000 claims description 14
- 239000007789 gas Substances 0.000 description 35
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 239000012530 fluid Substances 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
- 230000003068 static effect Effects 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
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/065—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front and aft fans
-
- 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
- F02C5/00—Gas-turbine plants characterised by the working fluid being generated by intermittent combustion
- F02C5/02—Gas-turbine plants characterised by the working fluid being generated by intermittent combustion characterised by the arrangement of the combustion chamber in the chamber in the plant
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/072—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with counter-rotating, e.g. fan rotors
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/56—Combustion chambers having rotary flame tubes
-
- 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)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
A turbofan gas turbine engine comprises a forward fan 50 and a rearward fan 80 arranged to rotate in opposite directions within a bypass duct 22. The forward fan 50 is driven by a first drum rotor 42 of a compressor 14, the first drum rotor rotating in the same direction but at a relatively slower speed than a second drum rotor 52 of the compressor. The rearward fan 80 is driven by a first turbine drum rotor 72 of a turbine 18, the first turbine drum rotor rotating in the opposite direction but at a relatively slower speed than a second turbine drum rotor 82 of the turbine. The second turbine drum rotor is arranged to drive the second drum rotor of the compressor. A combustor 16 is positioned axially between the forward and rearward fans, and may be stationary or arranged to rotate independently or with the first turbine drum rotor. Contra-rotating propellers may be driven in a similar way. <IMAGE>
Description
SPECIFICATION
Gas turbine engines
The present invention relates to gas turbine engines and in particular to turbofan, or bypass gas turbine engines, and turbopropeller gas turbine engines.
In turbofan, or bypass, gas turbine engines it is known in the prior art to provide a separate turbine stage downstream of a core engine to drive a fan, or fans, or a separate power turbine, with contra-rotating turbine stages, downstream of the core engine to drive contra-rotating fans.
In the case of contra-rotating fans it is desireable to have the fans axially spaced by as large a distance as is possible, on any particular gas turbine engine, so as to reduce the noise produced by the contra-rotating fans.
The contra-rotating fans driven by a power turbine, with contra-rotating turbine stages, have a relatively short axial spacing.
The present invention seeks to provide a bypass gas turbine engine with contra-rotating fans which has a relatively large axial spacing between the contra-rotating fans to reduce noise.
A turbopropeller gas turbine engine with contra-rotating propellers driven by contra-rotating turbine stages of a separate power turbine also has a relatively short axial spacing between the propellers.
The present invention seeks to provide a turbo-propeller gas turbine engine with contrarotating propellers which has a relatively large axial spacing between the contra-rotating propellers to reduce noise.
Accordingly the present invention provides a gas turbine comprising a core engine, and a forward and a rearward aerofoil arrangement, the core engine comprising in flow series a compressor, a combustor and a turbine, the compressor comprising a first rotor having a plurality of circumferentially arranged first blades extending radially inwards therefrom, a second rotor spaced radially inwardly from the first rotor and having a plurality of circumferentially arranged second blades extending radially outwards therefrom, the first and second blades being arranged axially alternately, the first and second rotors being adapted to rotate in the same direction, the turbine comprising a first turbine rotor having a plurality of circumferentially arranged first turbine blades extending radially inwards therefrom, a second turbine rotor spaced radially inwardly from the first turbine rotor and having a plurality of circumferentially arranged second turbine blades extending radially outwards therefrom, the first and second turbine rotors being adapted to contra-rotate, the second turbine rotor being adapted to drive the second rotor of the compressor, the forward aerofoil arrangement comprising a plurality of circumferentially arranged first aerofoil blades extending radially outwards from and being driven by the first rotor of the compressor, the rearward aerofoil arrangement comprising a plurality of circumferentially arranged second aerofoil blades extending radially outwards from and being driven by the first turbine rotor in contra-rotation to the forward aerofoil arrangement.
The present invention also provides a gas turbine engine comprising a core engine, and a forward and a rearward aerofoil arrangement, the core engine comprising in flow series a compressor, a combustor and a turbine, the compressor comprising a first rotor having a plurality of circumferentially arranged first blades extending radially inwards therefrom, a second rotor having a plurality of circumferentially arranged second blades extending radially outwards therefrom, the first and second blades being arranged axially alternately, the turbine comprising a first turbine rotor having a plurality of circumferentially arranged first turbine blades extending radially inwards therefrom, a second turbine rotor having a plurality of circumferentially arranged second turbine blades extending radially outwards therefrom, the first and second turbine rotors being adapted to contra-rotate, the first turbine rotor being adapted to drive the first rotor of the compressor, the second turbine rotor being adapted to drive the second rotor of the compressor, the first and second turbine rotors being adapted to contra-rotate the turbine being adapted to drive the forward aerofoil arrangement and rearward aerofoil arrangement in opposite directions.
The forward aerofoil arrangement may comprise a plurality of circumferentially arranged first aerofoil blades driven by the first turbine rotor, the rearward aerofoil arrangement may comprise a plurality of circumferentially arranged second aerofoil blades driven by the second turbine rotor in the opposite direction to the forward aerofoil arrangement.
The forward aerofoil arrangement may comprise a plurality of circumferentially arranged first aerofoil blades driven by the second turbine rotor, the rearward aerofoil arrangement may comprise a plurality of circumferentially arranged second aerofoil blades driven by the first turbine rotor in the opposite direction to the forward aerofoil arrangement.
The gas turbine engine may be a bypass gas turbine engine in which the forward and rearward aerofoil arrangements are a forward and a rearward fan arranged within a bypass duct.
The gas turbine engine may be a turbopropeller gas turbine engine in which the forward and rearward aerofoil arrangements are a forward and a rearward propeller.
A cylindrical body may be positioned coaxially within the second rotor of the compressor and the second turbine rotor, the cylindrical body extending axially from a first vane ar rangement at the upstream end of the compressor to a second vane arrangement at the downstream end of the turbine, the first and second rotors of the compressor and the first and second turbine rotors being rotatably mounted on the cylindrical body.
The combustor may be secured to a casing forming the outer boundary of the bypass duct by a number of struts extending across the bypass duct.
The combustor may be rotatably mounted on the cylindrical body. The combustor may be secured to the upstream end of the first turbine rotor and be adapted to rotate with the first turbine rotor.
The invention will be more fully described by way of reference to the accompanying drawings in which:
Figure 1 is a cut-away view of a gas turbine engine according to the present invention.
Figure 2 is a cut-away view of a second embodiment of a gas turbine engine according to the present invention.
Figure 3 is a cut-away view of a third embodiment of a gas turbine engine according to the present invention.
Figure 4 is a cut-away view of a fourth embodiment of a gas turbine engine according to the present invention.
Figure 5 is a cut-away view of a fifth embodiment of a gas turbine engine according to the present invention.
Figure 6 is a cut-away view of a further embodiment of a gas turbine engine according to the present invention.
A gas turbine engine according to the present invention is shown in Fig. 1, and is of the bypass or turbofan type. The gas turbine engine 10 comprises a core engine which has in flow series an inlet 12, a compressor 14, a combustor 16, a turbine 18 and an exhaust nozzle 20. A bypass duct 22 which has an inlet 24 and an outlet 26 is formed coaxially around the core engine by a fan casing 28.
A circumferential arrangement of vanes 30 is provided at the upstream end of the compressor 14 and the vanes 30 are secured to a compressor casing 32 at their outer radial ends and to an axially extending cylindrical body 34 at their inner radial ends. The cylindrical body 34 is positioned coaxially with the gas turbine engine and extends substantially the full length of the gas turbine engine to provide stiffness, for carrying the rotating parts of the engine. Likewise a circumferential arrangement of vanes 36 is provided at the downstream end of the turbine 18, and the vanes 36 are secured to the cylindrical body 34 at their radially inner ends and to a casing 38 at their radially outer ends. The casing 38 is secured to the downstream end of the fan casing 28 by a number of radially extending struts 40.
The compressor 14 comprises a first drum rotor 42 which has a plurality of circumferenti
ally arranged first blades 44 extending radially
inwardly therefrom, and which are arranged in
axially spaced stages on the drum rotor 42.
The upstream stage of first blades has a bear
ing 46 by which the first drum rotor 42 is
rotatably mounted onto the cylindrical body
34, and the downstream stage of first blades has a bearing 48 by which the first drum rotor is rotatably mounted onto the combustor
16. A forward fan 50 comprising a circumferential arrangement of radially outward extend
ing fan blades is positioned upstream of the combustor 16, the fan blades being secured to the downstream end of the first drum rotor 42 of the compressor 14.
The compressor 14 also comprises a second drum rotor 52 which is spaced radially inwardly from the first drum rotor 42 and which has a plurality of circumferentially arranged second blades 54 extending radially outwardly therefrom, and which are also arranged in axially spaced stages on the drum rotor 52, and are arranged axially alternately with the first blades 44 on the first drum rotor 42. The upstream end of the second drum rotor is rotatably mounted onto the cylindrical body 34 by a bearing 56, and the downstream end of the second drum rotor is connected to a shaft 58.
The combustor 16 comprises an annular flame tube 60 or a can-annular arrangement of flame tubes, which has or have inlet guide vanes 62 at the upstream end and outlet guide vanes 64 at the downstream end. The combustor 16 is secured to the fan casing 28 by a number of struts 66 which extend across the bypass duct 22 downstream of the forward fan 50.
Bearings 68 and 70 are interposed between static structures extending radially inwardly from the inlet and outlet guide vanes 62 and 64 of the combustor 16 and the shaft 58.
The turbine 18 comprises a first turbine drum rotor 72 which has a plurality of circumferentially arranged first turbine blades 74 extending radially inwardly therefrom, and which are arranged in axially spaced stages on the first turbine drum rotor 72. The stage of blades at the upstream end of the first rotor 72 has a bearing 76 by which the first rotor 72 is rotatably mounted onto the combustor 16, and the downstream stage of blades has a bearing 78 by which the first rotor is rotatably mounted onto the cylindrical body 34. A rearward fan 80 comprising a circumferential arrangement of radially outward extending fan blades is positioned downstream of the combustor 16, the fan blades being secured to the downstream end of the first turbine drum rotor 72 of the turbine 18.
The turbine 18 also comprises a second turbine drum rotor 82 whic is spaced radially inwardly from the first turbine drum rotor 72 and which has a plurality of circumferentially arranged second turbine blades 84 extending radially outwardly therefrom, and which are also arranged in axially spaced stages on the drum rotor 82, and are arranged axially alternately with the first turbine blades 74 on the first rotor 72.
The upstream end of the second turbine drum rotor 82 is secured to the shaft 58, and the downstream end of the second turbine drum rotor 82 is rotatably mounted on the cylindrical body 34 by a bearing 86.
The gas turbine engine 10 may be mounted to an aircraft fuesalage 90, for example, by a strut 88 which extends from the upstream end of the compressor casing 32, and a second strut 92 which extends from the fan casing 28. It may also be secured to an aircraft wing by suitable means.
In operation air flows into the inlet 12 of the core engine of the gas turbine engine 10 and is compressed by the compressor 14.
The first and second drum rotors 42,52 of the compressor 14 are arranged to rotate in the same direction, but the second drum rotor 52 rotates at a relatively higher speed than the first drum rotor 42. This enables the speed of rotation of the forward fan 50, which is driven by the first drum rotor 42, to be relatively low and optimised for the compression of air flowing through the bypass duct 22.
The compressed air leaving the compressor 14 is supplied into the combustor 16, where fuel is burnt in the air to produce hot gases which are supplied into the turbine 18. The first and second turbine drum rotors 72 and 82 are arranged to rotate in opposite directions, with the second turbine drum rotor 82 driving the second drum rotor 52 of the compressor 14. The first turbine drum rotor 72 drives the rearward fan 80 in the opposite direction to the forward fan 50, and the second turbine drum rotor 82 rotates at a relatively higher speed than the first turbine drum rotor 72. This enables the speed of rotation of the rearward fan 80 to be relatively low and optimised for the compression of the air flowing through the bypass duct 22 from the forward fan 50.
Seals must be provided at the upstream and downstream ends of the combustor to prevent leakage of working fluid into the bypass duct. The seals between the stationary combustor and the first rotor of the compressor or the first turbine rotor may be brush seals, or other suitable seals, as the first rotor of the compressor and the first turbine rotor are rotating at relatively low speeds.
The contra-rotating fans allow lower speeds of rotation to be achieved compared to a single fan for the same pressure ratio, so as to give lower tip speeds which reduce the noise generated and minimises blade containment problems.
The compressor does not require variable geometry vanes or blades at high pressure ratios to control the airflow because the speed of rotation of the first rotor varies. The strut 66 may be spaced two or more chord widths downstream of the forward fan for noise reduction purposes.
Fig. 2 shows a second embodiment of a gas turbine engine according to the present invention, which is similar to that shown in
Fig. 1, and like parts are demoted by like numerals. This embodiment differs in that the strut extending from the fan casing to the combustor has been removed, because the struts reduce efficiency and generate noise.
The combustor is therefor rotatably mounted onto the shaft 58 by the bearings 68 and 70, the downstream end of the first drum rotor 42 is rotatably mounted onto the shaft 58 by a bearing 94 and the upstream end of the first turbine drum rotor 72 is rotatably mounted onto the shaft 58 by a bearing 96.
In order to supply fuel to the combustor 16 a fuel system is provided which supplies fuel through piping within the vane 30 to a fuel pipe 98 which extends through cylindrical body 34. The fuel pipe 98 supplies fuel radially into a manifold 100 formed radially between the cylindrical body 34 and the shaft 58, and the two radially inwardly extending ribs 102,104. The shaft 58 has a number of apertures 106 to supply fuel to a second manifold 108 formed radially between the combustor 16 and the shaft 58, the manifold 108 supplies fuel to fuel injectors 110 which extend radially outwardly to the combustor head. Piping could be provided to extend through vane 36 and strut 40 to pipe 98.
Fig. 3 shows a third embodiment of a gas turbine engine according to the present invention, which is similar to that shown in Fig. 2 and like parts are denoted by like numerals.
This embodiment differs in that the downstream end of the combustor 16 is secured to or formed integral with the first turbine drum rotor 72 and is therefore arranged to rotate with the first turbine drum rotor 72, only one bearing 70 is required. This also overcomes the need for a seal between the combustor and the turbine.
Fig. 4 shows a fourth embodiment of a gas turbine according to the present invention, and is a turbo-propeller gas turbine engine, which is similar to that shown in Fig. 3 and like parts are shown by like numerals. This embodiment differs in that the turbo-propeller gas turbine engine 120 does not have a bypass duct and casing or forward and rearward fans, but the first rotor 42 of the compressor 14 drives a forward propeller 122 and the first turbine rotor 72 drives a rearward propeller 124. The propeller blades may be of the propfan type i.e. curved to reduce shock waves.
All the embodiments described above have a compact arrangement to produce a shorter engine which will have a reduced weight, and the fans or propellers are spaced further apart relatively for noise reduction. This is achieved by a novel drive system, in which a contrarotating turbine is used to drive a rearward fan, or propeller, and a second rotor of the compressor, and a first rotor of the compressor drives a forward fan, or propeller.
Figs. 5 and 6 show two further gas turbine arrangements according to the present invention, Fig. 5 and 6 show turbo-propeller gas turbine engines 140 and 160, which are similar to that shown in Fig. 4 and like parts are denoted by like numerals. The embodiment in
Fig. 5 differs in that the first rotor 42 is secured to and therefore driven by the first turbine 72 via a shaft or the casing of the combustor.
The first turbine rotor 72 therefore also drives the forward propeller 122 which is mounted on the first rotor 42. The forward propeller 122 could also be arranged on the forward row of blades 44 of the first rotor 42, or it could be arranged on the first turbine rotor 72. The second turbine rotor 82 drives the rearward propeller 124 which is positioned at the downstream end of the turbine, on the rearmost row of blades 84. The first rotor 42, the first turbine rotor 72 and the forward propeller 122 rotate in the opposite direction to the second rotor 52, the second turbine rotor 82 and the rearward propeller 124. The first turbine rotor and first rotor would rotate at a relatively slower speed than the second rotor and second turbine rotor. The first turbine rotor 72 is rotatably mounted onto the second turbine rotor 82 by bearing 78.
In Fig. 6 the first rotor 42 is secured to and driven by the first turbine rotor 72, and the first turbine rotor 72 also drives the rearward propeller 122, which is mounted on the first turbine rotor 72.
The second turbine rotor 82 drives the second rotor 52 and the forward propeller 122.
The foremost row of blades 54 of the second rotor 52 is secured to a casing 132, upon which the forward propeller 122 is mounted.
The first rotor 42 is rotatably mounted onto the second rotors 52 by the bearing 46. The first rotor 42, the first turbine rotor 72 and the rearward propeller 124 rotate in the opposite direction to the second rotor 52, the second turbine rotor 82 and the forward propeller 122.
It would be possible to have the same arrangements shown in Figs. 5 and 6, but in which the forward and rearward aerofoil arrangements are fans not propellers.
Prior art arrangements require a separate turbine to drive a compressor rotor, and a contra-rotating turbine to drive contra-rotating fans, or propellers.
The compressor does not require variable geometry vanes or blades at high pressure ratios because the speed of the first rotor varies.
Claims (13)
1. A gas turbine engine comprising a core engine, and a forward and a rearward aerofoil arrangement, the core engine comprising in flow series a compressor, a combustor and a turbine, the compressor comprising a first rotor having a plurality of circumferentially arranged first blades extending radially inwards therefrom, a second rotor spaced radially inwardly from the first rotor and having a plurality of circumferentially arranged second blades extending radially outwards therefrom, the first and second blades being arranged axially alternately, the first and second rotors being adapted to rotate in the same direction, the turbine comprising a first turbine rotor having a plurality of circumferentially arranged first turbine blades extending radially inwardly therefrom, a second turbine rotor spaced radially inwardly from the first turbine rotor and having a plurality of circumferentially arranged second turbine blades extending radially outwards therefrom, the first and second turbine rotors being adapted to contra-rotate, the second turbine rotor being adapted to drive the second rotor of the compressor, the forward aerofoil arrangement comprising a plurality of circumferentially arranged first aerofoil blades extending radially outwards from and being driven by the first rotor of the compressor, the rearward aerofoil arrangement comprising a plurality of circumferentially arranged second aerofoil blades extending radially outwards from and being driven by the first turbine rotor in contra-rotation to the forward aerofoil arrangement.
2. A gas turbine engine comprising a core engine, and a forward and a rearward aerofoil arrangement, the core engine comprising in flow series a compressor, a combustor and a turbine, the compressor comprising a first rotor having a plurality of circumferentially arranged first blades extending radially inwards therefrom, a second rotor having a plurality of circumferentially arranged second blades extending radially outwards therefrom, the first and second blades being arranged axially aiternately, the turbine comprising a first turbine rotor having a plurality of circumferentially arranged first turbine blades extending radially inwards therefrom, a second turbine rotor having a plurality of circumferentially arranged second turbine blades extending radially outwards therefrom, the first turbine being adapted to drive the first rotor of the compressor, the second turbine rotor being adapted to drive the second rotor of the compressor, the first and second turbine rotors being adapted to contra-rotate, the turbine being adapted to drive the forward aerofoil ar aerofoil a rangement and rearward aerofoil arrangement in opposite directions.
3. A gas turbine engine as claimed in claim 2 in which the forward aerofoil arrangement comprises a plurality of circumferentially arranged first aerofoil blades driven by the first turbine rotor, the rearward aerofoil arrangement comprises a plurality of circumferentially arranged second aerofoil blades driven by the second turbine rotor in the opposite direction to the forward aerofoil arrangement.
4. A gas turbine engine as claimed in claim 2 in which the forward aerofoil arrangement comprises a plurality of circumferentially arranged first aerofoil blades driven by the second turbine rotor, the rearward aerofoil arrangement comprises a plurality of circumferentially arranged second aerofoil blades driven by the first turbine rotor in the opposite direction to the forward aerofoil arrangement.
5. A gas turbine engine as claimed in any of claims 1 to 4 comprising a cylindrical body positioned coaxially within the second rotor of the compressor and the second turbine rotor, the cylindrical body extending axially from a first vane arrangement at the upstream end of the compressor to a second vane arrangement at the downstream end of the turbine, the first and second rotors of the compressor and the first and second turbine rotors being rotatably mounted on the cylindrical body.
6. A gas turbine engine as claimed in any of claims 1 to 5 in which the gas turbine engine is of the bypass type, the forward and rearward aerofoil arrangements are a forward and rearward fan arranged within a bypass duct.
7. A gas turbine engine as claimed in any of claims 1 to 6 in which the gas turbine engine is of the turbopropeller type, the forward and rearward aerofoil arrangements are a forward and rearward propeller.
8. A gas turbine engine as claimed in claim 6 in which the combustor is secured to a casing forming the outer boundary of the bypass duct by a number of struts extending across the bypass duct.
9. A gas turbine engine as claimed in any of claims 5 to 8 in which the combustor is rotatably mounted on the cylindrical body.
10. A gas turbine engine as claimed in claim 9 in which the combustor is secured to the upstream end of the first turbine rotor to rotate with the first turbine rotor.
11. A gas tubine engine as claimed in claim 9 or claim 10 comprising a fuel system having a fuel pipe extending radially through one of the vanes of the first vane arrangement or the second vane arrangement, the fuel pipe extending axially within the cylindrical body, at least one pipe extending radially to supply fuel to a fuel chamber formed radially between the cylindrical body and a shaft connecting the second rotor of the compressor to the second turbine rotor, the shaft having one or more apertures to supply fuel from the fuel chamber to a second fuel chamber and a plurality of circumferentially arranged radially outwards extending fuel injectors which supply fuel into the combustor.
12. A bypass gas turbine engine substantially as herein described with reference to and as shown in Figs. 1 to 3.
13. A turbopropeller gas turbine engine substantially as herein described with reference to and as shown in Figs. 4,5 and 6.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08610559A GB2189844A (en) | 1986-04-30 | 1986-04-30 | Gas turbine engines |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08610559A GB2189844A (en) | 1986-04-30 | 1986-04-30 | Gas turbine engines |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| GB8610559D0 GB8610559D0 (en) | 1986-11-26 |
| GB2189844A true GB2189844A (en) | 1987-11-04 |
Family
ID=10597108
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB08610559A Withdrawn GB2189844A (en) | 1986-04-30 | 1986-04-30 | Gas turbine engines |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2189844A (en) |
Cited By (16)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2230298A (en) * | 1989-01-03 | 1990-10-17 | Gen Electric | Geared counterrotating turbine/fan propulsion system |
| GB2192234B (en) * | 1986-07-02 | 1991-04-17 | Rolls Royce Plc | A turbofan gas turbine engine |
| US5058379A (en) * | 1989-04-18 | 1991-10-22 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation | High by-pass ratio turbojet engine with counterrotating upstream and downstream fans |
| EP0592817A1 (en) * | 1992-10-10 | 1994-04-20 | Asea Brown Boveri Ag | Gas turbine plant with a pressure wave machine |
| EP0753705A4 (en) * | 1993-08-08 | 1998-06-03 | Yanovsky Ilya Yakovlevich | Method of converting thermal energy to mechanical energy and a device for carrying out the same |
| WO2005008043A3 (en) * | 2003-07-23 | 2005-02-24 | Eugene Thomas David Archery | Tangentially exhausting revolving combustion chamber in a revolving ramjet, turbojet, turbofan, gas turbine or other jet engine |
| FR2866074A1 (en) * | 2004-02-11 | 2005-08-12 | Snecma Moteurs | High-bypass three-spool turbofan, has counter-rotating fans, and low pressure compressor with downstream rotating blades connecting rear fan to fan`s drive shaft and upstream rotating blades connected to drive shaft of front fan |
| EP1626002A1 (en) * | 2004-08-11 | 2006-02-15 | General Electric Company | Gas turbine engine turbine assembly |
| ES2342586A1 (en) * | 2008-04-30 | 2010-07-08 | Futur Investment Partners, S.A. | Turbopropulsor aeronautical (Machine-translation by Google Translate, not legally binding) |
| DE102014226696A1 (en) * | 2014-12-19 | 2016-06-23 | Rolls-Royce Deutschland Ltd & Co Kg | Pre-compressor device, retrofit kit and aircraft engine |
| US11118509B2 (en) * | 2017-06-01 | 2021-09-14 | Safran Aircraft Engines | Turbojet of the unducted rotor type |
| US11247780B2 (en) | 2018-08-22 | 2022-02-15 | Rolls-Royce Plc | Turbomachine having inner and outer fans with hub-tip ratios |
| US11306682B2 (en) | 2018-08-22 | 2022-04-19 | Rolls-Royce Plc | Concentric turbomachine with trailing edge |
| US11313327B2 (en) | 2018-08-22 | 2022-04-26 | Rolls-Royce Plc | Concentric turbomachine with electric machine |
| US11371467B2 (en) | 2018-08-22 | 2022-06-28 | Rolls-Royce Plc | Concentric turbomachine with electric machine |
| US11371350B2 (en) | 2018-08-22 | 2022-06-28 | Rolls-Royce Plc | Concentric turbomachine with electric machine |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1321657A (en) * | 1970-09-16 | 1973-06-27 | Secr Defence | Ducted fan gas turbine jet propulsion engine |
| GB2129502A (en) * | 1982-11-01 | 1984-05-16 | Gen Electric | Counter rotation power turbine |
-
1986
- 1986-04-30 GB GB08610559A patent/GB2189844A/en not_active Withdrawn
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1321657A (en) * | 1970-09-16 | 1973-06-27 | Secr Defence | Ducted fan gas turbine jet propulsion engine |
| GB2129502A (en) * | 1982-11-01 | 1984-05-16 | Gen Electric | Counter rotation power turbine |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2192234B (en) * | 1986-07-02 | 1991-04-17 | Rolls Royce Plc | A turbofan gas turbine engine |
| GB2230298B (en) * | 1989-01-03 | 1993-09-22 | Gen Electric | Geared counterrotating turbine/fan propulsion system |
| GB2230298A (en) * | 1989-01-03 | 1990-10-17 | Gen Electric | Geared counterrotating turbine/fan propulsion system |
| US5058379A (en) * | 1989-04-18 | 1991-10-22 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation | High by-pass ratio turbojet engine with counterrotating upstream and downstream fans |
| EP0592817A1 (en) * | 1992-10-10 | 1994-04-20 | Asea Brown Boveri Ag | Gas turbine plant with a pressure wave machine |
| EP0753705A4 (en) * | 1993-08-08 | 1998-06-03 | Yanovsky Ilya Yakovlevich | Method of converting thermal energy to mechanical energy and a device for carrying out the same |
| WO2005008043A3 (en) * | 2003-07-23 | 2005-02-24 | Eugene Thomas David Archery | Tangentially exhausting revolving combustion chamber in a revolving ramjet, turbojet, turbofan, gas turbine or other jet engine |
| US7412819B2 (en) | 2004-02-11 | 2008-08-19 | Snecma | Turbojet architecture with two fans at the front |
| FR2866074A1 (en) * | 2004-02-11 | 2005-08-12 | Snecma Moteurs | High-bypass three-spool turbofan, has counter-rotating fans, and low pressure compressor with downstream rotating blades connecting rear fan to fan`s drive shaft and upstream rotating blades connected to drive shaft of front fan |
| EP1564397A1 (en) * | 2004-02-11 | 2005-08-17 | Snecma Moteurs | Jet engine architecture having two fans at the front end |
| JP2005226643A (en) * | 2004-02-11 | 2005-08-25 | Snecma Moteurs | Turbojet structure with two fans at the front |
| RU2357092C2 (en) * | 2004-02-11 | 2009-05-27 | Снекма | Design of turbojet engine with doubled fan in front part |
| EP1626002A1 (en) * | 2004-08-11 | 2006-02-15 | General Electric Company | Gas turbine engine turbine assembly |
| US7185484B2 (en) | 2004-08-11 | 2007-03-06 | General Electric Company | Methods and apparatus for assembling a gas turbine engine |
| ES2342586A1 (en) * | 2008-04-30 | 2010-07-08 | Futur Investment Partners, S.A. | Turbopropulsor aeronautical (Machine-translation by Google Translate, not legally binding) |
| ES2342586B1 (en) * | 2008-04-30 | 2011-06-14 | Futur Investment Partners, S.A. | AERONAUTICAL TURBOPROPULSOR. |
| DE102014226696A1 (en) * | 2014-12-19 | 2016-06-23 | Rolls-Royce Deutschland Ltd & Co Kg | Pre-compressor device, retrofit kit and aircraft engine |
| US11118509B2 (en) * | 2017-06-01 | 2021-09-14 | Safran Aircraft Engines | Turbojet of the unducted rotor type |
| US11247780B2 (en) | 2018-08-22 | 2022-02-15 | Rolls-Royce Plc | Turbomachine having inner and outer fans with hub-tip ratios |
| US11306682B2 (en) | 2018-08-22 | 2022-04-19 | Rolls-Royce Plc | Concentric turbomachine with trailing edge |
| US11313327B2 (en) | 2018-08-22 | 2022-04-26 | Rolls-Royce Plc | Concentric turbomachine with electric machine |
| US11371467B2 (en) | 2018-08-22 | 2022-06-28 | Rolls-Royce Plc | Concentric turbomachine with electric machine |
| US11371350B2 (en) | 2018-08-22 | 2022-06-28 | Rolls-Royce Plc | Concentric turbomachine with electric machine |
Also Published As
| Publication number | Publication date |
|---|---|
| GB8610559D0 (en) | 1986-11-26 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| WAP | Application withdrawn, taken to be withdrawn or refused ** after publication under section 16(1) |