US3121526A - Gas turbine engines - Google Patents
Gas turbine engines Download PDFInfo
- Publication number
- US3121526A US3121526A US128173A US12817361A US3121526A US 3121526 A US3121526 A US 3121526A US 128173 A US128173 A US 128173A US 12817361 A US12817361 A US 12817361A US 3121526 A US3121526 A US 3121526A
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- Prior art keywords
- shaft
- compressor
- turbine
- fan
- gas turbine
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Classifications
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- 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
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
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- 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
- F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
- F05D2220/327—Application in turbines in gas turbines to drive shrouded, high solidity propeller
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/40—Transmission of power
- F05D2260/403—Transmission of power through the shape of the drive components
- F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
Definitions
- GAS TURBINE ENGINES Filed July 31, 1961 'n 17X 2/ i [g H ,1 2 ll /5 /6 25 22 22 Q 11 J g '2 r 22 22 I29 I 1 4 24 o 2/ A 5/ 5 9 lfia lnvenlor y MWMW r M United States Patent 3,121,526
- This invention relates to gas turbine engines and in particular, but not exclusively, to gas turbine engines having a single shaft connecting a turbine or turbines to the compressor or compressors.
- One of the objects of the present invention is to produce a gas turbine engine having a low pressure compressor or fan driven from the engine rotor shaft through reduction gearing, but not having the disadvantages associated with systems employing lay-shafts and spur gears.
- a gas turbine engine comprises at least a compressor, a turbine and a first shaft interconnecting the compressor and turbine, the engine also comprises at least one additional shaft mounted compressor or fan driven from said first shaft through reduction gearing, not employing a lay shaft, the longitudinal axis of the shaft of said additional oompressor or fan being displaced laterally with respect to the longitudinal axis of said first shaft.
- the reduction gearing comprises a pinion driven from the first shaft in driving engagement with an internally toothed annular gear driving said shaft of said additional compressor or fan, said annular gear having a larger internal diameter than the external diameter of said pinion.
- the pinion is formed integrally around or is mounted upon the front end of said first shaft and said annular gear forms part of the downstream end of said additional compressor or fan rotor.
- the pinion is mounted on a separate shaft having its longitudinal axis concentric with the longitudinal axis of the first shaft, said separate shaft being driven from said first shaft by coupling means.
- the additional compressor or fan is provided with a downstream hearing which is located in the plane of the gear teeth and the said first shaft or ice said separate shaft is provided with an upstream bearing located close to said downstream hearing.
- the gear teeth of the reduction gearing are preferably of helical form.
- some of the air compressed in the additional compressor or fan passes to a high pressure compressor, combustion equipment and then through a turbine to drive it before exhausting to atmosphere, the remainder of the compressed air being conveyed in a duct so as to bypass the combustion equipment and turbine and being mixed with the hot gases from the turbine downstream of said turbine, said additional compressor or fan having its rotational axis displaced laterally with respect to the rotational axis of said high pressure compressor and being driven from the high pressure compressor rotor through reduction gearing.
- FIGURE 1 shows a diagrammatic section through a gas turbine engine embodying the present invention
- FIGURE 2 is a more detailed view of the reduction gear employed in driving the low pressure compressor fan
- FIGURE 3 is a section taken on the line 33 indicated on FIGURE 2, and
- FIGURE 4 is a diagrammatic section through a two shaft engine embodying the invention.
- a gas turbine engine 10 comprises a high pressure compressor 11 which is driven from a turbine 12 through a first shaft 13.
- the shaft 13 is arranged to drive a lowpressure compressor or fan 14 through a reduction gear 16.
- Air compressed in the low pressure compressor or fan 14 is divided into two flow paths, one of which flows into the high pressure compressor 11 Where the air is further compressed before passing into combustion equipment 15 where fuel is burned in the air.
- the hot gases from the combustion equipment pass through the turbine 12 to drive it and are discharged to atmosphere as a pro pulsive jet through a propelling nozzle 17.
- the remainder of the air compressed in the low pressure compressor or fan 14 is conveyed to a by-pass duct 18 and is injected into the hot propulsive gas stream through mixer chutes '19 located downstream of the turbine 12.
- the upstream end of the inner wall of the bypass duct 18 is provided with a lip 18a which may be shaped circumferentially to ensure an even entry of air into the bypass duct 18.
- the reduction gear 16 is shown more cleanly in FIG- URES 2 and 3 and comprises an internally toothed annular gear 20 which forms the downstream end of a rotor shaft 21 upon which are mounted the rotor blades of the low pressure compressor or fan 14.
- the rotor shaft 21 is mounted in a thrust bearing 22 at its upstream end and a roller bearing 23 positioned in the plane of the annular gear 20.
- the longitudinal axis of the rotor shaft 21 is indicated at 24.
- a pinion 25 In driving engagement with the annular gear 20 is a pinion 25 which is formed or mounted on the upstream end of a shaft 26.
- the shaft '26 is mounted in a roller bearing 27 and a thrust bearing 28 and has its longitudinal axis concentric with the longitudinal axis of the first shaft 13, said axis being indicated at 29.
- the shaft 26 is driven from the first shaft 13 by means of a connecting member 30 having sets of splines which engage with sets of splines for-med on the internal surface of shafts 26 and 13.
- the pinion 25 could 3 be formed or mounted on the upstream end of the first shaft 13.
- the air compressed in the low pressure compressor or fan 14' is disch-argedinto a duct 31 which at its upstream end is concentric with the low pressure compressor or fan 14 and at its downstream end is concentric with the high pressure compressor 11.
- the rotational speed of the first shaft 113 will be governed so as to prevent the tip speeds of the rotor blades of the high pressure compressor 11 exceeding a predetermined value.
- This rotational speed will be reduced by the reduction gear 16 so that the rotational speed of the rotor shafts 21 will not cause the tip speeds of the rotor blades of the low pressure compressor or fan 14- to exceed a predetermined value.
- the bearings 23, 27 are mounted as close to the plane of the annular gear 20 and the pinion 25 as is possible.
- the gear teeth of the annular gear and pinion 25 are of helical form.
- FIGURE 4 is shown a two shaft gas turbine engine having a high pressure turbine 32 driving a high pressure compressor 33 through a shaft 34 and a low pressure turbine 35 driving an intermediate compressor 36 through a shaft 37.
- the shaft 37 is mounted concentrically within the shaft 34 and also drives a fan or low pressure compressor 38 through a reduction gear 39 in accordance with the present invention.
- a first compressor means and turbine means arranged in series and having a common axis of rotation, said compressor means and said turbine means having fluid flowing therethrough in the same axial direction; shaft means rotatable on said common axis of rotation and drivingly connecting said compressor means to said turbine means; a second compressor means arranged in series with said first compressor means and said turbine means, said second compressor means having fluid flowing therethrough in an axial direction substantially the same as the flow of fluid through said first compressor means and said turbine means, said second compressor means having a maximum diameter greater than the maximum diameter of said first compressor means, said second compressor means having a drive shaft and an axis of rotation displaced laterally from and parallel to the common axis of rotation of said first compressor means and said turbine means; and reduction gearing means drivingly connecting said shaft means of said first compressor means and turbine means to said drive shaft of said second compressor means for driving said second compressor means at a slower speed than said first compressor means, said reduction gearing means including a drive pinion gear rotated by said shaft means
- said shaft means includes a first shaft connected to said first compressor means and said turbine means and a second shaft extending forwardly of said first shaft on a longitudinal axis of said first shaft, and coupling means coupling said first shaft and said second shaft to each other, said pinion being mounted at the forward end of said second shaft.
- compressor means and turbine means arranged in series and having a common axis of rotation, said compressor means and said turbine means having fluid flowing therethrough in the same axial direction; shaft means drivingly connecting said compressor means to said turbine means; a multi-stage axiallflow fan arranged in series with said compressor means and said turbine means, said multi-stage axial-flow fan having an axis of rotation displaced laterally from and parallel to the common axis of rotation of said compressor means and said turbine means said multi-stage axial flow fan having fluid flowing therethrough in an axial direction substantially the same as the flow of fluid through said compressor means and said turbine means, and said multi-stage fan having a maximum diameter greater than the diameter of said compressor means; and means drivingly connecting said multi-stage axial-flow fan to said shaft means.
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- 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)
Description
Feb. 18, 1964 F. w. w. MORLEY 3,121,526
GAS TURBINE ENGINES Filed July 31, 1961 'n 17X 2/ i [g H ,1 2 ll /5 /6 25 22 22 Q 11 J g '2 r 22 22 I29 I 1 4 24 o 2/ A 5/ 5 9 lfia lnvenlor y MWMW r M United States Patent 3,121,526 GAS TURBINE ENGINES Frederick Wiliiarn Walton Morley, Castle Donington, Engiaud, assignor to Rolls-Royce Limited, Derby, England Filed July 31, 1%1, Ser. No. 128,173 Claims priority, application Great Britain Aug. 25, 1960 7 Ciainrs. (Cl. 230-416) This invention relates to gas turbine engines and in particular, but not exclusively, to gas turbine engines having a single shaft connecting a turbine or turbines to the compressor or compressors.
In single shaft gas turbine engines it is important to maintain the tip speeds of the compressor blades within predetermined limits. Where a single shaft engine is designed to rotate a high pressure compressor and a low pressure compressor or fan the diameter of the blade tips of the low pressure compressor or fan is usually much larger than the diameter of the blade tips of the high pressure compressor, therefore, if the rotational speed of the engine rotor shaft is chosen to give the correct tip speed of the high pressure compressor rotor blades then the tip speeds of the low pressure compressor or fan rotor blades becomes too great, resulting in an increase in noise and a loss in efficiency.
It is, therefore, desirable to rotate the low pressure compressor or fan at a much lower rotational speed in order to obtain the correct tip speeds of the rotor blades.
It has previously been proposed to drive the low pressure compressor or fan from the engine rotor shaft through reduction gearing comprising a system of spur gears mounted on lay-shafts, both the engine rotor shaft and the low pressure compressor or fan rotor shaft being mounted on the same rotational axis.
Such gear systems suffer from the disadvantages of low mechanical efficiency and high bearing loading usually associated with systems employing lay-shafts and spur gears.
One of the objects of the present invention is to produce a gas turbine engine having a low pressure compressor or fan driven from the engine rotor shaft through reduction gearing, but not having the disadvantages associated with systems employing lay-shafts and spur gears.
According to the present invention a gas turbine engine comprises at least a compressor, a turbine and a first shaft interconnecting the compressor and turbine, the engine also comprises at least one additional shaft mounted compressor or fan driven from said first shaft through reduction gearing, not employing a lay shaft, the longitudinal axis of the shaft of said additional oompressor or fan being displaced laterally with respect to the longitudinal axis of said first shaft.
Preferably the reduction gearing comprises a pinion driven from the first shaft in driving engagement with an internally toothed annular gear driving said shaft of said additional compressor or fan, said annular gear having a larger internal diameter than the external diameter of said pinion.
In one arrangement the pinion is formed integrally around or is mounted upon the front end of said first shaft and said annular gear forms part of the downstream end of said additional compressor or fan rotor.
In an alternative arrangement the pinion is mounted on a separate shaft having its longitudinal axis concentric with the longitudinal axis of the first shaft, said separate shaft being driven from said first shaft by coupling means.
Preferably the additional compressor or fan is provided with a downstream hearing which is located in the plane of the gear teeth and the said first shaft or ice said separate shaft is provided with an upstream bearing located close to said downstream hearing.
The gear teeth of the reduction gearing are preferably of helical form.
In a preferred arrangement some of the air compressed in the additional compressor or fan passes to a high pressure compressor, combustion equipment and then through a turbine to drive it before exhausting to atmosphere, the remainder of the compressed air being conveyed in a duct so as to bypass the combustion equipment and turbine and being mixed with the hot gases from the turbine downstream of said turbine, said additional compressor or fan having its rotational axis displaced laterally with respect to the rotational axis of said high pressure compressor and being driven from the high pressure compressor rotor through reduction gearing.
One embodiment of the present invention will now be described with reference to the drawings accompanying the sepcification and claims in which:
FIGURE 1 shows a diagrammatic section through a gas turbine engine embodying the present invention,
FIGURE 2 is a more detailed view of the reduction gear employed in driving the low pressure compressor fan,
FIGURE 3 is a section taken on the line 33 indicated on FIGURE 2, and
FIGURE 4 is a diagrammatic section through a two shaft engine embodying the invention.
A gas turbine engine 10 comprises a high pressure compressor 11 which is driven from a turbine 12 through a first shaft 13. The shaft 13 is arranged to drive a lowpressure compressor or fan 14 through a reduction gear 16.
Air compressed in the low pressure compressor or fan 14 is divided into two flow paths, one of which flows into the high pressure compressor 11 Where the air is further compressed before passing into combustion equipment 15 where fuel is burned in the air. The hot gases from the combustion equipment pass through the turbine 12 to drive it and are discharged to atmosphere as a pro pulsive jet through a propelling nozzle 17.
The remainder of the air compressed in the low pressure compressor or fan 14 is conveyed to a by-pass duct 18 and is injected into the hot propulsive gas stream through mixer chutes '19 located downstream of the turbine 12. The upstream end of the inner wall of the bypass duct 18 is provided with a lip 18a which may be shaped circumferentially to ensure an even entry of air into the bypass duct 18.
The reduction gear 16 is shown more cleanly in FIG- URES 2 and 3 and comprises an internally toothed annular gear 20 which forms the downstream end of a rotor shaft 21 upon which are mounted the rotor blades of the low pressure compressor or fan 14. The rotor shaft 21 is mounted in a thrust bearing 22 at its upstream end and a roller bearing 23 positioned in the plane of the annular gear 20. The longitudinal axis of the rotor shaft =21 is indicated at 24.
In driving engagement with the annular gear 20 is a pinion 25 which is formed or mounted on the upstream end of a shaft 26. The shaft '26 is mounted in a roller bearing 27 and a thrust bearing 28 and has its longitudinal axis concentric with the longitudinal axis of the first shaft 13, said axis being indicated at 29.
It will be seen that the longitudinal axis 24- of the rotor shaft 2 1 is displaced laterally with respect to the longiuudinal axis 29 of the first shaft 13.
The shaft 26 is driven from the first shaft 13 by means of a connecting member 30 having sets of splines which engage with sets of splines for-med on the internal surface of shafts 26 and 13.
In an alternative construction the pinion 25 could 3 be formed or mounted on the upstream end of the first shaft 13.
The air compressed in the low pressure compressor or fan 14' is disch-argedinto a duct 31 which at its upstream end is concentric with the low pressure compressor or fan 14 and at its downstream end is concentric with the high pressure compressor 11.
During operation of the gas turbine engine the rotational speed of the first shaft 113 will be governed so as to prevent the tip speeds of the rotor blades of the high pressure compressor 11 exceeding a predetermined value.
This rotational speed will be reduced by the reduction gear 16 so that the rotational speed of the rotor shafts 21 will not cause the tip speeds of the rotor blades of the low pressure compressor or fan 14- to exceed a predetermined value.
The bearings 23, 27 are mounted as close to the plane of the annular gear 20 and the pinion 25 as is possible.
The gear teeth of the annular gear and pinion 25 are of helical form.
In FIGURE 4 is shown a two shaft gas turbine engine having a high pressure turbine 32 driving a high pressure compressor 33 through a shaft 34 and a low pressure turbine 35 driving an intermediate compressor 36 through a shaft 37.
The shaft 37 is mounted concentrically within the shaft 34 and also drives a fan or low pressure compressor 38 through a reduction gear 39 in accordance with the present invention.
The remaining parts of the two shaft gas turbine engine are identical to those shown in FIGURE 1 and have been indicated by similar reference numerals.
What I claim is:
1. In a substantially axial flow gas turbine engine: a first compressor means and turbine means arranged in series and having a common axis of rotation, said compressor means and said turbine means having fluid flowing therethrough in the same axial direction; shaft means rotatable on said common axis of rotation and drivingly connecting said compressor means to said turbine means; a second compressor means arranged in series with said first compressor means and said turbine means, said second compressor means having fluid flowing therethrough in an axial direction substantially the same as the flow of fluid through said first compressor means and said turbine means, said second compressor means having a maximum diameter greater than the maximum diameter of said first compressor means, said second compressor means having a drive shaft and an axis of rotation displaced laterally from and parallel to the common axis of rotation of said first compressor means and said turbine means; and reduction gearing means drivingly connecting said shaft means of said first compressor means and turbine means to said drive shaft of said second compressor means for driving said second compressor means at a slower speed than said first compressor means, said reduction gearing means including a drive pinion gear rotated by said shaft means on its axis of rotation, and an annular gear on said drive shaft and having internal teeth meshing with said pinion gear, said pinion gear having a smaller external diameter than the internal diameter of said annular gear.
2. The gas turbine engine as claimed in claim 1 in which said pinion is mounted at the upstream end of said shaft means and said annular gear is mounted at the downstream end of said drive shaft of said second compressor means.
3. The gas turbine engine as claimed in claim 1 in which said shaft means includes a first shaft connected to said first compressor means and said turbine means and a second shaft extending forwardly of said first shaft on a longitudinal axis of said first shaft, and coupling means coupling said first shaft and said second shaft to each other, said pinion being mounted at the forward end of said second shaft.
4. The gas turbine engine as claimed in claim 2 ineluding a bearing for the drive shaft of said first compressor means, said bearing being positioned at the downstream end of said drive shaft in the plane of said annular gear, and a bearing for said shaft means positioned at the upstream end thereof immediately adjacent the bearing for said drive shaft of said second compressor means.
5. The gas turbine engine as claimed in claim 1 in which said internal gear teeth of said annular gear and the gear teeth of said pinion gear are helical gear teeth.
6. In a gas turbine engine: compressor means and turbine means arranged in series and having a common axis of rotation, said compressor means and said turbine means having fluid flowing therethrough in the same axial direction; shaft means drivingly connecting said compressor means to said turbine means; a multi-stage axiallflow fan arranged in series with said compressor means and said turbine means, said multi-stage axial-flow fan having an axis of rotation displaced laterally from and parallel to the common axis of rotation of said compressor means and said turbine means said multi-stage axial flow fan having fluid flowing therethrough in an axial direction substantially the same as the flow of fluid through said compressor means and said turbine means, and said multi-stage fan having a maximum diameter greater than the diameter of said compressor means; and means drivingly connecting said multi-stage axial-flow fan to said shaft means.
7. The gas turbine engine as claimed in claim 6 wherein said multi-stage axial-flow fan is positioned on the inlet side of said compressor means.
References Cited in the file of this patent UNITED STATES PATENTS 2,803,943 Rainbow Aug. 27, 1957 2,811,302 Hodge et a1 Oct. 29, 1957 3,025,672 Syrovy Mar. 20; 1962 FOREIGN PATENTS 586,567 Great Britain Mar. 24, 1947 698,783 Great Britain Oct. 21, 1953
Claims (1)
1. IN A SUBSTANTIALLY AXIAL FLOW GAS TURBINE ENGINE: A FIRST COMPRESSOR MEANS AND TURBINE MEANS ARRANGED IN SERIES AND HAVING A COMMON AXIS OF ROTATION, SAID COMPRESSOR MEANS AND SAID TURBINE MEANS HAVING FLUID FLOWING THERETHROUGH IN THE SAME AXIAL DIRECTION; SHAFT MEANS ROTATABLE ON SAID COMMON AXIS OF ROTATION AND DRIVINGLY CONNECTING SAID COMPRESSOR MEANS TO SAID TURBINE MEANS; A SECOND COMPRESSOR MEANS ARRANGED IN SERIES WITH SAID FIRST COMPRESSOR MEANS AND SAID TURBINE MEANS, SAID SECOND COMPRESSOR MEANS HAVING FLUID FLOWING THERETHROUGH IN AN AXIAL DIRECTION SUBSTANTIALLY THE SAME AS THE FLOW OF FLUID THROUGH SAID FIRST COMPRESSOR MEANS AND SAID TURBINE MEANS, SAID SECOND COMPRESSOR MEANS HAVING A MAXIMUM DIAMETER GREATER THAN THE MAXIMUM DIAMETER OF SAID FIRST COMPRESSOR MEANS, SAID SECOND COMPRESSOR
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB3121526X | 1960-08-25 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| US3121526A true US3121526A (en) | 1964-02-18 |
Family
ID=10922141
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| US128173A Expired - Lifetime US3121526A (en) | 1960-08-25 | 1961-07-31 | Gas turbine engines |
Country Status (1)
| Country | Link |
|---|---|
| US (1) | US3121526A (en) |
Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4038816A (en) * | 1974-07-16 | 1977-08-02 | Wright Charles H | Rotary engine and turbine assembly |
| US5313784A (en) * | 1992-10-15 | 1994-05-24 | Hughes Aircraft Company | Solid fuel pinwheel power plant and method |
| US5553448A (en) * | 1992-05-14 | 1996-09-10 | General Electric Company | Intercooled gas turbine engine |
| FR2759734A1 (en) * | 1997-02-20 | 1998-08-21 | Snecma | TURBOMACHINE WITH OPTIMIZED COMPRESSION SYSTEM |
| WO2003067054A2 (en) * | 2002-01-17 | 2003-08-14 | Vericor Power Systems Llc | Hybrid lean premixing catalytic combustion system for gas turbines |
| US20090092487A1 (en) * | 2007-10-09 | 2009-04-09 | United Technologies Corp. | Systems and Methods Involving Multiple Torque Paths for Gas Turbine Engines |
| WO2013180762A1 (en) * | 2012-01-27 | 2013-12-05 | United Technologies Corporation | Thrust balance system for gas turbine engine |
| US20160363055A1 (en) * | 2015-06-09 | 2016-12-15 | Rolls-Royce North American Technologies, Inc. | Off-set geared turbofan engine |
| WO2017193035A1 (en) * | 2016-05-06 | 2017-11-09 | General Electric Company | Gas turbine power generation system with a high pressure compressor and an added forward low pressure compressor |
| US9850821B2 (en) | 2012-02-28 | 2017-12-26 | United Technologies Corporation | Gas turbine engine with fan-tied inducer section |
| US20180105278A1 (en) * | 2016-10-14 | 2018-04-19 | Safran Aircraft Engines | Turbine engine having horizontally offset axes |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB586567A (en) * | 1942-11-12 | 1947-03-24 | Karl Baumann | Improvements in axial flow compressors |
| GB698783A (en) * | 1951-12-20 | 1953-10-21 | Inconex Handelsgesellschaft M | Compressor apparatus |
| US2803943A (en) * | 1953-12-30 | 1957-08-27 | Armstrong Siddeley Motors Ltd | Means for supporting and driving accessories which are exterior to a ductedfan turbo-jet engine |
| US2811302A (en) * | 1954-02-24 | 1957-10-29 | Power Jets Res & Dev Ltd | Gas turbine plant and control arrangements therefor |
| US3025672A (en) * | 1959-07-21 | 1962-03-20 | Gen Motors Corp | Engine accessory installation |
-
1961
- 1961-07-31 US US128173A patent/US3121526A/en not_active Expired - Lifetime
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB586567A (en) * | 1942-11-12 | 1947-03-24 | Karl Baumann | Improvements in axial flow compressors |
| GB698783A (en) * | 1951-12-20 | 1953-10-21 | Inconex Handelsgesellschaft M | Compressor apparatus |
| US2803943A (en) * | 1953-12-30 | 1957-08-27 | Armstrong Siddeley Motors Ltd | Means for supporting and driving accessories which are exterior to a ductedfan turbo-jet engine |
| US2811302A (en) * | 1954-02-24 | 1957-10-29 | Power Jets Res & Dev Ltd | Gas turbine plant and control arrangements therefor |
| US3025672A (en) * | 1959-07-21 | 1962-03-20 | Gen Motors Corp | Engine accessory installation |
Cited By (21)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4038816A (en) * | 1974-07-16 | 1977-08-02 | Wright Charles H | Rotary engine and turbine assembly |
| US5553448A (en) * | 1992-05-14 | 1996-09-10 | General Electric Company | Intercooled gas turbine engine |
| US5313784A (en) * | 1992-10-15 | 1994-05-24 | Hughes Aircraft Company | Solid fuel pinwheel power plant and method |
| FR2759734A1 (en) * | 1997-02-20 | 1998-08-21 | Snecma | TURBOMACHINE WITH OPTIMIZED COMPRESSION SYSTEM |
| EP0860593A1 (en) * | 1997-02-20 | 1998-08-26 | SOCIETE NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION -Snecma | Compression system for a turbomachine |
| WO2003067054A2 (en) * | 2002-01-17 | 2003-08-14 | Vericor Power Systems Llc | Hybrid lean premixing catalytic combustion system for gas turbines |
| WO2003067054A3 (en) * | 2002-01-17 | 2003-11-20 | Vericor Power Systems Llc | Hybrid lean premixing catalytic combustion system for gas turbines |
| US6658856B2 (en) * | 2002-01-17 | 2003-12-09 | Vericor Power Systems Llc | Hybrid lean premixing catalytic combustion system for gas turbines |
| US8266886B2 (en) | 2007-10-09 | 2012-09-18 | United Technologies Corp. | Systems and methods involving multiple torque paths for gas turbine engines |
| US8104289B2 (en) * | 2007-10-09 | 2012-01-31 | United Technologies Corp. | Systems and methods involving multiple torque paths for gas turbine engines |
| US20090092487A1 (en) * | 2007-10-09 | 2009-04-09 | United Technologies Corp. | Systems and Methods Involving Multiple Torque Paths for Gas Turbine Engines |
| US8621871B2 (en) | 2007-10-09 | 2014-01-07 | United Technologies Corporation | Systems and methods involving multiple torque paths for gas turbine engines |
| WO2013180762A1 (en) * | 2012-01-27 | 2013-12-05 | United Technologies Corporation | Thrust balance system for gas turbine engine |
| US9850821B2 (en) | 2012-02-28 | 2017-12-26 | United Technologies Corporation | Gas turbine engine with fan-tied inducer section |
| US20160363055A1 (en) * | 2015-06-09 | 2016-12-15 | Rolls-Royce North American Technologies, Inc. | Off-set geared turbofan engine |
| EP3106653A1 (en) * | 2015-06-09 | 2016-12-21 | Rolls-Royce North American Technologies, Inc. | Off-set geared turbofan engine |
| WO2017193035A1 (en) * | 2016-05-06 | 2017-11-09 | General Electric Company | Gas turbine power generation system with a high pressure compressor and an added forward low pressure compressor |
| CN109415975A (en) * | 2016-05-06 | 2019-03-01 | 通用电气公司 | With high pressure compressor and the additional preceding gas turbine power generation system to low pressure compressor |
| US20180105278A1 (en) * | 2016-10-14 | 2018-04-19 | Safran Aircraft Engines | Turbine engine having horizontally offset axes |
| FR3057543A1 (en) * | 2016-10-14 | 2018-04-20 | Safran Aircraft Engines | TURBOMACHINE WITH HORIZONTALLY DECAL AXES |
| US10640222B2 (en) * | 2016-10-14 | 2020-05-05 | Safran Aircraft Engines | Turbine engine having horizontally offset axes |
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