WO1999036688A1 - Gas turbine engine - Google Patents

Gas turbine engine Download PDF

Info

Publication number
WO1999036688A1
WO1999036688A1 PCT/GB1999/000074 GB9900074W WO9936688A1 WO 1999036688 A1 WO1999036688 A1 WO 1999036688A1 GB 9900074 W GB9900074 W GB 9900074W WO 9936688 A1 WO9936688 A1 WO 9936688A1
Authority
WO
WIPO (PCT)
Prior art keywords
gas turbine
turbine engine
diffuser
variable
engine
Prior art date
Application number
PCT/GB1999/000074
Other languages
French (fr)
Inventor
Philip Patrick Walsh
Paul Fletcher
James Melville
Original Assignee
Rolls-Royce Plc
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Priority to GB9800782.6 priority Critical
Priority to GBGB9800782.6A priority patent/GB9800782D0/en
Application filed by Rolls-Royce Plc filed Critical Rolls-Royce Plc
Publication of WO1999036688A1 publication Critical patent/WO1999036688A1/en

Links

Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16Control of working fluid flow
    • F02C9/20Control of working fluid flow by throttling; by adjusting vanes
    • F02C9/22Control of working fluid flow by throttling; by adjusting vanes by adjusting turbine vanes
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00Regulating or controlling by varying flow
    • F01D17/10Final actuators
    • F01D17/12Final actuators arranged in stator parts
    • F01D17/14Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/08Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/16Control of working fluid flow
    • F02C9/20Control of working fluid flow by throttling; by adjusting vanes

Abstract

A gas turbine engine comprises a centrifugal compressor (4), an air diffuser (8), a heat exchanger (10), combustion apparatus (12), and first and second turbines (14, 18). The compressor (4), diffuser (8) and turbines (14, 18) all comprise means (6, 8, 16, 20) for varying the mass flow area at their inlets such that in operation the amount of air mass through each component may be independently variable. Under part power conditions the mass flow is reduced and under full power conditions the mass flow is increased thereby maintaining a substantially constant gas cycle throughout the engine.

Description

Gas Turbine Engine
This invention relates to gas turbine engines and in particular relates to gas turbine engines for non-aero applications . One main consideration for the operation of gas turbine engines is the specific fuel consumption value (SFC) , measured in Kg/kWhrs . In general for certain gas turbine engine applications especially marine, automotive and even industrial, a significant proportion of operation is at low power. A gas turbine engine utilises hot working fluid expanding through a given expansion ratio in the turbines which produces a power in excess of that required for the compressor to produce the corresponding pressure ratio. This is due to pressure and temperature ratios being proportional to one another during compression or expansion in the simple gas turbine engine cycle, which means that temperature change, and hence work, is proportional to the initial temperature level . Therefore reducing the amount of fuel available at part power results in reduced temperature levels and hence a reduced speed and pressure ratio thus resulting in a significant increase in specific fuel consumption (SFC) .
Recuperated gas turbine engines use heat exchangers to return heat from the final turbine exhaust to pre-heat compressed air entering the combustor. This helps to conserve fuel by raising the combustor air temperature and therefore limiting the amount of fuel needed to achieve the turbine inlet temperature .
It is also known to provide a power turbine with variable area nozzles (VANs) to improve SFC at part power conditions . Such an arrangement of VANs is disclosed in GB2301868 and GB application No 9511269.4.
It is also known to provide a gas turbine compressor arrangement where the compressor vane angles are varied so as to alter the flow area . One such arrangement of variable compressor guide vanes is disclosed in GB2210108. Other arrangements for varying the mass flow through a gas turbine engine are disclosed in US3138923 and US3025688 and US4145875. These prior art patents disclose variable geometry arrangements in the form of angled vanes positioned to change the incident angle of gas flow with respect to either diffuser passages or nozzle passages.
As mentioned previously when the gas turbine engine is operated at low power the reduced levels of pressure ratio and temperature result in increased fuel consumption. It is desirable therefore that the gas turbine engine cycle is optimised at low power so that fuel consumption is reduced. In the prior art, for recuperated gas turbine engine cycles the variable area nozzle is closed as power is reduced maintaining the desired high temperature levels in the recuperator.
It is an object of this invention to provide a gas turbine engine with improved fuel consumption over various power conditions .
According to the present invention there is provided a gas turbine engine comprising a centrifugal compressor, a diffuser, a heat exchanger, combustion apparatus, and at least one turbine, wherein said centrifugal compressor, said diffuser and at least one turbine comprise means for varying the flow capacity at their inlets, such that in operation the flow capacity of each component is independently variable so that over a predetermined power range the gas turbine engine mass flow is variable whilst maintaining the temperature, pressure ratio and speed of rotation of the gas turbine engine substantially constant.
Preferably a first turbine is drivingly connected to the centrifugal compressor and a second turbine is drivingly connected to a load. The second turbine may be connected to the load via a gear unit .
Alternatively a first turbine is drivingly connected to the centrifugal compressor and is also drivingly connected to an electrical generator. The electrical generator may be electrically connected to at least one electrical motor. The electrical motor may be drivingly connected to a load.
The load may comprise a propeller of a marine vessel or a driving wheel of an automotive vehicle. Preferably the means for varying the flow capacity of the compressor comprises variable inlet guide vanes .
Preferably the means for varying the flow capacity of the diffuser comprises moveable diffuser vanes pivotable such that their leading edges move in a tangential direction with respect to the axis of the diffuser.
Preferably the variable diffuser vanes are adapted to be moveable in unison with the variable inlet guide vanes .
Preferably the means for varying the flow capacity of the at least one turbine comprises variable area nozzles positioned within the inlet to the at least one turbine.
Preferably the at least one turbine is arranged to have a choked operation over the predetermined power range to ensure the flow capacity remains proportional to the area of the variable area nozzles.
Preferably the combustion apparatus comprises a primary combustion stage, a secondary combustion stage and a tertiary combustion stage, each of the combustion stages is provided with premixed fuel and air. Preferably the tertiary stage includes opposite swirl to increase mixing and minimise swirl upstream of a dilution section.
Also according to the present invention there is provided a method of controlling a gas turbine engine wherein the engine includes a centrifugal compressor, diffuser means, a heat exchanger, combustion apparatus and at least one turbine, comprising the steps of independently varying the flow capacity of the centrifugal compressor, the diffuser, and the at least one turbine such that the mass flow through each component is proportional to the power requirements of the gas turbine engine .
Also according to the present invention there is provided a method of controlling a gas turbine engine wherein the engine includes a centrifugal compressor, a diffuser, a heat exchanger, combustion apparatus and at least one turbine, said compressor, diffuser and at least one turbine all comprise means for varying the flow capacity at their inlets, comprising the steps of independently varying the flow capacity of each component so that over a predetermined power range the gas turbine engine mass flow is variable whilst maintaining the temperature, pressure ratio and speed of rotation of the gas turbine engine substantially constant. The invention will now be described more fully with reference to the accompanying drawings in which:
Figure 1 is a schematic view of a gas turbine engine according to the present invention,
Figure 2 is a schematic view of a further gas turbine engine according to the present invention, and Figure 3 is a graph comparing specific fuel consumption against power for a diesel engine and a gas turbine engine according to the present invention.
The gas turbine engine proposed by the present invention is particularly suitable for marine and automotive applications which operate largely at low power. For example an automotive gas turbine engine may utilise 6% of the available power at speeds of approximately 30 mph, 18% of the available power at speeds of approximately 56 mph, 35% of the available power at speeds of approximately 75 mph for a vehicle with a maximum speed of 115 mph, as shown in figure 3.
The gas turbine engine arrangement shown in figure 1 is a preferred embodiment of the present invention. The gas turbine engine comprises a first centrifugal air compressor 2 comprising a radial inlet duct 4 incorporating variable inlet guide vanes 6 and a centrifugal impeller (not shown) . The centrifugal compressor 2 delivers air via a variable area radial diffuser 8 to a heat exchanger 10. The variable area radial diffuser 8 reduces the velocity of the air before it enters the heat exchanger 10.
The variable area inlet guide vanes 6 comprise an aerofoil cross section. These inlet guide vanes are fully open when full power is required thus allowing the air to reach the impeller without substantial swirl and even a small amount of anti-rotative swirl to ensure the maximum amount of flow is passed. When the power demand is reduced the variable area inlet guide vanes 6 are positioned in a semi- closed state. This causes rotative swirl of the working fluid reaching the impeller inlet, which reduces the relative velocity, because the rotative swirl velocity is effectively subtracted from the rotational speed vector. Thus the mass flow of the working fluid at any given speed is reduced. The centrifugal impeller produces an increase in static pressure and absolute velocity. The working fluid leaves the impeller at approximately Mach 1.
The working fluid passes into a variable area radial diffuser 8 which contributes to the compressor pressure rise by recovering velocity as static pressure. The diffuser vanes are pivoted so as to move in a tangential direction with respect to the diffuser and adjust the throat area. As these vanes are closed simultaneously with the variable inlet guide vanes, the leading edge incidence is optimised. In addition the degree of diffusion up to the throat is controlled. Too high an incidence or attempted diffusion would also result in surge. Surge is where the adverse flow conditions cause a high local pressure loss resulting in flow reversal as the pressure rise cannot be sustained. This working fluid is then passed through the heat exchanger 10, then into the combustion chamber 12 and then to a first turbine 14. The working fluid is preheated in the heat exchanger 10. Fuel is burned in the combustion chamber 12 and the resulting combustion products flow into the first turbine 14 which is drivingly connected to the centrifugal compressor 2.
The first turbine 14 incorporates a variable area nozzle 16 which is operated so as to close the vanes as power demand falls thus reducing flow capacity. The turbine design expansion ratio is chosen high enough to ensure choked operation over most of the power range, thus ensuring flow capacity remains proportional to nozzle throat area. If unchoked, the flow capacity would be set by the expansion ratio and rotor throat area, and sufficient variation could not be achieved via the nozzle area.
The exhaust gases from this first turbine 14 are then directed into a free power turbine 18. The power turbine 18 also comprise variable area nozzles. The power turbine 18 is connected to a power shaft 22 which is in turn connected through a gear unit 24 to an output shaft 26 coupled to any suitable load device (not shown) , for example, the driving wheels 28 of a motor vehicle or a propeller of a marine vessel. The hot exhaust gases from the power turbine 18 are directed back into the heat exchanger 10 to directly pre-heat the air from the diffuser 8 before it enters the combustion chamber 12.
The combustion chamber 12 is a staged combustor incorporating primary, secondary and tertiary stages of combustion arranged in flow series and a dilution section downstream of the tertiary stage of combustion. All stages of combustion are provided with premixed fuel and air, the downstream, tertiary, stage of combustion including opposite swirl to that of the intermediate, secondary, stage to increase mixing and minimise swirl upstream of the dilution section.
The engine provides a part load SFC curve which is comparable with a diesel engine, particularly in the idle to 25% power range as shown in figure 3. This is achieved by maintaining the cycle parameters (i.e. pressures and temperatures) substantially constant over a large portion of the power range by use of variable vanes . A further effect is to reduce the engine rotational speed variation required for a given power range, which reduces the acceleration time requirement because the angles of the variable vanes are adjustable at a faster rate than the rotational speed of the engine can change .
The gas turbine engine arrangement shown in figure 2 is a further embodiment of the present invention. The gas turbine engine comprises a first centrifugal air compressor 32 comprising a radial inlet duct 34 incorporating variable inlet guide vanes 36 and a centrifugal impeller (not shown) . The centrifugal compressor 2 delivers air via a variable area radial diffuser 8 to a heat exchanger 10. The variable area radial diffuser 8 reduces the velocity of the air before it enters the heat exchanger 10. The variable area inlet guide vanes 36 comprise an aerofoil cross section. These inlet guide vanes are fully open when full power is required thus allowing the air to reach the impeller without substantial swirl and even a small amount of anti-rotative swirl to ensure the maximum amount of flow is passed. When the power demand is reduced the variable area inlet guide vanes 36 are positioned in a semi- closed state. This causes rotative swirl of the working fluid reaching the impeller inlet, which reduces the relative velocity, because the rotative swirl velocity is effectively subtracted from the rotational speed vector. Thus the mass flow of the working fluid at any given speed is reduced. The centrifugal impeller produces an increase in static pressure and absolute velocity. The working fluid leaves the impeller at approximately ach 1.
The working fluid passes into a variable area radial diffuser 38 which contributes to the compressor pressure rise by recovering velocity as static pressure. The diffuser vanes are pivoted so as to move in a tangential direction with respect to the diffuser and adjust the throat area. As these vanes are closed simultaneously with the variable inlet guide vanes, the leading edge incidence is optimised. In addition the degree of diffusion up to the throat is controlled. Too high an incidence or attempted diffusion would also result in surge. Surge is where the adverse flow conditions cause a high local pressure loss resulting in flow reversal as the pressure rise cannot be sustained.
This working fluid is then passed through the heat exchanger 10, then into the combustion chamber 42 and then to a turbine 44. Fuel is burned in the combustion chamber 42 and the resulting combustion products flow into the turbine 44 which is drivingly connected to the centrifugal compressor 32.
The turbine 44 incorporates a variable area nozzle 46 which is operated so as to close the vanes as power demand falls thus reducing flow capacity. The turbine design expansion ratio is chosen high enough to ensure choked operation over most of the power range, thus ensuring flow capacity remains proportional to nozzle throat area. If unchoked, the flow capacity would be set by the expansion ratio and rotor throat area, sufficient variation could not be achieved via the nozzle area. The turbine 44 is also connected to an output shaft 48 which drives an electrical generator 50. The electrical generator 50 is arranged to supply electricity to one or more electrical motors 54 via electrical connections 52, for example, for driving the wheels 58 of a motor vehicle or a propeller of a marine vessel . The hot exhaust gases from the turbine 44 are directed back into the heat exchanger 40 to directly pre-heat the air from the diffuser 38 before it enters the combustion chamber 42.
The combustion chamber 42 is a staged combustor incorporating primary, secondary and tertiary stages of combustion arranged in flow series and a dilution section downstream of the tertiary stage of combustion. All stages of combustion are provided with premixed fuel and air, the downstream, tertiary, stage of combustion including opposite swirl to that of the intermediate, secondary, stage to increase mixing and minimise swirl upstream of the dilution section.
The engine provides a part load SFC curve which is comparable with a diesel engine, particularly in the idle to 25% power range as shown in figure 3. This is achieved by maintaining the cycle parameters (i.e. pressures and temperatures) substantially constant over a whole power range by use of variable vanes. A further effect is to reduce the engine rotational speed variation required for a given power range, which reduces the acceleration time requirement because the angles of the variable vanes are adjustable at a faster rate than the rotational speed of the engine can change .
The objective of the variable vanes at the inlet to the centrifugal compressor, diffuser, first turbine and power turbine is to allow reduced mass flow whilst maintaining pressure ratio and efficiency. The variable vanes are adjusted to reduce the mass flow for the compressor, diffuser and all the turbines as the power demand falls, or conversely the variable vanes are adjusted to increase mass flow for the compressor, diffuser and all the turbines over a predetermined wide power range. The values of the cycle temperature, speed and pressure ratio are maintained substantially constant over the predetermined wide power range and therefore the specific fuel consumption is maintained substantially constant over this predetermined wide power range. The variable area inlet guide vanes for the centrifugal compressor enhance the centrifugal compressor's ability to deliver reduced mass flow at constant engine rotational speed.

Claims

Claims : -
1. A gas turbine engine comprising a centrifugal compressor (2) , a diffuser (8) , a heat exchanger (10) , combustion apparatus (12), and at least one turbine (14,18), wherein said diffuser (8) and at least one turbine (14,18) comprise means (8,16,20) for varying the flow capacity at their inlets, characterised in that the centrifugal compressor (2) comprises means (6) for varying the flow capacity at its inlet such that in operation the flow capacity of each component (2,8,14,18) is independently variable so that over a predetermined power range the gas turbine engine mass flow is variable whilst maintaining the temperature, pressure ratio and speed of rotation of the gas turbine engine substantially constant .
2. A gas turbine engine according to claim 1 comprising a first turbine (14) drivingly connected to the centrifugal compressor (2) and a second turbine (18) drivingly connected to a load.
3. A gas turbine engine according to claim 2 wherein the second turbine (18) is connected to the load via a gear unit
(24) .
4. A gas turbine engine according to claim 1 comprising a first turbine (44) drivingly connected to the centrifugal compressor (32) and drivingly connected to an electrical generator (50) .
5. A gas turbine engine according to claim 4 wherein the electrical generator (50) is electrically connected (52) to at least one electrical motor (54) .
6. A gas turbine engine according to claim 5 wherein the electrical motor (54) is drivingly connected to a load (58) .
7. A gas turbine engine as claimed in any of claims 2, 3 or 6 wherein the load (28,58) comprises a propeller of a marine vessel or a driving wheel of an automotive vehicle.
8. A gas turbine engine according to any of claims 1 to 7 wherein the means (6) for varying the flow capacity of the compressor (2) comprises variable inlet guide vanes.
9. A gas turbine engine according to any of claims 1 to 8 wherein the means for varying the flow capacity of the diffuser (8) comprises moveable diffuser vanes pivotable such that their leading edges move in a tangential direction with respect to the axis of the diffuser.
10. A gas turbine engine according to claim 8 wherein the variable diffuser vanes are adapted to be moveable in unison with the variable inlet guide vanes (6) .
11. A gas turbine engine according to any of claims 1 to 10 wherein the means (16,20) for varying the flow capacity of the at least one turbine (14,18) comprises variable area nozzles positioned within the inlet to the at least one turbine (14, 18) .
12. A gas turbine engine according to claim 10 wherein the at least one turbine (14) is arranged to have a choked operation over the predetermined power range to ensure the flow capacity remains proportional to the area of the variable area nozzles (16) .
13. A gas turbine engine according to any of claims 1 to 12 wherein the combustion apparatus (12) comprises a primary combustion stage, a secondary combustion stage and a tertiary combustion stage, each of the combustion stages is provided with premixed fuel and air.
14. A gas turbine engine according to claim 13 wherein the tertiary stage includes opposite swirl to increase mixing and minimise swirl upstream of a dilution section.
15. A method of controlling a gas turbine engine wherein the engine includes a centrifugal compressor (2) , diffuser means (8) , a heat exchanger (10) , combustion apparatus (12) and at least one turbine (14,18), comprising the steps of independently varying the flow capacity of the centrifugal compressor (2), the diffuser(8), and the at least one turbine (14,18) such that the mass flow through each component (2,8,14,18) is proportional to the power requirements of the gas turbine engine .
16. A method of controlling a gas turbine engine wherein the engine includes a centrifugal compressor (2) , a diffuser (8) , a heat exchanger (10) , combustion apparatus (12) and at least one turbine (14,18), said compressor (2), diffuser (8) and at least one turbine (14,18) all comprise means (6,8,16,20) for varying the flow capacity at their inlets, comprising the steps of independently varying the flow capacity of each component (2,8,14,18) so that over a predetermined power range the gas turbine engine mass flow is variable whilst maintaining the temperature, pressure ratio and speed of rotation of the gas turbine engine substantially constant.
PCT/GB1999/000074 1998-01-15 1999-01-08 Gas turbine engine WO1999036688A1 (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
GB9800782.6 1998-01-15
GBGB9800782.6A GB9800782D0 (en) 1998-01-15 1998-01-15 Gas turbine engine

Publications (1)

Publication Number Publication Date
WO1999036688A1 true WO1999036688A1 (en) 1999-07-22

Family

ID=10825292

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/GB1999/000074 WO1999036688A1 (en) 1998-01-15 1999-01-08 Gas turbine engine

Country Status (2)

Country Link
GB (1) GB9800782D0 (en)
WO (1) WO1999036688A1 (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2003094265A1 (en) * 2002-05-03 2003-11-13 Rolls-Royce Plc A gas turbine engine and fuel cell stack combination
EP1362984A2 (en) * 2002-05-16 2003-11-19 ROLLS-ROYCE plc Gas turbine engine

Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB701505A (en) * 1949-09-24 1953-12-30 Centrax Power Units Ltd Improvements relating to centrifugal compressors
US2715814A (en) * 1949-03-25 1955-08-23 Centrax Power Units Ltd Fuel-flow for plural radial inwardflow gas turbines
US3025688A (en) 1959-06-11 1962-03-20 Draper Corp Pattern mechanism for knitting machines
US3138923A (en) 1956-03-24 1964-06-30 Volvo Ab Automotive gas turbine power power plant
US3625003A (en) * 1970-09-08 1971-12-07 Gen Motors Corp Split compressor gas turbine
US3981140A (en) * 1975-06-23 1976-09-21 General Motors Corporation Gas turbine engine geometry control
US4145875A (en) 1976-10-15 1979-03-27 General Motors Corporation Variable flow capacity gas turbine engine for improved part load fuel economy
US4195473A (en) * 1977-09-26 1980-04-01 General Motors Corporation Gas turbine engine with stepped inlet compressor
US4414805A (en) * 1981-11-27 1983-11-15 General Motors Corporation Hybrid gas turbine engine and flywheel propulsion system
GB2172340A (en) * 1985-03-08 1986-09-17 Hitachi Shipbuilding Eng Co Turbocharger for diesel engine and method of controlling same
US4821506A (en) * 1987-10-08 1989-04-18 Sundstrand Corporation Radial turbine with variable axial nozzle
GB2210108A (en) 1987-09-26 1989-06-01 Rolls Royce Plc A variable guide vane arrangement for a compressor
EP0515746A1 (en) * 1991-05-22 1992-12-02 Mitsubishi Jukogyo Kabushiki Kaisha Gas turbine and operating method of the same
GB2301868A (en) 1995-06-05 1996-12-18 Rolls Royce Plc Actuator mechanism with emergency drive for variable angle vane arrays

Patent Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2715814A (en) * 1949-03-25 1955-08-23 Centrax Power Units Ltd Fuel-flow for plural radial inwardflow gas turbines
GB701505A (en) * 1949-09-24 1953-12-30 Centrax Power Units Ltd Improvements relating to centrifugal compressors
US3138923A (en) 1956-03-24 1964-06-30 Volvo Ab Automotive gas turbine power power plant
US3025688A (en) 1959-06-11 1962-03-20 Draper Corp Pattern mechanism for knitting machines
US3625003A (en) * 1970-09-08 1971-12-07 Gen Motors Corp Split compressor gas turbine
US3981140A (en) * 1975-06-23 1976-09-21 General Motors Corporation Gas turbine engine geometry control
US4145875A (en) 1976-10-15 1979-03-27 General Motors Corporation Variable flow capacity gas turbine engine for improved part load fuel economy
US4195473A (en) * 1977-09-26 1980-04-01 General Motors Corporation Gas turbine engine with stepped inlet compressor
US4414805A (en) * 1981-11-27 1983-11-15 General Motors Corporation Hybrid gas turbine engine and flywheel propulsion system
GB2172340A (en) * 1985-03-08 1986-09-17 Hitachi Shipbuilding Eng Co Turbocharger for diesel engine and method of controlling same
GB2210108A (en) 1987-09-26 1989-06-01 Rolls Royce Plc A variable guide vane arrangement for a compressor
US4821506A (en) * 1987-10-08 1989-04-18 Sundstrand Corporation Radial turbine with variable axial nozzle
EP0515746A1 (en) * 1991-05-22 1992-12-02 Mitsubishi Jukogyo Kabushiki Kaisha Gas turbine and operating method of the same
GB2301868A (en) 1995-06-05 1996-12-18 Rolls Royce Plc Actuator mechanism with emergency drive for variable angle vane arrays

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2003094265A1 (en) * 2002-05-03 2003-11-13 Rolls-Royce Plc A gas turbine engine and fuel cell stack combination
EP1362984A2 (en) * 2002-05-16 2003-11-19 ROLLS-ROYCE plc Gas turbine engine
EP1362984A3 (en) * 2002-05-16 2006-09-06 ROLLS-ROYCE plc Gas turbine engine

Also Published As

Publication number Publication date
GB9800782D0 (en) 1998-03-11

Similar Documents

Publication Publication Date Title
EP1055809B1 (en) A gas turbine engine and a method of controlling a gas turbine engine
US5911679A (en) Variable pitch rotor assembly for a gas turbine engine inlet
US5768884A (en) Gas turbine engine having flat rated horsepower
US6003298A (en) Steam driven variable speed booster compressor for gas turbine
CA1229493A (en) Fluid injection gas turbine engine and method for operating
EP1362984B1 (en) Gas turbine engine
EP1055879B1 (en) A combustion chamber assembly and a method of operating a combustion chamber assembly
EP1252424B1 (en) Method of operating a variable cycle gas turbine engine
US5155993A (en) Apparatus for compressor air extraction
US4660376A (en) Method for operating a fluid injection gas turbine engine
US3514952A (en) Variable bypass turbofan engine
JP5325367B2 (en) Method and apparatus for operating a gas turbine engine
US5119624A (en) Gas turbine engine power unit
US4641495A (en) Dual entry radial turbine gas generator
RU2575837C9 (en) Apparatus and method for reducing air mass flow for extended range low emissions combustion for single shaft gas turbines
US5160080A (en) Gas turbine engine and method of operation for providing increased output shaft horsepower
US4428714A (en) Pre-swirl inlet guide vanes for compressor
EP1967717A1 (en) Gas turbine with a bypass conduit system
EP1368560A1 (en) Turbine engine
RU2140001C1 (en) Method of operation of supersonic hybrid air-jet engine plant
US20060248899A1 (en) Method for producing gas turbines and gas turbine assembly
US4640091A (en) Apparatus for improving acceleration in a multi-shaft gas turbine engine
JP2954754B2 (en) Operation control device for gas turbine system and pressurized fluidized bed boiler power plant
GB2377973A (en) Gas bleed system for a gas turbine
US4170874A (en) Gas turbine unit

Legal Events

Date Code Title Description
AK Designated states

Kind code of ref document: A1

Designated state(s): US

AL Designated countries for regional patents

Kind code of ref document: A1

Designated state(s): AT BE CH CY DE DK ES FI FR GB GR IE IT LU MC NL PT SE

DFPE Request for preliminary examination filed prior to expiration of 19th month from priority date (pct application filed before 20040101)
121 Ep: the epo has been informed by wipo that ep was designated in this application