GB2064006A - Dual Fuel System for a Gas Turbine Engine - Google Patents

Dual Fuel System for a Gas Turbine Engine Download PDF

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Publication number
GB2064006A
GB2064006A GB7940687A GB7940687A GB2064006A GB 2064006 A GB2064006 A GB 2064006A GB 7940687 A GB7940687 A GB 7940687A GB 7940687 A GB7940687 A GB 7940687A GB 2064006 A GB2064006 A GB 2064006A
Authority
GB
United Kingdom
Prior art keywords
fuel
combustion
liquid fuel
vapourising
liquid
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
GB7940687A
Other versions
GB2064006B (en
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
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
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Priority to GB7940687A priority Critical patent/GB2064006B/en
Priority to DE19803043698 priority patent/DE3043698A1/en
Priority to FR8024809A priority patent/FR2470249A1/en
Priority to JP16520380A priority patent/JPS5696121A/en
Publication of GB2064006A publication Critical patent/GB2064006A/en
Application granted granted Critical
Publication of GB2064006B publication Critical patent/GB2064006B/en
Expired legal-status Critical Current

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Classifications

    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/36—Supply of different fuels
    • 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
    • F02C3/20—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products
    • F02C3/26—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being solid or pulverulent, e.g. in slurry or suspension
    • F02C3/28—Gas-turbine plants characterised by the use of combustion products as the working fluid using a special fuel, oxidant, or dilution fluid to generate the combustion products the fuel or oxidant being solid or pulverulent, e.g. in slurry or suspension using a separate gas producer for gasifying the fuel before combustion
    • 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
    • F02C7/22—Fuel supply systems
    • 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
    • Y02T10/00—Road transport of goods or passengers
    • Y02T10/10—Internal combustion engine [ICE] based vehicles
    • Y02T10/30—Use of alternative fuels, e.g. biofuels

Abstract

The system (10) comprises a gaseous fuel duct (30) feeding a manifold (32) and gaseous fuel burners (24) located in the engine combustion chamber or chambers (20). Liquid fuel can be burnt by first burning a small proportion of the liquid fuel in a vapourising combustion chamber (38), injecting the remaining liquid fuel into the vapouriser and ducting the vapourised liquid fuel into the gaseous fuel duct (30). The engine combustion chambers (20) are preferably of the type in which the fuel is burnt rich in the primary zone and is quenched almost immediately by dilution air so as to reduce the NOx emissions. <IMAGE>

Description

SPECIFICATION Improvements in or Relating to Gas Turbine Engine Fuel Supply Systems This invention relates to the fuel supply systems of gas turbine engines, more particularly to those gas turbine engines which are arranged to operate on liquid and gaseous fuels.
In order to meet the emission requirements as determined by the United States Environmental Protection Agency (EPA), considerable effort has been expended by the gas turbine industry to reduce inter alia, the emissions of nitrogen oxides, otherwise known as NOx. The present invention is particularly concerned with reducing NOx emissions in engines which are designed to operate on liquid fuels such as kerosene and diesel and gaseous fuels such as gaseous propane and methane. A known method of reducing Nox is to burn in the engine combustion chamber primary zone a rich fuel and air mixture, e.g. not less than 1.3 equivalence ratio, and then introduce all the dilution air immediately downstream of the primary zone and mix it rapidly as possible with the combustion gases.This method is known as "rich burn rapid quench", and tests have shown that NOx can be reduced substantially with gaseous propane and methane, at least to a degree sufficient to meet the EPA requirements.
In the case of liquid fuels the application of the rich burn rapid quench method of combustion has not been so advantageous, even though from a combustion and NOx production point of view there is little difference between the liquid diesel fuel and the gaseous propane fuel.
The present invention seeks to provide a method and means of preparing the liquid fuel in a gas turbine engine dual fuel supply system so that the NOx emissions produced by the combustion of liquid fuels is comparable to the NOx emissions produced by the combustion of gaseous fuels.
In one aspect of the invention, there is provided a method of burning fuel in a gas turbine engine in which the fuel supply comprises a liquid fuel, a gaseous fuel or a mixture of liquid and gaseous fuels, the method comprising supplying the fuel to the engine combustion means which has only gaseous fuel burner means, the liquid fuel being vapourised externally of the engine combustion means, the heat for said vapourisation being provided by burning a proportion of the liquid fuel, the vapourised liquid fuel being supplied to the engine combustion means.
In another aspect of the invention there is provided a dual fuel system for a gas turbine engine comprising liquid fuel supply means and gaseous fuel supply means, the gaseous fuel supply means including gaseous fuel burners located in the engine combustion means, the liquid fuel supply means comprising vapourising combustion means located externally of the engine combustion means and in fluid communication with the gaseous fuel supply means, the vapourising combustion means being arranged to receive a proportion of the air delivered by the engine compressor, having a fuel burner means arranged to receive a proportion of the liquid fuel, the remainder of the liquid fuel being injected into the vapourising combustion means, the vapourised liquid fuel being arranged to pass into the gaseous fuel supply means.
The gaseous fuel supply means may comprise a gaseous fuel supply duct, a gaseous fuel manifold and the gaseous fuel burners located in the combustion chamber or chambers of the engine.
The liquid fuel vapourising means may comprise a combustion chamber with a liquid fuel burner having pilot and main fuel nozzles which operate on a proportion of the liquid fuel flow, the remainder of the liquid fuel flow being injected into the vapourising combustion chamber downstream of the liquid fuel burner.
The vapourising combustion chamber may receive compressed air from the engine via a bleed from the engine compressor and a boost compressor can be used to increase the pressure of the engine compressor delivery. Both the liquid and gaseous fuels may initially flow through a common fuel proportioning valve which controls the flow of each fuel according to whether liquid fuel, gaseous fuel or a mixture of liquid and gaseous fuel is required. From the fuel proportioning valve the gaseous fuel flows direct to the gaseous fuel burners and the liquid fuel flows to the gaseous fuel burners via the liquid fuel vapourising combustion means.
Airflow control means may be provided upstream of the vapourising combustion chamber to control the airflow to the vapourising combustion chamber according to the temperature of the vapourised liquid fuel leaving the vapourising combustion chamber.
It will be noted that an added benefit of the invention is that downstream of the vapourising combustion chamber, the fuel supply system is common for both initially liquid fuels, and gaseous fuels, there being no need for separate fuel lines, manifolds and specially designed dual fuel burners.
The present invention will now be more particularly described with reference to the accompanying drawings in which, Figure 1 is a schematic layout of a gas turbine engine having a dual fuel supply system according to the present invention, Figure 2 shows the liquid fuel vapourising combustion chamber of Figure 1, Figure 3 shows the liquid fuel burner of the combustion chamber shown in Figure 2 and, Figure 4 shows one of the combustion chambers of the gas turbine engine shown in Figure 1.
Referring to the Figures, a gas turbine engine 10 comprises a compressor 12 driven by a turbine 14 via a shaft 1 6. Combustion equipment 1 8 comprising a series of cylindrical combustion chambers 20 in an annular housing 22 each having a fuel burner 24 receives compressed air from the compressor 12 and fuel from a fuel supply system 26, the products of combustion flowing from the combustion chambers to drive the turbine 14.
The fuel supply system 26 comprises a gas and liquid fuel proportioning valve 28 arranged to receive gas and liquid fuels from their respective stores (not shown), the gaseous fuel then flowing along a gas duct 30 and a gas manifold 32 around the annular housing 22 and thence to the burners 24, which are a conventional type of gas burner and are illustrated in Figure 4.
From the valve 28, the liquid fuel flows to a liquid fuel proportioning valve 34 which divides the liquid fuel into two portions, one relatively small portion flowing to a fuel burner 36 of a liquid fuel vapourising combustion chamber 36, whilst the remaining fuel flows into the combustion chamber 38 downstream of the burner 36. The flow of liquid fuel to the burner 36 is controlled by a valve unit 40 which contains a fuel shut-off valve and a pressurising valve (both not shown). The valve unit 40 has two outlets 42, 44, outlet 42 for the pilot fuel to the burner 40 shown in more detail in Figure 3 and outlet 44 for the main fuel, the pressurising valve opening when the fuel is at sufficient pressure to allow the main fuel supply to flow.
The air supply to the combustion chamber 38 comprises a bleed from the engine compressor 12 and a boost compressor 46 driven from an external source or off the engine 10, can be provided in the air supply to ensure the correct air pressure at the inlet to the combustion chamber 38.
The mass flow of the air supply can be regulated by a valve 48 which is controlled by an air valve controller 50, the operation of the valve being dependent on the temperature of the gases downstream of the combustion chamber as detected by a temperature sensor 52. Thus if the temperature is too low, the controller 50 operates to open up the valve 46 and increases the air flow to the combustion chamber and vice versa if the temperature is too high.
Referring to Figure 2, the liquid fuel vapourising combustion chamber 38 has the fuel burner 36 located at its upstream end in a ring of swirler vanes 54 which are positioned in a circular section passage 56. Air from the compressor 12 flows into the combustion chamber 38 through the passage 36 and through inlets 58 in the combustion chamber wall via an annular space 60 and openings (not shown) in an outer casing 62.
As mentioned previously a small proportion of the liquid fuel flows through the burner 36 in the form of pilot and main fuel, whilst the remaining liquid fuel is injected into the chamber downstream of the burner. The remaining liquid fuel flows into a manifold 64 and enters the chamber through a number of injection holes 66.
Figure 3 shows in more detail the construction of the liquid fuel burner 36. The pilot fuel flows from the duct 42 to a pilot orifice 68 via a duct 70, filter 72, manifold 74 and outlets 76, whilst the main fuel flows from duct 44 to an annular main orifice 78 via a duct 80, manifold 82, drillings 84, a second manifold 86 and outlets 88.
A shroud 90 surrounds the end of the burner to guide a flow of air through apertures 92 over the main orifice to prevent the formation of carbon.
Figure 4 shows one of the combustion chambers 20 in more detail. The chamber 20 is of the "rich burn rapid quench" type which has been found to be effective in reducing the emissions of NOx in the combustion of gaseous fuels. The burner 24 at the upstream end of the chamber is of a standard type and is known as a "pepperpot" burner, e.g. it has nozzles 24a formed in the otherwise closed end of the burner. The primary air enters through a ring of swirl vanes 94 and the secondary air to complete the combustion enters through a number of secondary ports 96. The combustion gases are almost immediately quenched by air entering through rapid quench ports 98 which are located relatively close to the ports 96.The fuel and air mixture in the primary zone is fuel rich, having an equivalence ratio of not less than 1.3 and the rapid quenching of the combustion gases almost immediately downstream of the primary zone has been found to reduce the NOx emissions by up to 50% for one particular type of gas turbine engine when burning gaseous propane and methane.
Referring back to Figure 1, the operation of the fuel supply system 26 is as follows:- When the engine 10 is to be run on gaseous fuel only, the proportioning valve 28 allows gas fuel only to flow along the duct 30 to the burners 24 in the combustion chambers 20.
When the engine is to be run on liquid fuel, the proportioning valve 28 closes off the gaseous fuel flow to the duct 30 and allows liquid fuel to flow to the valve 34. Approximately 3% of the total liquid fuel flow passes to the burner 36, whilst the remainder is injected into the vapourising combustion chamber 38 through holes 66.
Approximately 1% of the compressor delivery air is bled from the compressor outlet and passed through the boost compressor 46 which in this application has a compression ratio of 1.1, to the chamber 38. The heat released by the burning of the 3% liquid fuel is sufficient to vapourise the remainder of the liquid fuel and the vapourised fuel passes into the duct 30 and thence to the burners 24. Thus the liquid fuel enters the "rich burn quick quench" combustion chambers 20 effectively as a gaseous fuel with the consequent advantages of an initially gaseous fuel in respect of NOx emissions. It will also be appreciated that the combustion system according to the invention also has the advantages of a common fuel supply system and simple fuel burners as opposed to the relatively complicated known dual fuel burners.
Since the fuel to air ratio of the engine 10 varies over the operating range and during transient conditions, it is necessary to control the flow of air to the vapourising combustion chamber to control the temperature of the delivery flow from the chamber 38. The temperature sensor 52 is located at a point well downstream of the chamber 38 where the reaction and mixing is substantially complete.
The engine 10 can be run on gaseous or liquid fuels alone or on a mixture of these fuels. For practical reasons, if the liquid fuel proportion falls below 25% the boost compressor 46 and valve 48 are closed off, and the valve 28 allows only gaseous fuel to flow.
Although the invention has been specifically described in relation to a 'can-annular' combustion system, it can also be applied to other gas turbine engine combustion systems, e.g. a single annular combustion chamber.

Claims (10)

Claims
1. A gas turbine engine dual fuel system comprising liquid fuel supply means and gaseous fuel supply means, the gaseous fuel supply means including gaseous fuel burners located in the engine combustion means, the liquid fuel supply means comprising vapourising combustion means located externally of the engine combustion means and in fluid communication with the gaseous fuel supply means, the vapourising combustion means being arranged to receive a proportion of the compressed air delivered by the gas turbine engine compressor and having fuel burner means arranged to receive a proportion of the liquid fuel, the remainder of the liquid fuel being injected into the vapourising combustion means.
2. A fuel system as claimed in claim 1 in which the liquid fuel vapourising means comprises a combustion chamber having a liquid fuel burner receiving a proportion of the liquid fuel, the remainder of the liquid fuel being injected into the vapourising combustion chamber downstream of the burner.
3. A fuel system as claimed in claim 2 in which the vapourising combustion chamber receives a proportion of the compressed air from the gas turbine engine compressor.
4. A fuel system as claimed in claim 3 in which the said proportion of compressed air passes through a boost compressor upstream of the vapourising combustion chamber.
5. A fuel system as claimed in any one of the preceding claims having fuel proportioning means arranged to control the supply of liquid and gaseous fuels to the vapourising combustion means and the gaseous fuel burners respectively.
6. A fuel system as claimed in claim 5 having a liquid fuel proportioning valve downstream of the fuel proportioning valve and upstream of the vapouring combustion means.
7. A fuel system as claimed in any one of the preceding claims having airflow control means comprising valve means to control the proportion of engine compressor air delivered to the vapourising combustion means, control means to operate the said valve means, and temperature sensing means located downstream of the vapouring combustion means, the function of the control means being dependent on the temperature sensed by the temperature sensing means.
8. A fuel system as claimed in any one of the preceding claims in which the gas turbine engine combustion means is arranged to burn fuel in the primary zone at an equivalence ratio of not less 1.3 and the dilution air enters the combustion means immediately downstream of the primary zone.
9. A method of burning fuel in a gas turbine engine in which the fuel supply comprises a liquid fuel, a gaseous fuel or a mixture of liquid and gaseous fuels, the method comprising supplying the fuel to the engine combustion means which has only gaseous fuel burner means, the liquid fuel being vapourised externally of the engine combustion means, the heat for said vapourisation being provided by burning a proportion of the liquid fuel, the vapourised liquid fuel being supplied to the engine combustion means.
10. A gas turbine engine fuel system constructed and arranged for use and operation substantially as herein described with reference to and as shown in the accompanying drawings.
GB7940687A 1979-11-24 1979-11-24 Dual fuel system for a gas turbine engine Expired GB2064006B (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
GB7940687A GB2064006B (en) 1979-11-24 1979-11-24 Dual fuel system for a gas turbine engine
DE19803043698 DE3043698A1 (en) 1979-11-24 1980-11-19 DOUBLE FUEL SYSTEM FOR GAS TURBINE ENGINES
FR8024809A FR2470249A1 (en) 1979-11-24 1980-11-21 FUEL CIRCUIT OF A GAS TURBINE ENGINE FOR LIQUID OR GAS FUEL
JP16520380A JPS5696121A (en) 1979-11-24 1980-11-22 Composite fuel system for gas turbine engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB7940687A GB2064006B (en) 1979-11-24 1979-11-24 Dual fuel system for a gas turbine engine

Publications (2)

Publication Number Publication Date
GB2064006A true GB2064006A (en) 1981-06-10
GB2064006B GB2064006B (en) 1983-09-14

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
GB7940687A Expired GB2064006B (en) 1979-11-24 1979-11-24 Dual fuel system for a gas turbine engine

Country Status (4)

Country Link
JP (1) JPS5696121A (en)
DE (1) DE3043698A1 (en)
FR (1) FR2470249A1 (en)
GB (1) GB2064006B (en)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6247299B1 (en) * 1996-11-16 2001-06-19 Abb Research Ltd. Method for feeding a gas turbine with both liquid and gaseous fuels
US9097208B2 (en) 2012-12-14 2015-08-04 Electro-Motive Diesel, Inc. Cryogenic pump system for converting fuel
CN105556188A (en) * 2013-09-20 2016-05-04 西门子股份公司 Pipe connection arrangement, high-pressure fluid line system of a double internal combustion engine, double internal combustion engine and use of a tension nut

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19539246A1 (en) * 1995-10-21 1997-04-24 Asea Brown Boveri Airblast atomizer nozzle

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FR1143041A (en) * 1954-09-10 1957-09-25 Henschel & Sohn Gmbh Gas turbine and pressure carburetor with different pressure level
US2958189A (en) * 1955-05-31 1960-11-01 Phillips Petroleum Co Method and apparatus for providing improved combustion in jet engines
CH343714A (en) * 1956-09-10 1959-12-31 Bbc Brown Boveri & Cie Control device for the burner of a gas turbine combustion chamber for the simultaneous combustion of a gaseous and a liquid fuel
DE1944307A1 (en) * 1969-09-01 1971-03-11 Metallgesellschaft Ag Turbine power plant process
US4161164A (en) * 1972-07-17 1979-07-17 Siemens Aktiengesellschaft Internal combustion engine fuel supply system
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DE2260515A1 (en) * 1972-12-11 1974-06-12 Siemens Ag GAS TURBINE WITH EVEN, COMPLETE COMBUSTION OF THE FUEL SUPPLIED TO IT
US3866411A (en) * 1973-12-27 1975-02-18 Texaco Inc Gas turbine process utilizing purified fuel and recirculated flue gases
GB1470867A (en) * 1973-12-27 1977-04-21 Texaco Development Corp Gas turbine process utilizing purified fuel and recirculated fuel gas
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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6247299B1 (en) * 1996-11-16 2001-06-19 Abb Research Ltd. Method for feeding a gas turbine with both liquid and gaseous fuels
US9097208B2 (en) 2012-12-14 2015-08-04 Electro-Motive Diesel, Inc. Cryogenic pump system for converting fuel
CN105556188A (en) * 2013-09-20 2016-05-04 西门子股份公司 Pipe connection arrangement, high-pressure fluid line system of a double internal combustion engine, double internal combustion engine and use of a tension nut
US10281068B2 (en) 2013-09-20 2019-05-07 Siemens Aktiengesellschaft Pipe connection arrangement, high-pressure fluid line system of a dual fuel engine, dual fuel engine and use of a tension nut
CN105556188B (en) * 2013-09-20 2019-08-02 西门子股份公司 The use of the assembling pipe joint, high-pressure fluid line system, double internal combustion engines and tensioning nut of double internal combustion engines

Also Published As

Publication number Publication date
GB2064006B (en) 1983-09-14
FR2470249A1 (en) 1981-05-29
JPS5696121A (en) 1981-08-04
DE3043698A1 (en) 1981-06-11

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