GB2034874A - Gas turbine engine combustor - Google Patents
Gas turbine engine combustor Download PDFInfo
- Publication number
- GB2034874A GB2034874A GB7933578A GB7933578A GB2034874A GB 2034874 A GB2034874 A GB 2034874A GB 7933578 A GB7933578 A GB 7933578A GB 7933578 A GB7933578 A GB 7933578A GB 2034874 A GB2034874 A GB 2034874A
- Authority
- GB
- United Kingdom
- Prior art keywords
- cooling
- zone
- air
- chamber
- primary
- 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- Withdrawn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
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
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
-
- 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
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
- F23R3/08—Arrangement of apertures along the flame tube between annular flame tube sections, e.g. flame tubes with telescopic sections
-
- 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)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A gas turbine air cooled combustor 40 comprises a combustor liner means 46, 48 define a combustion chamber. A first portion 62, 64, of the liner means defines a primary combustion zone 70 while a second portion 76, 86, of the liner means defines a secondary dilution zone 72 downstream of the primary zone. Air, associated with cooling of the first liner portion defining the primary zone, is prevented from being discharged into the primary zone thus reducing the formation of gaseous emissions. <IMAGE>
Description
SPECIFICATION
Gas turbine engine combustor
This invention relates to combustion apparatus and, more particularly, to means for reducing the gaseous emissions from the combustor apparatus associated with a gas turbine engine.
Aircraft engines presently in operational use and those under development for future applications are designed to operate at extremely high temperatures. Combustors associated with high pressure ratio gas turbine engines not only must be compatible with the high temperature environment but also must perform satisfactorily
for long periods of time before removal from the engine for repair and maintenance. Present day materials cannot operate for extended periods of time at temperatures typically found in the combustor section of high pressure ratio gas turbine engines. Efficient and reliable means for cooling the combustor components must be provided.
State of the art cooling techniques have included the provision for a moving film of cooling air between the inner surface of the liners of the combustor and the hot gases,burning inside the combustion chamber. The film of cooling air prevents the hot gases from contacting the combustor liner. Generally, the protective film is introduced into the combustion chamber from a plenum surrounding the combustor.
The technique of film cooling of combustor liners has generally met with success in the aircraft engine industry. As a result of this technique combustor life has been significantly extended thereby reducing combustor replacement frequency and time between combustor overhaul.
In order to effect improvements in the quality of the environment, various federal agencies have issued regulations controlling the amounts and concentrations of pollutants which may be discharged to the atmosphere. Gaseous emissions of aircraft gas turbine engines have been addressed by the agencies and have resulted in very stringent restrictions especially those relating to emissions of nitrogen oxides, hydrocarbons and carbon monoxide. In some instances, these restrictions cannot be met without significant advances in the state of the art of combustion technology.
The utilization of film cooling in aircraft gas turbine engine combustors has been found to contribute significantly to the gaseous emissions of the engine particularly to those emissions of carbon monoxide when the engine is operating at ground idle conditions. More specifically, under such conditions the film of cooling air acts to upset the delicate balance of the fuel to air ratio present in the primary combustion zone of the com6ustor. Typically the amount of air introduced for film cooling of the combustor at the primary combustion zone can be approximately 10--15 per cent of the total air flow throug the combustor (this percentage also constitutes 20-40 percent of the primary zone air) resulting in the leaning of the mixture (especially at the liner walls) and a resultant increase in the emission of carbon monoxide, and unburned hydrocarbons.
Additionally, since at idle conditions the temperature of the film of cooling air is relatively cool, approximately 4000 F, while the temperature of the flame in the primary zone approximates 3000"F, local quenching of the flame occurs at its interface with the film of cooling air. Quenching of the flame lowers the temperature at which burning at the interface occurs such that burning is incomplete and eletated levels of carbon monoxide and unburned hydrocarbons emissions are encountered. The present invention addresses these problems associated with the incompatibility of the application of film cooling techniques with restriction on the gaseous emissions of a gas turbine engine.
It is therefore an object of the present invention to provide combustion apparatus wherein cooling air is used to cool the combustor liner without the production of significant gaseous emissions.
It is another object of the present invention to provide combustion apparatus wherein cooling air used to cool the combustor is prevented from quenching the burning fuel and air mixture in the primary combustion zone.
Briefly stated, this and other objects, which will become apparent hereinafter, are accomplished by the present invention which provides, in one form, combustor liner means defining a primary combustion zone and secondary dilution zone within a combustion chamber. First cooling means are provided for cooling a first portion of the liner means defining the primary combustion zone.
Means are provided for preventing discharge of the cooling air into the primary combustion zone.
Second cooling means are provided for cooling a second portion of the liner means defining the secondary dilution zone.
While the specification concludes with claims distinctly claiming and particularly pointing out the invention described herein, it will be more readily understood by reference to the description to follow together with the appended drawings in which:
Fig. 1 is a cross-sectional representation of annular gas turbine engine combustor embodying the present invention.
Referring to Fig. 1 an annular combustor, shown generally at 40, is comprised of axially and circu mferentia Ily extending radially outer wall 42 spaced radially apart from axially and circumferentially extending radially inner wall 44.
Liner means, in the form of first and second radially outer and inner combustor liners 46 and 48, respectively, are disposed within the annular spacing between walls 42 and 44. Liners 46 and 48 extend circumferentially and axially and are radially spaced from each other to define a longitudinally extending generally annular combustion chamber 50. Outer liner 46 cooperates in spaced-apart relationship with outer wall 42 to form a first radially outward annular longitudinally extending plenum 52 therebetween while inner liner 48 is disposed in spaced-apart relationship with inner wall 44 to define therebetween a second radially inward annular and longitudinally extending plenum 54. Plenum 52 and 54 are in communication at their upstream ends with a source of pressurized air such as the compressor (not shown) associated with the gas turbine engine.As stated above the general configuration of combustor 40 is of the generally well-known type having an annular shape disposed symmetrically about center line x-x. For purposes of the description hereinafter to follow the terms "axially" or "longitudinally" shall mean a direction generally along the line x-x, the term "radial" shall mean a direction radial to the line x-x and the term "circumferential" shall mean a direction circumferentially about the line x-x.
A plurality of circumferentially spaced apart fuel nozzle and swirler assemblies 56 (only one of which is shown) are disposed at the upstream end of combustion chamber 50 and are mounted upon a pair of circumferentially extending flanges 58 and 60. Flanges 58 and 60 depend radially toward each other from liners 46 and 48 respectively at the upstream end of combustion chamber 50. Fuel nozzle and swirler assemblies 56 serve to
introduce a mixture of compressed air and fuel into the upstream end of combustion chamber50 at a relatively high velocity. Fuel and air introduced in this manner is burned in chamber 40 and the hot gas resulting from combustion flows afterward and is expelled through exit 57 into the turbine section of a gas turbine engine.
Outer liner 46 is comprised of a generally longitudinally and circumferentially extending continuous upstream portion or segment 62 disposed immediately proximate fuel nozzle and swirler assemblies 56. Similarly, inner liner 48 includes a generally longitudinally and circumferentially extending continuous upstream portion or segment 64 disposed immediately proximate fuel nozzle and swirler assemblies 56.
It is observed that continuous, uninterrupted liner portions or segments 62 and 64 define a first or upstream portion of combustor liner means 46 and 48 and are disposed radially opposite to one another at the same general axial location.
Combustion chamber 50 includes an annular primary combustion zone 70 bounded generally on its upstream side by the discharge or exit plane 66 of fuel nozzle and swirler assembly 56; at its
radially outward and inner extremities by segments 62 and 64 respectively. Primary combustion zone 70 is a zone wherein burning of the
air and fuel mixture is substantially completed. To
accomplish substantially complete burning, the
flowing gas must reside within zone 70 for a
period of time sufficient to permit the required
degree of conversion of hydrocarbons and carbon
monoxide to carbon dioxide necessary to meet
low emission requirements. Consequently, zone
70 extends downstream from exit plane 66 of
assembly 56 and terminates at a distance d which
has a preselected magnitude sufficient to permit the aforementioned conversion.Segments 62 and 64 extend downstream from exit plane 66 and include lip portions 68 and 69, respectively, which terminate at approximately the same distance from plane 66 as does primary zone 70. The
magnitude of distance d may be preselected for various combustors in accordance with known
procedures and for most, if not all, combustors will
be approximately equal to radial spacing h
between lines 46 and 48 proximate plane 66.
Liners 46 and 48 further define and
circumscribe a secondary or dilution zone 72
within combustion chamber 50 downstream of
zone 70. Secondary zone 72 comprises a zone
wherein the hot gases, resulting from burning of
the fuel and air in zone 70, are mixed with dilution
air to improve the radial and circumferential
temperature distrIbution of the gases as they exit
the combustor. Accordingly, liner means 46
further comprises a circumferentially and
longitudinally extending stepped wall 74 disposed
immediately downstream of segments 62 and
comprised of a plurality of circumferentially and
axially extending segments 76 disposed
consecutively one after the other in the axial
direction. Each segment 76 includes downstream
lip portion 78 and upstream step portion 80.Step portion 80 of each segment 76 overlaps downstream lip portion 78 of the next adjacent upstream segment 76. The step portion 80 of the
segment 76 adjacent segment 62 overlaps
cooling lip 68 or segment 62. Each step 80
includes cooling means in the form of a plurality of
circumferentially spaced-apart apertures 82 to
provide for the passage of cooling air from plenum
52 into the secondary dilution zone 72 of chamber
50 forfilm cooling of the downstream segments
76 of liner 46 in the convention manner well
known in the art.
Secondary dilution zone 72 is further defined by circumferentially and longitudinally extending stepped wall 84 of liner 48. Stepped wall 84 is disposed immediately downstream of segment 64 and is comprised of a plurality of circumferentially and axially extending segments 86 disposed one after the other in the axial direction. Each segment 86 include a downstream lip portion 88 and an upstream step portion 90. Step portion 90 of each segment 86 overlaps the downstream lip portion 88 of the next adjacent upstream segment 86.
The step portion 90 of the segment 86 adjacent segment 64 overlaps lip portion 69 of segment 64. Cooling means in the form of a plurality of circumferentially spaced-apart apertures 92 are disposed in step portion 90 for admitting cooling air from plenum 54 into the secondary dilution zone 72 of chamber 50 for film cooling downstream segments 86 in the conventional manner well-known in the art.
It is observed that all cooling air introduced into chamber 50 is admitted at a location downstream of primary combustion zone 70. That is to say lips 68, 69, 78 and 88 each introduce cooling air into chamber 50 at a distance greater than the distance d which defines the downstream
end of primary combustion zone 70. Hence
cooling air discharged through apertures 82 and
92 will not quench the burning fuel/air mixture in
the primary combustion zone 70 nor alter the ratio
of fuel to air in zone 70. Said another way,
burning of the fuel/air mixture is substantially
completed upstream of the point at which the
cooling air is introduced into chamber 50.
Cooling means in the form of outer and inner
impingement liners 94 and 96 respectively are
provided for impingement cooling upstream liner segments 62 and 64 respectively. More specifically, circumferentially and axially extending impingement liner 94 is disposed between plenum 52 and liner segment 62 and radially adjacent liner segment 62 so as to cooperate therewith to form circumferentially and axially extending cooling chamber 98. Liner 94 includes a stepped portion 100 defining the upstream end of chamber 98 while the downstream end of chamber 98 is terminated at step portion 80 associated with the first downstream segment 76.
A plurality of apertures 102 provide for the admittance of a cooling air from plenum 52 into chamber 98 and are radially directed to provide impingement of cooling air upon the radially outer surface of liner segment 62. After impingement cooling liner segment 62 in this manner, the cooling air is discharged away from the primary combustion zone 70 through cooling air discharge means, in the form of apertures 82, disposed in the first upstream segment 76 of stepped wall 74 to provide film cooling of the liner segment 76 in zone 72.
Circumferentially and axially extending impingement liner 96 is disposed between plenum
54 and liner segment 64 and radially adjacent liner segment 64 so as to defined therebetween a circumferentially and axially extending cooling chamber 104. Stepped portion 106 of liner 96 forms the upstream end of chamber 104 while the downstream end of chamber 104 is terminated at step portion 90 of downstream segment 86.
Cooling air from plenum 54 is admitted to chamber 104 through cooling means in the form of a plurality of circumferentially extending spaced-apart radially oriented apertures 107 in liner 96. Cooling air admitted to chamber 104 impinges upon the radially inner surface of liner segment 64. Impingement cooling of liner segment 64 is thereby accomplished and thence the cooling air is discharged away from primary combustion zone 70 through cooling air discharge means, in the form of apertures 92, disposed in the first upstream segment 86 of stepped wall 84 to provide film cooling of the liner segment 86 in zone 72.
It may be readily observed that the continuous construction of segments 62 and 64 precludes passage of cooling airfrom impingement chambers 98 and 104 from entering primary combustion zone 70. In this manner then - the present invention provides means for preventing cooling air used to cool segments 62 and 64 from discharge into the primary
combustion zone.
It is clear that the structure hereinbefore
described is well adapted to accomplish the
objects of the present invnetion. More specifically,
the present invention provides combustion
apparatus wherein the liners associated with the
combustion zone are impingment cooled proximate
the primary combustion zone and film cooled
proximate the secondary combustion zone. The
cooling air used in impingement cooling is
discharged into the combustion zone downstream
of the primary combustion zone and hence does
not contribute to the formation of undesirable
gaseous emissions in the primary combustion
zone. Elimination of film cooling in the primary
combustion zone achieves reduced levels of
gaseous emissions present in prior art combustors
for a number of principal reasons.First, the
quenching zone present in the state of the art
combustors is eliminated thus increasing the
volume available for primary combustion.
Secondly, since film cooling requires large
quantities of air moving at substantial velocities,
elimination of film cooling in the primary
combustion zone increases both the average
temperature of combustion and the residence time
of the burning gases in the primary combustion
zone. More complete combustion is therefore
promoted. Thirdly, the ratio of fuel to air in the
primary combustion zone is increased to provide a
rich mixture of fuel and air which tends to form
lower emissions of carbon monoxide and
unburned hydrocarbons.
understood that- modifications or alterations of the
invention may be made within the spirit of my
invention or within the scope of the invention as
set forth in the appended claims.
Claims (7)
1. For use in reducing the gaseous emissions of
a combustor assembly, the invention comprising:
combustor liner means for defining a
longitudinally extending combustion chamber, a
first portion of said liner means defining a primary
combustion zone within said chamber and a
second portion of said liner means defining a
secondary dilution zone within said chamber
downstream of said primary combustion zone;
a fuel nozzle and swirler assembly for
introducing a mixture of fuel and air into said
chamber for burning of said mixture in said
chamber;
first cooling means providing for cooling of said
first portion of said liner means with air;
means for preventing air used in cooling said
first portion from being discharged into said
primary combustion zone; and
second cooling means providing for cooling of
said second portion of said liner means with air.
2. The invention as set forth in Claim 1 wherein
said air used to cool said first portion is discharged
away from said primary combustion zone and into
said secondary dilution zone.
3. The invention as set forth in Claim 2 wherein
said secondary means discharges said air used to cool said second portion into said secondary dilution zone.
4. For use in reducing the gaseous emissions of an annular combustor assembly associated with a high pressure ratio gas turbine engine, the invention comprising:
first and second radially spaced-apart annular combustor liners defining a longitudinally extending combustion chambertherebetween each of said first and second liners including an upstream segment defining a primary combustion zone within said combustion chamber and a downstream segment defining a secondary dilution zone within said combustion chamber;
a fuel nozzle and swirler assembly adapted to introduce a mixture of fuel and air into said chamber for burning of said mixture in said primary zone;;
first and second impingement liners cooperating with and spaced apart from said upstream segments, to define first and second cooling chambers radially adjacent each of said upstream segments, each of said impingement liners including at least one aperture disposed therein for admitting cooling air into said cooling chamber for cooling of said upstream segments;
cooling air discharge means associated with each of said cooling chambers for exhausting said cooling air from said cooling chambers into said secondary dilution zone; and
cooling means for providing for cooling of said downstream segments with air.
5. For use in reducing gaseous emissions of a combustor assembly associated with a high pressure ratio gas turbine engine, the invention comprising:
first and second radially spaced-apart combustor liners defining a longitudinally extending combustion chamber therebetween, said chamber including at its upstream end a primary combustion zone wherein fuel and air entering said combustor are substantially burned;
a fuel nozzle and swirler assembly disposed upstream of said zone for introducing a mixture of fuel and air into said primary zone, said mixture flowing downstream through said primary zone and burning within said primary zone;
means for air cooling a first portion of said liners radially adjacent said primary zone;
means for preventing discharge of said cooling air into said primary combustion zone; and
means for discharging said cooling air into said chamber at a preselected distance downstream of said assembly, said distance having a magnitude preselected to provide for substantially complete burning of said mixture upstream of said preselected distance.
6. The invention as set forth in Claim 5 wherein said distance is substantially equal to said radial spacing between said liners proximate the upstream end of said primary zone.
7. The combination substantially in accordance with any embodiment (or modification thereof) of the invention claimed in Claim 1,4 or 5 and described and/or illustrated herein.
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US95759178A | 1978-11-03 | 1978-11-03 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| GB2034874A true GB2034874A (en) | 1980-06-11 |
Family
ID=25499810
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB7933578A Withdrawn GB2034874A (en) | 1978-11-03 | 1979-09-27 | Gas turbine engine combustor |
Country Status (5)
| Country | Link |
|---|---|
| JP (1) | JPS5582231A (en) |
| DE (1) | DE2944139A1 (en) |
| FR (1) | FR2440524A1 (en) |
| GB (1) | GB2034874A (en) |
| IT (1) | IT1124886B (en) |
Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2531748A1 (en) * | 1982-08-16 | 1984-02-17 | Gen Electric | DOUBLE FUNCTION COOLING AIR COMBUSTION SHIRT |
| GB2134243A (en) * | 1983-01-27 | 1984-08-08 | Rolls Royce | Combustion equipment for a gas turbine engine |
| GB2161914A (en) * | 1980-12-10 | 1986-01-22 | Rolls Royce | Combustion equipment for a gas turbine engine |
| US4566280A (en) * | 1983-03-23 | 1986-01-28 | Burr Donald N | Gas turbine engine combustor splash ring construction |
| US5381652A (en) * | 1992-09-24 | 1995-01-17 | Nuovopignone | Combustion system with low pollutant emission for gas turbines |
| EP1253378A2 (en) * | 2001-04-24 | 2002-10-30 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor having bypass passage |
| US6679063B2 (en) | 2000-10-02 | 2004-01-20 | Rolls-Royce Deutschland Ltd & Co Kg | Combustion chamber head for a gas turbine |
| EP2261565A1 (en) * | 2009-06-09 | 2010-12-15 | Siemens Aktiengesellschaft | Gas turbine reactor and gas turbines |
Families Citing this family (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2604509B1 (en) * | 1986-09-25 | 1988-11-18 | Snecma | PROCESS FOR PRODUCING A COOLING FILM FOR A TURBOMACHINE COMBUSTION CHAMBER, FILM THUS PRODUCED AND COMBUSTION CHAMBER COMPRISING SAME |
| US4916906A (en) * | 1988-03-25 | 1990-04-17 | General Electric Company | Breach-cooled structure |
| DE19547703C2 (en) * | 1995-12-20 | 1999-02-18 | Mtu Muenchen Gmbh | Combustion chamber, in particular ring combustion chamber, for gas turbine engines |
| FR2901349B1 (en) * | 2006-05-19 | 2008-09-05 | Snecma Sa | COMBUSTION CHAMBER OF A TURBOMACHINE |
| RU2451881C2 (en) * | 2009-10-06 | 2012-05-27 | Открытое акционерное общество "Всероссийский дважды ордена Трудового Красного Знамени теплотехнический научно-исследовательский институт" | Premixing combustion chamber of gas turbine plant |
Family Cites Families (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2482E (en) * | 1902-04-30 | 1904-04-27 | Gaston Charles Emile De Bonnec | Continuous combustion turbo-engine system |
| CH255541A (en) * | 1947-05-12 | 1948-06-30 | Bbc Brown Boveri & Cie | Cooled metal combustion chamber for generating heating and propellant gases. |
| US2729062A (en) * | 1951-07-06 | 1956-01-03 | Dresser Operations Inc | Power plant combustion chamber |
| US2974485A (en) * | 1958-06-02 | 1961-03-14 | Gen Electric | Combustor for fluid fuels |
| GB845971A (en) * | 1958-07-21 | 1960-08-24 | Gen Electric | Improvements relating to combustion chambers for gas turbine engines |
| US3338051A (en) * | 1965-05-28 | 1967-08-29 | United Aircraft Corp | High velocity ram induction burner |
| US3369363A (en) * | 1966-01-19 | 1968-02-20 | Gen Electric | Integral spacing rings for annular combustion chambers |
| FR2055720A1 (en) * | 1969-08-01 | 1971-04-30 | Snecma | |
| CA980584A (en) * | 1972-11-10 | 1975-12-30 | Edward E. Ekstedt | Double walled impingement cooled combustor |
| US3986347A (en) * | 1973-12-06 | 1976-10-19 | Phillips Petroleum Company | Combustor process for low-level NOx and CO emissions |
| GB1498269A (en) * | 1975-05-23 | 1978-01-18 | Snecma | Gas turbine combustion chambers |
| DE2636520A1 (en) * | 1976-08-13 | 1978-02-16 | Daimler Benz Ag | Spherical combustion chamber for gas turbines - has spherical secondary wall with inlets accelerating swirling in combustion chamber |
-
1979
- 1979-09-27 GB GB7933578A patent/GB2034874A/en not_active Withdrawn
- 1979-10-31 JP JP13998779A patent/JPS5582231A/en active Pending
- 1979-10-31 IT IT26982/79A patent/IT1124886B/en active
- 1979-10-31 FR FR7927035A patent/FR2440524A1/en not_active Withdrawn
- 1979-11-02 DE DE19792944139 patent/DE2944139A1/en not_active Withdrawn
Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2161914A (en) * | 1980-12-10 | 1986-01-22 | Rolls Royce | Combustion equipment for a gas turbine engine |
| US4610135A (en) * | 1980-12-10 | 1986-09-09 | Rolls-Royce Plc | Combustion equipment for a gas turbine engine |
| FR2531748A1 (en) * | 1982-08-16 | 1984-02-17 | Gen Electric | DOUBLE FUNCTION COOLING AIR COMBUSTION SHIRT |
| GB2134243A (en) * | 1983-01-27 | 1984-08-08 | Rolls Royce | Combustion equipment for a gas turbine engine |
| US4566280A (en) * | 1983-03-23 | 1986-01-28 | Burr Donald N | Gas turbine engine combustor splash ring construction |
| US5381652A (en) * | 1992-09-24 | 1995-01-17 | Nuovopignone | Combustion system with low pollutant emission for gas turbines |
| US6679063B2 (en) | 2000-10-02 | 2004-01-20 | Rolls-Royce Deutschland Ltd & Co Kg | Combustion chamber head for a gas turbine |
| EP1253378A2 (en) * | 2001-04-24 | 2002-10-30 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor having bypass passage |
| EP1253378A3 (en) * | 2001-04-24 | 2003-10-08 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor having bypass passage |
| US6860098B2 (en) * | 2001-04-24 | 2005-03-01 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor having bypass and annular gas passage for reducing uneven temperature distribution in combustor tail cross section |
| EP2261565A1 (en) * | 2009-06-09 | 2010-12-15 | Siemens Aktiengesellschaft | Gas turbine reactor and gas turbines |
Also Published As
| Publication number | Publication date |
|---|---|
| FR2440524A1 (en) | 1980-05-30 |
| DE2944139A1 (en) | 1980-05-14 |
| IT1124886B (en) | 1986-05-14 |
| JPS5582231A (en) | 1980-06-20 |
| IT7926982A0 (en) | 1979-10-31 |
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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) |