Connect public, paid and private patent data with Google Patents Public Datasets

Gas turbine for aircrafts

Download PDF

Info

Publication number
US3651645A
US3651645A US3651645DA US3651645A US 3651645 A US3651645 A US 3651645A US 3651645D A US3651645D A US 3651645DA US 3651645 A US3651645 A US 3651645A
Authority
US
Grant status
Grant
Patent type
Prior art keywords
heat
turbine
means
cooling
air
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.)
Expired - Lifetime
Application number
Inventor
Hubert Josef Grieb
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.)
Daimler-Benz AG
MTU Aero Engines GmbH
Original Assignee
MTU Aero Engines GmbH
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
Grant date

Links

Images

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
    • F02C7/00Features, 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
    • F02C7/12Cooling of plants
    • F02C7/14Cooling of plants of fluids in the plant, e.g. lubricant or fuel
    • YGENERAL 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
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies
    • Y02T50/67Relevant aircraft propulsion technologies
    • Y02T50/671Measures to reduce the propulsor weight
    • YGENERAL 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
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies
    • Y02T50/67Relevant aircraft propulsion technologies
    • Y02T50/675Enabling an increased combustion temperature by cooling

Abstract

A gas turbine for aircrafts, particularly jet engines with aircooled turbine blades whereby the cooling air is taken off either directly or indirectly from a compressor stage; a heat-exchanger with a closed circulation is provided in the cooling system whose heat-absorbing part is arranged in a cooling air channel disposed between the compressor and the turbine rotor.

Description

United States Patent Grieb 51 Mar. 28, 1972 GAS TURBINE FOR AIRCRAFTS 21 Appl. No.: 79,508

3,301,526 1/1967 Chamberlain ..60/39.66 2,992,529 7/1961 Sampletro ..60/39.66 3,083,532 4/1963 Cook ..60/39.66 3,253,406 5/1966 Grieb... .....60/262 3,418,808 12/1968 Rich ..60/226 3,437,313 4/1969 Moore ..60/39.66 3,584,458 6/1971 Wetzler ..60/39.66

FOREIGN PATENTS OR APPLICATIONS 861 ,632 2/1961 Great Britain ..60/39.66

[30] Foreign Application Pnonty Data Primary EXaminer Mark M Newman Oct. 11, 1969 Germany ..P 19 51 356.5 Assistant ExaminerRichard Rothman Attorney-Craig, Antonelli, Stewart & Hill [52] U.S. Cl ..60/262, 60/3966 [51] Int. Cl. ..F02c 7/12, F02c 7/14 [57] ABSTRACT [58] Field oISearch ..60/262, 226, 39665523655651, A gas turbine for aircrafts particularly jet engines with aib cooled turbine blades whereby the cooling air is taken off [56] References Cited either directly or indirectly from a compressor stage; a heatexchanger with a closed circulation is provided in the cooling UNITED STATES PATENTS system whose heat-absorbing part is arranged in a cooling air channel disposed between the compressor and the turbine ro- 2,465,099 3/1949 Johnson ..60/262 X ton 2,501,633 3/1950 Price ...60/262 2,703,477 3/1955 Anxionnaz ..60/262 16 Claims, 4 Drawing Figures N. :l S

PATENTEDMAR28 1972 3. 651 ,645

SHEET 1 OF 2 INVENTOR HUBERT J. GRIEB ATTORNEYS PATENTED MR 2 8 I972 SHEET 2 [IF 2 GAS TURBINE FOR AIRCRAFIS The present invention relates to a gas turbine for aircrafts, especially to a jet engine, with air-cooled turbine blades whose cooling air is taken off either directly or indirectly from a compressor stage. With such types of gas turbines, one aims at an increase of the output by an increase of the turbine inlet temperature of the operating gases. However, limitations are imposed thereon by the heat resistance of the available materials. The use of strongly compressed air for the improvement of the blade cooling results in a favorable pressure drop for the cooling air stream, however, on the other hand, it reduces the cooling effect by the heating of the cooling air connected with the compression. The present invention has as its purpose to eliminate this shortcoming and to enable an output increase of the gas turbine by an effective cooling of the turbine blades. This is realized according to the present invention by a heatexchanger with closed circulation, whose heat-absorbing part is arranged in the cooling air channel between the compressor and the turbine rotor. By the use of such an arrangement a good blade-cooling can be achieved also with strongly compressed cooling air so that also high turbine inlet temperatures are possible. In this manner, an increase of the turbine output or, with with the same output, a smaller structural weight and smaller dimensions of the gas turbine are achieved by the present invention.

The heat-exchanger may be so arranged according to the present invention in the gas turbine that the heat-absorbing part of the heat-exchanger is arranged in an annular space between the rotor shaft and the combustion chambers, which part is connected on the one hand, with the compressor outlet diffusors and, on the other, with the blades at least of the first turbine stage. Oneobtains with this arrangement highly compressed air for the purpose of blade cooling, which is able to absorb a sufficient amount of heat and therebeyond produces a good pressure drop. By reason of the fact that only the thermally most strongly stressed turbine stages are supplied with cooled compressor air, one is able to get along in an advantageous manner with a small heat-exchanger so that the structural expenditures are of no significance in comparison to the output gain.

Theheat-releasing part of the heat-exchanger, i.e., the part of the heat-exchanger that gives off heat, may be arranged according to the present invention at the inlet of the compressor. This has the advantage that the air flowing past this place has its lowest temperature. According to another embodiment of the present invention for ducted-fan-jet power plants or engines, the heat-releasing part of the heat-exchanger which gives off the heat, is arranged in the bypass channel. This arrangement makes possible with slight additional structural expenditures a favorable heat transfer to the air flowing in the bypass channel.

Accordingly, it is an object of the present invention to provide a gas turbine for aircrafts which avoids by simple means the aforementioned shortcomings and drawbacks encountered in the prior art.

Another object of the present invention resides in a gas turbine which permits an increase in the output by an increase of the turbine inlet temperature of the working gases.

A further object of thepresent invention resides in a gas turbine for aircrafts which allows a good cooling of the blades also with a strongly compressed cooling air.

A still further. object of the present invention resides in a gas turbine for jet engines which results in a relatively slight structural weight and small dimensions of the turbine for the relatively high output of the turbine.

These and further objects, features and advantages of the present invention will become more obvious from the following description when taken in connection with the accompanying drawing which shows, for purposes of illustration only, several embodiments in accordance with the present invention, and wherein:

FIG. 1 is a schematic longitudinal cross-sectional view through a pure jet engine in accordance with the present invention;

FIG. 2 is a schematic longitudinal cross-sectional view through a ducted-fan-jet engine in accordance with the present invention;

FIG. 3 is-a partial longitudinalcross-sectional view illustrating the heat-exchanger for the blade cooling system of a jet engine according to FIG. 2; and

FIG. 4 is a schematic control diagram of a heat-exchanger in accordance with the present invention.

Referring now to the drawing, an d more particularly to FIG. 1, a low pressure compressor 13 and a low pressure turbine 14 are mounted on and secured to a shaft 12 within the housing 11 of the pure jet engine illustrated in this figure while a high-pressure compressor 16 and a high-pressure turbine 17 are secured on a hollow shaft 15 coaxial to the shaft 12. The combustion chambers of the jet engine are designated by reference numeral 18. The blades 19 of the first turbine stage 20 are cooled by compressed air which is taken off from the last stage 21 of the high-pressure compressor 16. The cooling air is conducted for that purpose through the annular space 22 formed between the hollow shaft 15 and the combustion chambers 18. The heat-absorbing part 23 of a heat-exchanger generally designated by reference numeral 24 with closed circulatory system is arranged in the annular space 22. The heatreleasing or heat-transferring part 25 of the heat-exchanger 24 which gives off heat, is secured at the compressor inlet 26 and is connected with the heat-absorbing part 23 by lines 27 and 28.

In contradistinction thereto, in the ducted-fan-jet engine illustrated in FIG. 2, the heat-transferring or heat-releasing part 29 of a heat-exchanger 30 also with closed circulatory system is accommodated in the bypass channel 33 surrounded by a jacket or casing 31 and delimited inwardly by a housing 32. The heat-absorbing part 34 is arranged, similar as in the example according to FIG. 1, in an annular space 35 between the combustion chambers 36 and a hollow shaft 37. The hollow shaft 37 connects a high pressure compressor 38 with a highpressure turbine 39. A shaft 40 for a low pressure compressor 41 and a low pressure turbine 42 is extended through the hollow shaft 37. Lines 43 and 44 connect the heat-absorbing part 34 with the heat-releasing part 29 of the heat-exchanger 30.

The ducted-fan-jet engine partially illustrated in FIG. 3 corresponds in its essential construction to that of FIG. 2. A shaft 46 for a low-pressure compressor (not shown) and a low-pressure turbine 48 as well as a hollow shaft 49 surrounding the shaft 46 for a high-pressure compressor 50 and a high-pressure turbine 51 are rotatably supported in the housing 45. Reference numeral 52 designates rotor blades and reference numeral 53 guide blades of the last stage generally designated by reference numeral 54 of the high pressure compressor 50. Compressor outlet diffusors 55 terminate in the combustion chambers 56, in each of which are arranged a flame tube 57 of conventional construction with a fuel feed line 58. Guide blades 59 and rotor blades 60 of the first turbine stage generally designated by reference numeral 61 adjoin the same. The bypass channel of the ducted fan-jet engine is designated by reference numeral 62.

The combustion chambers 56 are provided each within the area of the compressor outlet diffusors 55 with apertures 63, through which a part of the compressed air is able to flow over into an annular space 64 disposed between the hollow shaft 49 and the combustion chambers 56. From there, the air is fed by way of an annular channel 66 sealed off by labyrinth seals to the first turbine stage generally designated by reference numeral 61. Therotor blades 60 of this stage 61 have hollow spaces 67 which are in communication with the blade surface so that the compressed air is able to escape and is able to cool the thermally particularly strongly stressed turbine blades 60.

The heat-absorbing part 68 of a heat-exchanger 69 with closed circulation is arranged in the annular space 64. A line 70 leads from the heat-absorbing part 68 within the area of the compressor outlet diffusor55 to the'heat-releasing part 71 which'gives off heat and is arranged in the bypass channel 62 of the jet engine. A line 72 leads from the heat-transferring part 71 to a circulating pump (not shown) for the cooling medium and from there a further line 73, parallel to the line 70, leads back to the heat-absorbing part 68 of the heatexchanger 69. The two lines 70 and 73 are lined on the inside of the compressor outlet diffusor 55 by a hollow rib 74. The heat-absorbing part 68 and the heat-transferring part 71 are arranged in parallelly connected heat-exchanger groups, for example, consisting of pipe coils and uniformly distributed over the circumference of the annular space 64 and the bypass channel 62.

In operation, the heat-absorbing part 68 of the heatexchanger 69 cools the highly compressed and strongly heated air taken off from the compressor outlet diffusor 55, which flows through the annular space 64. The cooling medium heated up thereby, flows through the line 70 to the heat-transferring part 71 of the heat-exchanger 69 and thereby gives off heat to the less strongly compressed and relatively cool air flowing in the bypass channel 62. The cooling medium is again conducted back to the heat-absorbing part 68 by way of the line 73 by means of the circulating pump (not shown).

The cooled compressor air flows from the annular space 64 by way of the annular channel 66 into the hollow spaces of the rotor blades 60 of the first turbine stage 61. The air escapes out of the rotor blades 60 by way of openings (not shown) and cools the thermally highly stressed surfaces thereof. The temperature decrease of the highly compressed cooling air achieved by the heat-exchanger of the present invention improves the blade cooling and permits thereby the application of higher pressures and temperatures of the working gases. This leads to a considerable increase of the specific thrust or permits the construction of more lightweight and smaller jet engines.

The cooling air may be taken off, instead of from the compressor diffusor as shown in the illustrated embodiment, also directly from a compressor stage. Cooled compressor air may, in case of need, be branched off also for further stages in addition to the first turbine stage, whereas the remaining turbine stage, insofar as necessary, are supplied with uncooled compressor air. The turbine guide blades may also be cooled with uncooled compressor air whereby possibly the rate of air flow of the first stage is slightly increased.

The heat-exchanger 75 illustrated in FIG. 4 in a schematic diagram which has a closed circulator system, essentially consists of the heat-absorbing part 76 and of the heat-transferring or heat-releasing part 77 as well as of a circulating pump 78 and the lines 79 and 80. The heat-absorbing part 76 is subdivided into parallel pipe lines 81. The compressed air branched off from the compressor for the blade cooling flows through the pipe lines 81 which is indicated by an arrow 82. The heattransferring part 77 subdivided in a similar manner into pipe lines 83 is located in an air stream indicated by an arrow 84 which, for example, is conducted through the bypass. The circulating pump 78 as well as a charging pump 85 are driven in this example by an electric motor 86.

The charging pump 85 supplies cooling medium under pressure out ofa tank 87 through the lines 88 and 89 into the line 80. The operating pressure is appropriately chosen so high that the boiling temperature of the cooling medium is not attained. A check valve 90 disposed in the line 89 closes after the charging of the circulatory system and thus prevents a return flow of the cooling medium. In case ofa decrease of the pressure, the charging pump 85 supplies cooling medium to the circulatory system until the minimum pressure is again reached. In this manner, also in case of an occurrence of a leakage place, the operation can be maintained at least for a short period of time. An excess pressure valve 91 is disposed in a line 91 leading from the line 88 to the tank 87 which prevents that the pressure of the supplied cooling medium exceeds a predetermined value. Lines 93 and 94 conduct cooling medium that is discharged out of the bearings 95 and 96 of the circulating pump 78 and of the charging pump 85, respectively, back into the tank 87. v

A line 97 is connected to the line 79 which branches off into lines 98 and 99 that lead to an equalization tank generally designated by reference numeral 100 and to an excess pressure valve 101, respectively. The equalization tank 100 absorbs against the pressure of an air cushion 102 the excess cooling liquid caused by heat expansion and returns the same again to the circulatory system when the cooling liquid cools off again. The excess pressure valve 101 opens a return line 103, when the permissive operating pressure is exceeded, and thus avoids damages in the installation.

A warning installation 104 consists of a pressure-measuring device 105 disposed in the circulatory system which closes the energizing circuit of a warning lamp 106 when the pressure drops below a minimum operating pressure. The illumination of the warning lamp 106 indicates that the cooling circulatory system no longer operates satisfactorily. With a turbine operating at full load, the fuel supply can be timely throttled thereafter so that an overheating of the turbine is avoided. in the same manner, also the operation of the circulating pump can be monitored by a differential pressure-measuring device connected in parallel thereto which in case of a pump damage cases a warning lamp to light up. It is also possible to automatically limit the fuel supply by any conventional means in case of failure of the cooling circulatory system.

Water may be utilized a cooling liquid whose properties can be improved by suitable additives and/or admixtures, for example, for the purpose of increasing its lubricating capacity. For flights at high altitude and during the winter, an antifreeze agent, for example, glycol, has to be admixed to the cooling water.

While I have shown and described several embodiments in accordance with the present invention, it is understood that the same is not limited thereto, but is susceptible of numerous changes and modifications as known to those skilled in the art, and I therefore do not wish to be limited to the details shown and described herein, but intend to cover all such change and modifications as are encompassed by the scope of the appended claims.

What I claim is:

1. A gas turbine for aircraft having a compressor means and air-cooled turbine blade means whose cooling air is taken off from a compressor stage, characterized by a heat-exchanger means with closed circulatory system whose heat-absorbing part is arranged in a cooling air channel means between the compressor means and a turbine rotor means.

2. A gas turbine according to claim 1, characterized in that the cooling air is directly taken off from a compressor stage.

3. A gas turbine according to claim 1, characterized in that the cooling air is indirectly taken off from a compressor stage.

4. A gas turbine with a rotor shaft and combustion chamber means according to claim 1, characterized in that the heat-absorbing part of the heat-exchanger means is arranged in an annular space between the rotor shaft and the combustion chamber means, said annular space being in communication, on the one hand, with compressor outlet diffusor means and, on the other, with rotor blade means.

5. A gas turbine with several turbine stages according to claim 4, characterized in that said rotor blade means are part of at least the first turbine stage.

6. A gas turbine according to claim 4, characterized in that the heat-exchanger means includes a heat-transferring part giving off heat which is arranged at the air inlet of the compressor means.

7. A gas turbine according to claim 4, for a ducted fan-jet engine, characterized in that the heat-exchanger means includes a heat-transferring part giving off heat whichds arranged in a bypass channel means of the jet engine.

8. A gas turbine according to claim 4, with a circulating pump means for the closed circulatory system, characterized in that the minimum pressure of the cooling medium is maintained by a charging pump means operatively connected with the circulatory system.

9. A gas turbine according to claim 8, characterized in that an equalization tank means is connected to the circulatory system.

10. A gas turbine according to claim 9, characterized by a warning device which becomes operable when the pressure drops below a predetermined pressure.

11. A gas turbine according to claim 10, characterized in that the warning device becomes operative by illumination of 5 a warning lamp.

12. A gas turbine according to claim 1, characterized in that the heat-exchanger means includes a heat-transferring part giving off heat which is arranged at the air inlet of the compressor means.

13. A gas turbine according to claim 1, for a ducted fan-jet engine, characterized in that the heat-exchanger means includes a heat-tranferring part giving off heat which is arranged in a bypass channel means of the jet engine.

14. A gas turbine according to claim 1, with a circulating pump means for the closed circulatory system, characterized in that the minimum pressure of the cooling medium is maintained by a charging pump means operatively connected with the circulatory system.

15. A gas turbine according to claim 14, characterized in that an equalization tank means is connected to the circulatory system.

16. A gas turbine according to claim 15, characterized by a warning device which becomes operable when the pressure drops below a predetermined pressure.

Claims (16)

1. A gas turbine for aircraft having a compressor means and aircooled turbine blade means whose cooling air is taken off from a compressor stage, characterized by a heat-exchanger means with closed circulatory system whose heat-absorbing part is arranged in a cooling air channel means between the compressor means and a turbine rotor means.
2. A gas turbine according to claim 1, characterized in that the cooling air is directly taken off from a compressor stage.
3. A gas turbine according to claim 1, characterized in that the cooling air is indirectly taken off from a compressor stage.
4. A gas turbine with a rotor shaft and combustion chamber means according to claim 1, characterized in that the heat-absorbing part of the heat-exchanger means is arranged in an annular space between the rotor shaft and the combustion chamber means, said annular space being in communication, on the one hand, with compressor outlet diffusor means and, on the other, with rotor blade means.
5. A gas turbine with several turbine stages according to claim 4, characterized in that said rotor blade means are part of at least the first turbine stage.
6. A gas turbine according to claim 4, characterized in that the heat-exchanger means includes a heat-transferring part giving off heat which is arranged at the air inlet of the compressor means.
7. A gas turbine according to claim 4, for a ducted fan-jet engine, characterized in that the heat-exchanger means includes a heat-transferring part giving off heat which is arranged in a by-pass channel means of the jet engine.
8. A gas turbine according to claim 4, with a circulating pump means for the closed circulatory system, characterized in that the minimum pressure of the cooling medium is maintained by a charging pump means operatively connected with the circulatory system.
9. A gas turbine according to claim 8, characterized in that an equalization tank means is connected to the circulatory system.
10. A gas turbine according to claim 9, characterized by a warning device which becomes operable when the pressure drops below a predetermined pressure.
11. A gas turbine according to claim 10, characterized in that the warning device becomes operative by illumination of a warning lamp.
12. A gas turbine according to claim 1, characterized in that the heat-exchanger means includes a heat-transferring part giving off heat which is arranged at the air inlet of the compressor means.
13. A gas turbine according to claim 1, for a ducted fan-jet engine, characterized in that the heat-exchanger means includes a heat-tranferring part giving off heat which is arranged in a by-pass channel means of the jet engine.
14. A gas turbine according to claim 1, with a circulating pump means for the closed circulatory system, characterized in that the minimum pressure of the cooling medium is maintained by a charging pump means operatively connected with the circulatory system.
15. A gas turbine according to claim 14, characterized in that an equalization tank means is connected to the circulatory system.
16. A gas turbine according to claim 15, characterized by a warning device which becomes operable when the pressure drops below a predetermined pressure.
US3651645A 1969-10-11 1970-10-09 Gas turbine for aircrafts Expired - Lifetime US3651645A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
DE19691951356 DE1951356C3 (en) 1969-10-11 1969-10-11

Publications (1)

Publication Number Publication Date
US3651645A true US3651645A (en) 1972-03-28

Family

ID=5747939

Family Applications (1)

Application Number Title Priority Date Filing Date
US3651645A Expired - Lifetime US3651645A (en) 1969-10-11 1970-10-09 Gas turbine for aircrafts

Country Status (4)

Country Link
US (1) US3651645A (en)
DE (1) DE1951356C3 (en)
FR (1) FR2065179A5 (en)
GB (1) GB1287983A (en)

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5440910A (en) * 1977-07-25 1979-03-31 Gen Electric Method of cooling blade of gas turbine engine and its device
US4195474A (en) * 1977-10-17 1980-04-01 General Electric Company Liquid-cooled transition member to turbine inlet
US4254618A (en) * 1977-08-18 1981-03-10 General Electric Company Cooling air cooler for a gas turbofan engine
US4991394A (en) * 1989-04-03 1991-02-12 Allied-Signal Inc. High performance turbine engine
US5003773A (en) * 1989-06-23 1991-04-02 United Technologies Corporation Bypass conduit for gas turbine engine
US5012646A (en) * 1988-11-28 1991-05-07 Machen, Inc. Turbine engine having combustor air precooler
US5697208A (en) * 1995-06-02 1997-12-16 Solar Turbines Incorporated Turbine cooling cycle
WO2002038938A1 (en) * 2000-11-10 2002-05-16 Kovac Marek Bypass gas turbine engine and cooling method for working fluid
US6510684B2 (en) * 2000-05-31 2003-01-28 Honda Giken Kogyo Kabushiki Kaisha Gas turbine engine
US20050050877A1 (en) * 2003-09-05 2005-03-10 Venkataramani Kattalaicheri Srinivasan Methods and apparatus for operating gas turbine engines
US20070022732A1 (en) * 2005-06-22 2007-02-01 General Electric Company Methods and apparatus for operating gas turbine engines
EP1881182A2 (en) 2006-07-19 2008-01-23 Snecma Cooling system for a downstream cavity of a centrifugal compressor impeller
US20080141954A1 (en) * 2006-12-19 2008-06-19 United Technologies Corporation Vapor cooling of detonation engines
US20080310955A1 (en) * 2007-06-13 2008-12-18 United Technologies Corporation Hybrid cooling of a gas turbine engine
US20100107649A1 (en) * 2007-03-28 2010-05-06 Ulf Nilsson Gas Turbine Engine With Fuel Booster
US20100263350A1 (en) * 2009-04-17 2010-10-21 Yang Liu Apparatus and method for cooling a turbine using heat pipes
US20100319892A1 (en) * 2008-04-02 2010-12-23 United Technologies Corporation Heat exchanging structure
US20110296845A1 (en) * 2009-01-28 2011-12-08 Jonathan Jay Felnstein Combined heat and power with a peak temperature heat load
US20140345292A1 (en) * 2013-05-22 2014-11-27 General Electric Company Return fluid air cooler system for turbine cooling with optional power extraction
US20140352315A1 (en) * 2013-05-31 2014-12-04 General Electric Company Cooled cooling air system for a gas turbine
EP3163052A1 (en) * 2015-10-26 2017-05-03 General Electric Company Method and system for managing heat flow in an engine

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3514352A1 (en) * 1985-04-20 1986-10-23 Mtu Muenchen Gmbh Gas turbine engine with facilities for diversion of compressor air for cooling of hot share
FR2656657B1 (en) * 1989-12-28 1994-04-22 Snecma
US8845819B2 (en) * 2008-08-12 2014-09-30 General Electric Company System for reducing deposits on a compressor

Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2465099A (en) * 1943-11-20 1949-03-22 Allis Chalmers Mfg Co Propulsion means comprising an internal-combustion engine and a propulsive jet
US2501633A (en) * 1943-06-28 1950-03-21 Lockheed Aircraft Corp Gas turbine aircraft power plant having ducted propulsive compressor means
US2703477A (en) * 1951-03-16 1955-03-08 Rateau Soc Rotary jet propulsion unit
GB861632A (en) * 1958-06-25 1961-02-22 Rolls Royce Method and apparatus for cooling a member such, for example, as a turbine blade of agas turbine engine
US2992529A (en) * 1956-08-23 1961-07-18 Thompson Ramo Wooldridge Inc Turbine blade cooling
US3083532A (en) * 1953-09-07 1963-04-02 Rolls Royce Gas turbine engine with air-cooling means and means to control the temperature of cooling air by liquid injection
US3253406A (en) * 1966-05-31 Turbine propulsion unit
US3301526A (en) * 1964-12-22 1967-01-31 United Aircraft Corp Stacked-wafer turbine vane or blade
US3418808A (en) * 1966-07-05 1968-12-31 Rich David Gas turbine engines
US3437313A (en) * 1966-05-18 1969-04-08 Bristol Siddeley Engines Ltd Gas turbine blade cooling
US3584458A (en) * 1969-11-25 1971-06-15 Gen Motors Corp Turbine cooling

Patent Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3253406A (en) * 1966-05-31 Turbine propulsion unit
US2501633A (en) * 1943-06-28 1950-03-21 Lockheed Aircraft Corp Gas turbine aircraft power plant having ducted propulsive compressor means
US2465099A (en) * 1943-11-20 1949-03-22 Allis Chalmers Mfg Co Propulsion means comprising an internal-combustion engine and a propulsive jet
US2703477A (en) * 1951-03-16 1955-03-08 Rateau Soc Rotary jet propulsion unit
US3083532A (en) * 1953-09-07 1963-04-02 Rolls Royce Gas turbine engine with air-cooling means and means to control the temperature of cooling air by liquid injection
US2992529A (en) * 1956-08-23 1961-07-18 Thompson Ramo Wooldridge Inc Turbine blade cooling
GB861632A (en) * 1958-06-25 1961-02-22 Rolls Royce Method and apparatus for cooling a member such, for example, as a turbine blade of agas turbine engine
US3301526A (en) * 1964-12-22 1967-01-31 United Aircraft Corp Stacked-wafer turbine vane or blade
US3437313A (en) * 1966-05-18 1969-04-08 Bristol Siddeley Engines Ltd Gas turbine blade cooling
US3418808A (en) * 1966-07-05 1968-12-31 Rich David Gas turbine engines
US3584458A (en) * 1969-11-25 1971-06-15 Gen Motors Corp Turbine cooling

Cited By (35)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5440910A (en) * 1977-07-25 1979-03-31 Gen Electric Method of cooling blade of gas turbine engine and its device
JPS628615B2 (en) * 1977-07-25 1987-02-24 Gen Electric
US4254618A (en) * 1977-08-18 1981-03-10 General Electric Company Cooling air cooler for a gas turbofan engine
US4195474A (en) * 1977-10-17 1980-04-01 General Electric Company Liquid-cooled transition member to turbine inlet
US5012646A (en) * 1988-11-28 1991-05-07 Machen, Inc. Turbine engine having combustor air precooler
US4991394A (en) * 1989-04-03 1991-02-12 Allied-Signal Inc. High performance turbine engine
US5003773A (en) * 1989-06-23 1991-04-02 United Technologies Corporation Bypass conduit for gas turbine engine
US5697208A (en) * 1995-06-02 1997-12-16 Solar Turbines Incorporated Turbine cooling cycle
US6510684B2 (en) * 2000-05-31 2003-01-28 Honda Giken Kogyo Kabushiki Kaisha Gas turbine engine
WO2002038938A1 (en) * 2000-11-10 2002-05-16 Kovac Marek Bypass gas turbine engine and cooling method for working fluid
US20050050877A1 (en) * 2003-09-05 2005-03-10 Venkataramani Kattalaicheri Srinivasan Methods and apparatus for operating gas turbine engines
US6990797B2 (en) * 2003-09-05 2006-01-31 General Electric Company Methods and apparatus for operating gas turbine engines
US20070022732A1 (en) * 2005-06-22 2007-02-01 General Electric Company Methods and apparatus for operating gas turbine engines
EP1881182A3 (en) * 2006-07-19 2009-02-18 Snecma Cooling system for a downstream cavity of a centrifugal compressor impeller
US20080019829A1 (en) * 2006-07-19 2008-01-24 Snecma System for cooling a downstream cavity of a centrifugal compressor impeller
FR2904034A1 (en) * 2006-07-19 2008-01-25 Snecma Sa cooling system of a centrifugal compressor impeller downstream cavity.
JP2008025580A (en) * 2006-07-19 2008-02-07 Snecma Cooling system of downstream side cavity of impeller of centrifugal compressor
EP1881182A2 (en) 2006-07-19 2008-01-23 Snecma Cooling system for a downstream cavity of a centrifugal compressor impeller
US8029238B2 (en) 2006-07-19 2011-10-04 Snecma System for cooling a downstream cavity of a centrifugal compressor impeller
US7748211B2 (en) * 2006-12-19 2010-07-06 United Technologies Corporation Vapor cooling of detonation engines
US20080141954A1 (en) * 2006-12-19 2008-06-19 United Technologies Corporation Vapor cooling of detonation engines
US8448447B2 (en) * 2007-03-28 2013-05-28 Siemens Aktiengesellschaft Gas turbine engine with fuel booster
US20100107649A1 (en) * 2007-03-28 2010-05-06 Ulf Nilsson Gas Turbine Engine With Fuel Booster
US8656722B2 (en) 2007-06-13 2014-02-25 United Technologies Corporation Hybrid cooling of a gas turbine engine
US20080310955A1 (en) * 2007-06-13 2008-12-18 United Technologies Corporation Hybrid cooling of a gas turbine engine
US8056345B2 (en) * 2007-06-13 2011-11-15 United Technologies Corporation Hybrid cooling of a gas turbine engine
US20100319892A1 (en) * 2008-04-02 2010-12-23 United Technologies Corporation Heat exchanging structure
US20110296845A1 (en) * 2009-01-28 2011-12-08 Jonathan Jay Felnstein Combined heat and power with a peak temperature heat load
US8112998B2 (en) 2009-04-17 2012-02-14 General Electric Company Apparatus and method for cooling a turbine using heat pipes
US20100263350A1 (en) * 2009-04-17 2010-10-21 Yang Liu Apparatus and method for cooling a turbine using heat pipes
US20140345292A1 (en) * 2013-05-22 2014-11-27 General Electric Company Return fluid air cooler system for turbine cooling with optional power extraction
US9429072B2 (en) * 2013-05-22 2016-08-30 General Electric Company Return fluid air cooler system for turbine cooling with optional power extraction
US20140352315A1 (en) * 2013-05-31 2014-12-04 General Electric Company Cooled cooling air system for a gas turbine
US9422063B2 (en) * 2013-05-31 2016-08-23 General Electric Company Cooled cooling air system for a gas turbine
EP3163052A1 (en) * 2015-10-26 2017-05-03 General Electric Company Method and system for managing heat flow in an engine

Also Published As

Publication number Publication date Type
FR2065179A5 (en) 1971-07-23 application
GB1287983A (en) 1972-09-06 application
DE1951356B2 (en) 1979-12-13 application
DE1951356C3 (en) 1980-08-28 grant
DE1951356A1 (en) 1971-04-29 application

Similar Documents

Publication Publication Date Title
US3269120A (en) Gas turbine engine with compressor and turbine passages in a single rotor element
US3300965A (en) Gas turbine engine fuel heating and oil cooling system
US3238719A (en) Liquid cooled gas turbine engine
US3500636A (en) Gas turbine plants
US3734639A (en) Turbine cooling
US3382670A (en) Gas turbine engine lubrication system
US3450109A (en) Supercharged diesel motors
US3600890A (en) Turbine cooling construction
US3658153A (en) Lubricating oil system for a prime mover
US3705496A (en) Reaction propulsion engine and method of operation
US3220697A (en) Hollow turbine or compressor vane
US3831373A (en) Pumped air storage peaking power system using a single shaft gas turbine-generator unit
US4317439A (en) Cooling system
US4561246A (en) Bearing compartment for a gas turbine engine
US4107927A (en) Ebullient cooled turbocharger bearing housing
US4351150A (en) Auxiliary air system for gas turbine engine
US6378293B1 (en) Gas turbine engine bearing arrangement
US2219994A (en) Gas turbine plant and regulating system therefor
US6295803B1 (en) Gas turbine cooling system
US3034298A (en) Turbine cooling system
US4041697A (en) Oil cooling system for a gas turbine engine
US3014639A (en) High pressure air compressor
US4213296A (en) Seal clearance control system for a gas turbine
US4918923A (en) Internal combustion engine turbosystem and method
US3827236A (en) Cooling systems for turbocharger mechanisms