US2588532A - Jet propulsion unit - Google Patents

Jet propulsion unit Download PDF

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US2588532A
US2588532A US486655A US48665543A US2588532A US 2588532 A US2588532 A US 2588532A US 486655 A US486655 A US 486655A US 48665543 A US48665543 A US 48665543A US 2588532 A US2588532 A US 2588532A
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portion
turbine
means
low pressure
compressing
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US486655A
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Johnson John Algot
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Allis Chalmers Corp
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Allis Chalmers Corp
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/08Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof
    • F02K3/105Heating the by-pass flow
    • F02K3/115Heating the by-pass flow by means of indirect heat exchange

Description

March l952 J. AfJoHNsoN 2,588,532

JET PROPULSION UNIT Filed May 12, 1945 Patented Mar. 11, 1952 JET PROPULSION UNIT JohnAlgot Johnson, Wauwatosa, Wis., assignor to Allis-Chalmers Manufacturing Company, Milwaukee, Wis, a corporation of Delaware Application May 12, 1943, Serial No. 486,655

(Cl. Gil-35.6)

3 Claims,

This invention relates generally to power units and more particularly to jet propulsion apparatus for aircraft and other types of vehicles utilizing a propelling jet or jets formed by separate hot and cold gas streams.

An illustrative dual stream unit of the type hereinabove referred to is disclosed and claimed in my copending application Serial No. 484,159, filed April 23, 1943, now abandoned, and the primary object of this invention is to provide an improved dual or multiple stream jet propulsion unit effective to materially reduce the power requirements of the compressing apparatus employed and to materially increase the effective thrust of the propelling jet or jets.

Another object of this invention is to provide an improved dual or multiple stream jet propulsion unit which in addition to reducing the power requirements and/or increasing the efiective thrust of the propelling jet or jets, makes possible a simplified and durable construction using lighter materials.

In accordance with this invention, the above stated objects may be accomplished in whole or in part by correlating and operatively interconnecting high and low pressure gas compressing apparatus, prime mover apparatus, and nozzle means to provide separate hot and cold gas conducting passages which connect the discharge portions of the high and low pressure compressing apparatus with said nozzle means and which are formed in part by a common heat exchanging wall structure intercoolingly surrounding or otherwise intercoolingly enclosing or partially enclosing the high pressure compressing apparatus in order to increase the temperature and thereby the thrust producing efiectiveness of the cold gas stream, and in order to reduce the power required to drive the compressing apparatus by an amount greater than the increase in heat energy input necessary to raise the temperature of the hot gas stream to the same degree attained without the use of such an intercooling wall structure.

The invention accordingly consists of the various combinations of elements and arrangements of parts as is more fully set forth in the appended claims and in the detailed description, reference being had to the accompanying drawing, in which:

Fig. 1 schematically illustrates in longitudinal section a jet propulsion unit embodying the invention;

Fig. 2 is a transverse section taken on line II-II of Fig. 1; and

Fig. 3 is a transverse section taken On lin III-III of Fig. 1.

Referring to the drawing, it is seen that the invention may be embodied in a unit comprising an axial flow compressor I having a low pressure section 2 and a high pressure section 3, an axial flow turbine 4 axially spaced from and drivingly connected with the adjacent end of the compressor l by means of shaft 6, an inner shell 1 surrounding shaft 6, an intermediate shell 8 surrounding and forming with the active blade carrying portion of the high pressure compressor section 3, with the inner shell 1 and with the active blade carrying portion of the turbine 4 an inner passage 9 including a plurality of fuel distributing nozzles ll disposed therein between the discharge end of the compressor section 3 and the inlet end of the turbine 4, and an outer shell 12 surrounding and forming with the intermediate shell 8 an outer passage I3 receiving a portion of the air discharged from the low pressure compressor section 2; the remainder of the air compressed in the low pressure section 2 entering the high pressure section 3 in which it is further compressed and discharged into the inner passage 9.

The compressor end of the outer shell 12 extends forwardly beyond the low pressure compressor section 2 and forms with a coaxially extending cone-shaped casing l4 enclosing the adjacent shaft portion of the compressor I, a forwardly facing inlet passage l6 through which air enters the active blade carrying portion of the low pressure section 2. The turbine end of the outer shell I2 extends rearwardly beyond the exhaust end of the turbine 4 and forms with a coaxially extending cone-shaped casing I! a thrust producing nozzle structure embodying in this illustration only a single rearwardly directed expanding jet forming passage I8 combiningly received therein the hot gases issuing from the turbine and the relatively cool gas flowing through the outer surrounding passage I3 the active blade carrying portion of the turbine 4 and the exhaust end thereof forming a continuation of and constituting the nozzle end of the inner passage 9.

The compressor end of the intermediate shell 8 terminates in or merges with a heat exchanger comprising a fluted or finned wall portion l9 forming in side-by-side relation (see Fig. 2) longitudinally extending and relatively narrow channel 2] through which flows a portion of the gas issuing from the low pressure portion or section 2; said portion being thus separated from the gas undergoing further compression in the high pressure portion or section-3. In other words, the high and low pressure compressing means, whether constructed and arranged in the partial series flow relation herein shown or otherwise, embodies a common heat exchanging wall structure separating the gas issuing from the low pressure compressing means from that undergoing compression in the high pressure compressing means, thereby increasing the temperature and the thrust producing effectiveness of the relatively cool low pressure stream and simultaneously intercooling the high pressure compressing means sufliciently to reduce the power required to drive said compressing means by an amount greater than the increase in heat energy input necessary to raise the temperature of the hot gas stream to the same degree attained without the use of the heat exchanging wall structure. The nozzle end of the intermediate shell 8 also terminates in or merges with a fluted portion 23 forming in sideby-side relation (see Fig. 3) relatively narrow alternate continuations 24 and 26 of the inner and outer passages 9 and 13, respectively; said continuations functioning to improve the thrust producing effectiveness of the combined streams as is more fully disclosed in applicants aforementioned copending application. However, since this construction forms no part of the present-invention, a further description of same is deemed unnecessary.

The compressor end of the shaft 6 and the adjacent end of the rotorof compressor l connected thereto are supported in a common bearing structure 21 which is in turn supported by a plurality of inwardly extending struts 28 carried by the adjacent portion of the inner shell 1. The forward or inlet end of the rotor of compressor I is supported in journal and thrust bearings 29 and 3|, respectively, which are in turn supported by inwardly extending ribs or the like 32 and 33 carried by the cone-shaped shell 14. The turbine end of the shaft 6 and the ad acent end of the rotor of turbine 4 connected thereto are also supported in a common bearing structure 34 which is in turn supported by inwardly extending struts 36 carried by the adjacent end portion of the inner shell I. sequently, the compressor l, turbine 4 and interconnecting shaft 6 are supported as a unit by three axially spaced bearings with the turbine disposed in overhung relation to the bearing structure 34, thereby presenting a simplified and compact arrangement.

In operation, the air entering the forwardly facing inlet passage I6 is first compressed to a predetermined degree in the low pressure section 2 and issues therefrom in two streams, one of which the inner stream in this case, enters the high pressure section 3 in which it undergoes further compression and issues therefrom into the combustion chamber portion of the passage 9, that is, the portion containing the fuel distributing nozzles II, in which burning takes place, thereby highly heating the inner stream before it enters the turbine 4 from which it is discharged into the nozzle passage 18 and the other of which, the outer stream in this case, enters the outer passage l3 through which it flows directly into the nozzle passage l8 in which it is mixed with the hot gas stream issuing from the turbine 4, the resulting mixture being then expanded in the nozzle passage to produce a rearwardly directed propelling jet or blast having the velocity desired for a given wake efllciency, that is, the efficiency of. the jet or blast as a propelling means.

The feature of separating the relatively 9901 Conlow pressure stream issuing from the compressing section or means 2 from that undergoing further compression in the high pressure compressing section or means 3 by a heat exchanging wall structure effective to increase the temperature and thereby the thrust producing effectiveness of the cool gas stream while intercooling the high pressure compressing means sufficiently to obtain a net gain in energy input results in an increase in both thermal efficiency and propulsive thrust. This improved result is obtained irrespective of whether the cool and hot gas streams are rapidly mixed and then expanded in the manner disclosed in my copending application hereinbefore identified and irrespective of whether the high and low pressure compressing means are constructed and arranged as herein shown and described or otherwise, since the only requirement in this respect is to separate the relatively cool gas stream issuing from a low pressure compressing means from that undergoing compression in a high pressure compressing means by a heat exchanging wall structure which is effective to produce the aforementioned improved results. In this connection, such a separating wall structure can be readily designed by anyone skilled in the art knowing the altitude at which the unit is to be operated, its translational velocity, and whether the temperature and pressure of the jet producing hot gas stream are increased or decreased and to what extent by the prime mover utilizing same, since from this information the temperature of the incoming air, the relative quantities undergoing compression in the high and low pressure compressing means and the ultimate temperatures and velocities of the thrust producing hot and cold gas streams can be determined.

The invention is appl cable generallv to jet propulsion apparatus utilizing a propelling jet or jets formed at least in part by separate streams discharged from high and low pressure compressing means irrespective of whether said streams are separately expanded to produce propellin'g blasts or .iets or mixingly combined and then expanded to produce a single propelling jet or blast and irrespective of whether the prime mover means embodied in such apparatus is of the turbine, internal combustion en ine or other gas utilizing type or combinations thereof; and it should there ore be understood that it is not intended to limit the invention to the exact details of construction herein shown and described, as various modifications within the scope of the appended claims may occur to persons skilled in the art.

It is claimed and desired to secure by Letters Patent:

1. A jet propulsion unit comprising a gas compressing means having a forwardly facing inlet and rearwardly d scharging high and low pressure compressing portions, a turbine spaced from and drivingly connected with said compressing means, a combustion chamber interposed between and connecting the discharge of said high pressure compressing portion with the inlet of said turbine, a rearwardly directed thrust producing nozzle structure, means including said combustion chamber and turbine forming a first fluid confining passage connecting the discharge of said high pressure compressing portion with said nozzle structure, the major extent of the high pressure compressing portion being surrounded by a heat exchanging wall structure separating the gas issuing from the discharge of the low pressure compressing portion from that undergoing further compression in the high pressure compressing portion, said heat exchanging wall structure abstracting heat from the gas undergoing further compression for substantially reducing the power required for driving said gas compressing means and increasing the total heat energy delivered to said nozzle structure, and an outer shell means forming with said heat exchanging wall structure and said turbine a second passage surrounding said combustion chamber and directly connecting the discharge of said low pressure compressing portion with said nozzle structure.

2. A jet propulsion unit comprising a gas compressing means having a forwardly facing inlet and rearwardly discharging high and low pressure compressing portions, a turbine coaxially spaced from and drivingly connected with said compressing means, a combustion chamber interposed between and connecting the discharge of said high pressure compressing portion with the inlet to said turbine, a coaxial, rearwardly directed thrust producing nozzle structure, means including said combustion chamber and turbine forming an inner fluid confining passage connecting the discharge of said high pressure compressing portion with said nozzle structure, and means, including a wall structure providing in side-by-side relation a coaxial series of longitudinally extending channels surrounding the major extent of the high pressure compressing portion and separating a portion of the gas issuing from the discharge of said low pressure compressing portion from that undergoing further compression in the high pressure portion, said wall structure abstracting heat from the gas undergoing further compression for substantially reducing the power required for driving said gas compressing means and thereby increasing the quantity of heat energy carried to said nozzle structure by the exhaust from said turbine and by the low pressure gas forming an outer surrounding passage directly connecting the discharge of said low pressure compressing portion with said nozzle structure.

3. A jet propulsion unit comprising in coaxial combination a gas compressing means having a forwardly facing inlet and rearwardly discharging high and low pressure compressing portions, a turbine spaced from and drivingly connected 6 with said compressing means, a combustion chamber defining means interposed between and directly connecting the discharge of said high pressure compressing portion with the inlet to said turbine, a rearwardly directed thrust producing nozzle structure, means including said combustion chamberdefining means and turbine forming an inner fluid confining passage connecting the discharge of said high pressure compressing portion with said nozzle structure, the major extent of the high pressure compressing portion being surrounded by a radially finned wall structure separating the gas issuing from the discharge of the low pressure compressing portion from that undergoing further compression in the high pressure portion, said finned wall abstracting heat from the gas undergoing further compression for substantially reducing the power required for driving said gas compressing means whereby the unused power is delivered to the nozzle structure as an increased quantity of heat energy, and an outer shell means forming with said finned wall structure, said combustion chamber defining means and said turbine a surrounding passage directly connecting the discharge of said low pressure compressing portion with said nozzle structure.

JOHN ALGOT JOHNSON.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS Number Name Date 2,396,911 Anxionnaz et a1. 'Mar. 19, 1946 2,397,816 Sorensen Apr. 2, 1946 FOREIGN PATENTS Number Country Date 818,703 France June 21, 1937 216,157 Great Britain Nov. 6, 1924 538,022 Great Britain July 17, 1941 OTHER REFERENCES Serial No. 367,666, Anxionnaz et al. (A. P. 0.), published May 25, 1943.

Serial No. 367,677, Anxionnaz et al. (A. P. 0.), published April 27, 1943.

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Cited By (31)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2704440A (en) * 1952-01-17 1955-03-22 Power Jets Res & Dev Ltd Gas turbine plant
US2709337A (en) * 1952-03-28 1955-05-31 United Aircraft Corp Boundary layer control for the diffuser of a gas turbine
US2780060A (en) * 1951-02-14 1957-02-05 Rolls Royce Combustion equipment and nozzle guide vane assembly with cooling of the nozzle guide vanes
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
US2796732A (en) * 1953-05-12 1957-06-25 Napier & Son Ltd Gas producing apparatus of the gas turbine type
US2834181A (en) * 1950-10-07 1958-05-13 Snecma Jet propulsion unit comprising pulse jet units having ejector tubes within a ramjet unit
US2882992A (en) * 1957-04-15 1959-04-21 United Aircraft Corp Low-drag exhaust silencer
DE1083086B (en) * 1956-10-13 1960-06-09 George Simpson Ledgerwood Jet engine
US2940692A (en) * 1953-01-13 1960-06-14 Rolls Royce Aircraft structures with power plants
US2978865A (en) * 1956-02-06 1961-04-11 Curtiss Wright Corp Turbo fan exhaust mixing device
US2982495A (en) * 1951-08-16 1961-05-02 Rolls Royce Aircraft with tiltable lift engines
US2999672A (en) * 1958-04-09 1961-09-12 Curtiss Wright Corp Fluid mixing apparatus
US3009319A (en) * 1955-06-29 1961-11-21 Gregory D Filipenco Turbojet engine
DE1118538B (en) * 1957-04-03 1961-11-30 Rolls Royce Shunt gas turbine jet engine
US3016698A (en) * 1959-07-02 1962-01-16 Gen Motors Corp Bypass engine
US3043101A (en) * 1959-03-13 1962-07-10 Roils Royce Ltd By-pass gas turbine engine employing reheat combustion
US3048376A (en) * 1958-04-09 1962-08-07 Curtiss Wright Corp Fluid mixing apparatus
US3060680A (en) * 1957-12-30 1962-10-30 Rolls Royce By-pass gas-turbine engine and control therefor
DE1148817B (en) * 1959-06-23 1963-05-16 Rolls Royce Two current gas turbine jet engine
US3122886A (en) * 1958-09-02 1964-03-03 Davidovitch Vlastimir Gas turbine cycle improvement
DE1184562B (en) * 1960-08-02 1964-12-31 Rolls Royce Two current gas turbine jet engine
US3196608A (en) * 1959-06-23 1965-07-27 Rolls Royce Apparatus to admix by-pass air with exhaust gases in a by-pass gas-turbine engine
US3269114A (en) * 1966-08-30 Marchant etal j et propulsion engines
US3273340A (en) * 1963-11-22 1966-09-20 Gen Electric Gas turbine powerplant having an extremely high pressure ratio cycle
US3282052A (en) * 1964-04-08 1966-11-01 Lagelbauer Ernest Bypass ramjet engine with heat exchanger
US3289413A (en) * 1964-08-19 1966-12-06 Gen Electric Fluid mixing apparatus for turbofan engines
US3540216A (en) * 1967-01-23 1970-11-17 Snecma Two-flow gas turbine jet engine
US3871174A (en) * 1972-04-25 1975-03-18 Snecma Jet engines
US5706651A (en) * 1995-08-29 1998-01-13 Burbank Aeronautical Corporation Ii Turbofan engine with reduced noise
US5761900A (en) * 1995-10-11 1998-06-09 Stage Iii Technologies, L.C. Two-stage mixer ejector suppressor
DE102008024335A1 (en) * 2008-05-20 2009-11-26 Frank Schuster Gas-turbine engine for use as airplane engine, has cooling fins attached as series intermediate cooler on outer casing of high pressure compressor for cooling main air stream without removing main air stream from flow path

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB216157A (en) * 1923-05-17 1924-11-06 Marcel Pretot Arrangement for cooling air-compressors and improving the output therefrom
FR818703A (en) * 1936-03-04 1937-10-02 Improvements to devices for aircraft propulsion
GB538022A (en) * 1940-01-15 1941-07-17 Gustav Eichelberg Improvements in the propulsion of aircraft
US2396911A (en) * 1939-12-04 1946-03-19 Anxionnaz Rene Reaction propelling device for aircraft
US2397816A (en) * 1942-02-09 1946-04-02 Ford Motor Co Exhaust turbosupercharger

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB216157A (en) * 1923-05-17 1924-11-06 Marcel Pretot Arrangement for cooling air-compressors and improving the output therefrom
FR818703A (en) * 1936-03-04 1937-10-02 Improvements to devices for aircraft propulsion
US2396911A (en) * 1939-12-04 1946-03-19 Anxionnaz Rene Reaction propelling device for aircraft
GB538022A (en) * 1940-01-15 1941-07-17 Gustav Eichelberg Improvements in the propulsion of aircraft
US2397816A (en) * 1942-02-09 1946-04-02 Ford Motor Co Exhaust turbosupercharger

Cited By (33)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3269114A (en) * 1966-08-30 Marchant etal j et propulsion engines
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
US2834181A (en) * 1950-10-07 1958-05-13 Snecma Jet propulsion unit comprising pulse jet units having ejector tubes within a ramjet unit
US2780060A (en) * 1951-02-14 1957-02-05 Rolls Royce Combustion equipment and nozzle guide vane assembly with cooling of the nozzle guide vanes
US2982495A (en) * 1951-08-16 1961-05-02 Rolls Royce Aircraft with tiltable lift engines
US2704440A (en) * 1952-01-17 1955-03-22 Power Jets Res & Dev Ltd Gas turbine plant
US2709337A (en) * 1952-03-28 1955-05-31 United Aircraft Corp Boundary layer control for the diffuser of a gas turbine
US2940692A (en) * 1953-01-13 1960-06-14 Rolls Royce Aircraft structures with power plants
US2796732A (en) * 1953-05-12 1957-06-25 Napier & Son Ltd Gas producing apparatus of the gas turbine type
US3009319A (en) * 1955-06-29 1961-11-21 Gregory D Filipenco Turbojet engine
US2978865A (en) * 1956-02-06 1961-04-11 Curtiss Wright Corp Turbo fan exhaust mixing device
DE1083086B (en) * 1956-10-13 1960-06-09 George Simpson Ledgerwood Jet engine
DE1118538B (en) * 1957-04-03 1961-11-30 Rolls Royce Shunt gas turbine jet engine
US3100627A (en) * 1957-04-03 1963-08-13 Rolls Royce By-pass gas-turbine engine
US2882992A (en) * 1957-04-15 1959-04-21 United Aircraft Corp Low-drag exhaust silencer
US3060680A (en) * 1957-12-30 1962-10-30 Rolls Royce By-pass gas-turbine engine and control therefor
US2999672A (en) * 1958-04-09 1961-09-12 Curtiss Wright Corp Fluid mixing apparatus
US3048376A (en) * 1958-04-09 1962-08-07 Curtiss Wright Corp Fluid mixing apparatus
US3122886A (en) * 1958-09-02 1964-03-03 Davidovitch Vlastimir Gas turbine cycle improvement
US3043101A (en) * 1959-03-13 1962-07-10 Roils Royce Ltd By-pass gas turbine engine employing reheat combustion
DE1148817B (en) * 1959-06-23 1963-05-16 Rolls Royce Two current gas turbine jet engine
US3196608A (en) * 1959-06-23 1965-07-27 Rolls Royce Apparatus to admix by-pass air with exhaust gases in a by-pass gas-turbine engine
US3016698A (en) * 1959-07-02 1962-01-16 Gen Motors Corp Bypass engine
DE1184562B (en) * 1960-08-02 1964-12-31 Rolls Royce Two current gas turbine jet engine
US3273340A (en) * 1963-11-22 1966-09-20 Gen Electric Gas turbine powerplant having an extremely high pressure ratio cycle
US3282052A (en) * 1964-04-08 1966-11-01 Lagelbauer Ernest Bypass ramjet engine with heat exchanger
US3289413A (en) * 1964-08-19 1966-12-06 Gen Electric Fluid mixing apparatus for turbofan engines
US3540216A (en) * 1967-01-23 1970-11-17 Snecma Two-flow gas turbine jet engine
US3871174A (en) * 1972-04-25 1975-03-18 Snecma Jet engines
US5706651A (en) * 1995-08-29 1998-01-13 Burbank Aeronautical Corporation Ii Turbofan engine with reduced noise
US5943856A (en) * 1995-08-29 1999-08-31 Burbank Aeronautical Corporation Ii Turbofan engine with reduced noise
US5761900A (en) * 1995-10-11 1998-06-09 Stage Iii Technologies, L.C. Two-stage mixer ejector suppressor
DE102008024335A1 (en) * 2008-05-20 2009-11-26 Frank Schuster Gas-turbine engine for use as airplane engine, has cooling fins attached as series intermediate cooler on outer casing of high pressure compressor for cooling main air stream without removing main air stream from flow path

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