US3172260A - Gas turbine engine - Google Patents

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US3172260A
US3172260A US151931A US15193161A US3172260A US 3172260 A US3172260 A US 3172260A US 151931 A US151931 A US 151931A US 15193161 A US15193161 A US 15193161A US 3172260 A US3172260 A US 3172260A
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shaft
turbine
unit
air
wheel
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Chute Richard
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Continental Motors Corp
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/08—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage
    • F02C3/085—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage the turbine being of the radial-flow type (radial-radial)
    • 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

Definitions

  • This invention relates to gas turbine engines and more particularly to a compact, dual stage turbine unit utilizing back to back radial flow compressors and turbine wheels.
  • the present emphasis on turbine engine development has produced a need for a compact, efiicient turbine power plant.
  • duct work has presented one of the biggest obstacles in designing such a unit.
  • the present invention minimizes the ductwork necessary in a dual stage turbine engine by a unique arrangement of necessary components.
  • a further object of the invention is to produce a compact dual-staged turbine system having efiicient and smooth gas flow characteristics.
  • FIG. 1 is a longitudinal cross sectional view of a preferred embodiment of the present invention.
  • FIG. 2 is a cross sectional view similar to FIG. 1 but showing another preferred embodiment of the present invention.
  • FIG. 3 is a cross-sectional view taken substantially on the line 33 of FIG. 2.
  • FIG. 1 shows a gas turbine engine 10 as comprising a turbine unit 11 and a compressor unit 12 carried at opposite ends of a shaft 13.
  • a burner unit 14 is disposed intermediate the compressor unit 12 and the turbine unit 11.
  • the compressor unit 12 comprises a compressor housing 15 rotatably carrying the shaft 13 by means of bearings 13A carried at the inner end of the compressor housing 15.
  • the compressor housing 15 is provided with an air intake port 17. Air enters the port 17 as indicated by the arrows beginning at A in FIG. 1, and is directed by stator vanes 18 into the radial flow impeller vanes 19 carried on a wheel 20 secured to the end of the shaft 13. The air as compressed leaves the tips of the vanes 19 and flows past diffuser vanes 19A into a toroidal chamber 21 provided in the compressor housing 15 and which is substantially concentric to the radial flow wheel 20.
  • the air passes arcuately through the chamber 21 and is directed by a second set of stator vanes 23 into the path of radial flow impeller vanes 24 of a second radial flow wheel 25.
  • the radial flow wheel 25 is constructed substantially the same as the wheel 20 and is secured to the shaft 13 in back to back relation with the wheel 20.
  • the air at increased pressure, flows into a central toroidal chamber 22.
  • the compressed air passes through a compressed multiple air outlet port 29 provided in the compressor housing 15 and into air duct 30 to the burner unit 14.
  • the portions of the compressor housing 15 which form the toroidal chambers 21 and 22 are preferably strengthened and supported by annularly spaced through bolts 27 or similar means. Labyrinth seals 28 prevent leakage of the compressed air along the rotating shaft 13.
  • the burner unit 14 comprises a burner housing 31 enclosing a compressed air chamber 32 which in turn encompasses a burner liner 33 defining a combustion chamber 34.
  • the burner housing 31 is provided with a compressed air inlet port 35 communicating with the air duct 30.
  • the compressed air passes through the inlet port 35 into the compressed air chamber 32 and through ports 36 provided in the burner liner 33 into the combustion chamber 34.
  • a fuel nozzle 37 sprays fuel into the chamber 34 and an igniter plug 38 provides ignition to the fuel-air mixture.
  • the combustion gases are directed by a gas duct 39 to the turbine unit 11.
  • the turbine unit 11 comprises a turbine housing 40 rotatably carrying the shaft 13 by means of bearings 41 at an end of the shaft opposite the position of the compressor housing 15.
  • the turbine housing 40 defines a toroidal chamber 42 which is like the chamber 21.
  • the housing 40 is provided with an exhaust gas inlet 45 which provides communication between the exhaust gas duct 39 and a central toroidal chamber 44.
  • the exhaust gases are directed from the chamber 44 radially inwardly into the path of a radial inflow turbine wheel 46 secured to the shaft 13. Rotative force applied to the wheel 46 by the gases is utilized to turn the shaft 13 and thus rotate the compressor wheels 20 and 25.
  • the turbine unit 11 is also provided with labyrinth seals 51 to prevent the escape of gases along the axis of the shaft 13.
  • the housing 40 is provided with annularly spaced through bolts 52 to add strength and to support the portions of the housing 40 encompassing the toroidal chambers 42 and 44.
  • FIGS. 1 and 2 illustrate possible applications of the gas turbine engine of the present invention.
  • the shaft 13 is geared as at 13B to provide a power take off which may have many different uses.
  • a power takeoff could be geared to the transmission of a vehicle or it could be used to provide power to propel marine craft or aircraft.
  • FIG. 2 illustrates a generator 53 as having an armature 54 being rotated by the shaft 13. It will be noted that this embodiment of the invention provides a very compact generator unit.
  • the present invention provides a method of combining two radial flow compressors and two radial flow turbines, producing a highly compact and efficient gas turbine engine.
  • the reduction in the amount of ductwork required as well as the space saved in positioning the compressor wheels and turbine wheels back to back has produced a unit which provides a very high horsepower for the amount of space required and its total weight.
  • these features help to reduce manufacturing and shipping costs.
  • a gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft,
  • said compressor unit comprising,
  • first and second compressor wheels arranged in back-to-back relation with their respective vanes spaced along said passage
  • first and second turbine wheels carried on said shaft and having their respective vanes spaced along said last passage, said first turbine wheel receiving exhaust gases from said last inlet and delivering it to said second turbine wheel, and
  • a gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft,
  • a gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft, (b) a compressor unit carried at one end of said shaft,
  • said compressor unit comprising a toroidal air passage of spiral cross section extending coaxially about said shaft
  • first and second axial-inflow radial-outflow compressor wheels arranged in back-to-back relation with their respective vanes spaced along said passage
  • first and second radial-inflow axial-outflow turbine wheels carried on said shaft and having their respective vanes spaced along said last passage, said first turbine wheel receiving exhaust gases radially from said last inlet and delivering it radially to said second turbine wheel, and
  • a gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft,
  • said turbine unit comprising a toroidal passage of spiral cross-section extending coaxially about said shaft, an exhaust gas inlet at the inner end of said spiral receiving gases from said communicating means,
  • first and second radial-inflow axial-outflow turbine wheels carried on said shaft and having their respective vanes spaced along said last passage, said first turbine wheel receiving exhaust gases radially from said last inlet and delivering it radially to said second turbine Wheel, and
  • each of the turbine wheels having vanes respectively spaced in said passage

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Description

March 9, 1965 R. CHUTE 3,172,260
GAS TURBINE ENGINE Filed Nov. 15, 1961 2 Sheets-Sheet l INVENTOR. I?! am RD CHUTE ATTO RNEYS March 9, 1965 R. CHUTE 3,172,260
GAS TURBINE ENGINE Filed Nov. 13. 1961 2 Sheets-Sheet 2 mmvroa. P/c'HA/w CHUTE ATTO RN EYS United States Patent 3,172,260 GAS TURBINE ENGINE Richard Chute, Huntington Woods, Mich, assignor to Continental Motors Corporation, Muskegon, Mich., a corporation of Virginia Filed Nov. 13, 1961, Ser. No. 151,931 Claims. (Cl. oil-39.75)
This invention relates to gas turbine engines and more particularly to a compact, dual stage turbine unit utilizing back to back radial flow compressors and turbine wheels.
The present emphasis on turbine engine development has produced a need for a compact, efiicient turbine power plant. Heretofore, duct work has presented one of the biggest obstacles in designing such a unit. The present invention minimizes the ductwork necessary in a dual stage turbine engine by a unique arrangement of necessary components.
It is an object of the present invention to minimize the space requirements of a dual stage gas turbine engine by arranging the necessary components so as to produce a compact unit.
It is a further object of the present invention to reduce manufacturing costs of a turbine engine by providing a simply constructed, low loss, two stage unit.
A further object of the invention is to produce a compact dual-staged turbine system having efiicient and smooth gas flow characteristics.
Further objects and advantages will be apparent to one skilled in the art to which the invention pertains upon reference to the following drawings illustrating preferred embodiments of the invention in which like characters refer to like parts throughout the several views and in which:
FIG. 1 is a longitudinal cross sectional view of a preferred embodiment of the present invention.
FIG. 2 is a cross sectional view similar to FIG. 1 but showing another preferred embodiment of the present invention, and
FIG. 3 is a cross-sectional view taken substantially on the line 33 of FIG. 2.
Referring now to the drawings for a more detailed description of the present invention, FIG. 1 shows a gas turbine engine 10 as comprising a turbine unit 11 and a compressor unit 12 carried at opposite ends of a shaft 13. A burner unit 14 is disposed intermediate the compressor unit 12 and the turbine unit 11.
The compressor unit 12 comprises a compressor housing 15 rotatably carrying the shaft 13 by means of bearings 13A carried at the inner end of the compressor housing 15. The compressor housing 15 is provided with an air intake port 17. Air enters the port 17 as indicated by the arrows beginning at A in FIG. 1, and is directed by stator vanes 18 into the radial flow impeller vanes 19 carried on a wheel 20 secured to the end of the shaft 13. The air as compressed leaves the tips of the vanes 19 and flows past diffuser vanes 19A into a toroidal chamber 21 provided in the compressor housing 15 and which is substantially concentric to the radial flow wheel 20.
The air passes arcuately through the chamber 21 and is directed by a second set of stator vanes 23 into the path of radial flow impeller vanes 24 of a second radial flow wheel 25. The radial flow wheel 25 is constructed substantially the same as the wheel 20 and is secured to the shaft 13 in back to back relation with the wheel 20. The air, at increased pressure, flows into a central toroidal chamber 22.
The compressed air passes through a compressed multiple air outlet port 29 provided in the compressor housing 15 and into air duct 30 to the burner unit 14.
The portions of the compressor housing 15 which form the toroidal chambers 21 and 22 are preferably strengthened and supported by annularly spaced through bolts 27 or similar means. Labyrinth seals 28 prevent leakage of the compressed air along the rotating shaft 13.
The burner unit 14 comprises a burner housing 31 enclosing a compressed air chamber 32 which in turn encompasses a burner liner 33 defining a combustion chamber 34. The burner housing 31 is provided with a compressed air inlet port 35 communicating with the air duct 30. The compressed air passes through the inlet port 35 into the compressed air chamber 32 and through ports 36 provided in the burner liner 33 into the combustion chamber 34. A fuel nozzle 37 sprays fuel into the chamber 34 and an igniter plug 38 provides ignition to the fuel-air mixture. The combustion gases are directed by a gas duct 39 to the turbine unit 11.
The turbine unit 11 comprises a turbine housing 40 rotatably carrying the shaft 13 by means of bearings 41 at an end of the shaft opposite the position of the compressor housing 15. The turbine housing 40 defines a toroidal chamber 42 which is like the chamber 21. The housing 40 is provided with an exhaust gas inlet 45 which provides communication between the exhaust gas duct 39 and a central toroidal chamber 44. The exhaust gases are directed from the chamber 44 radially inwardly into the path of a radial inflow turbine wheel 46 secured to the shaft 13. Rotative force applied to the wheel 46 by the gases is utilized to turn the shaft 13 and thus rotate the compressor wheels 20 and 25.
Since energy still remains in the exhaust gases, they are directed by a set of stator vanes 47 into the toroidal chamber 42 and from there radially inwardly into the path of a second radial inflow turbine wheel 48. The turbine wheels 46 and 48, like their counterparts compressor wheels 20 and 25, are secured to the shaft 13 in a back to back relationship. The exhaust gases are now directed by a second set of stator vanes 49 through an exhaust outlet port 50 provided in the turbine housing 40.
Like the compressor unit 12, the turbine unit 11 is also provided with labyrinth seals 51 to prevent the escape of gases along the axis of the shaft 13. Also the housing 40 is provided with annularly spaced through bolts 52 to add strength and to support the portions of the housing 40 encompassing the toroidal chambers 42 and 44.
FIGS. 1 and 2 illustrate possible applications of the gas turbine engine of the present invention. In FIG. 1 the shaft 13 is geared as at 13B to provide a power take off which may have many different uses. Such a power takeoff could be geared to the transmission of a vehicle or it could be used to provide power to propel marine craft or aircraft. FIG. 2 illustrates a generator 53 as having an armature 54 being rotated by the shaft 13. It will be noted that this embodiment of the invention provides a very compact generator unit.
It is apparent from the foregoing description that the present invention provides a method of combining two radial flow compressors and two radial flow turbines, producing a highly compact and efficient gas turbine engine. The reduction in the amount of ductwork required as well as the space saved in positioning the compressor wheels and turbine wheels back to back has produced a unit which provides a very high horsepower for the amount of space required and its total weight. In addition to permitting use of gas turbine engines in situations where heretofore weight or space limitations were prohibitive, these features help to reduce manufacturing and shipping costs.
Although I have described but a few embodiments of the invention, it will be apparent to one skilled in vthe art to which the invention pertains that various .2 changes and modifications may be made therein without departing from the spirit of the invention as expressed by the scope of the appended claims.
I claim:
1. A gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft,
(b) a compressor unit carried at one end of said shaft,
(c) a turbine unit carried at the other end of said shaft,
(d) means providing communication between said compressor unit and said turbine unit,
(2) a burner disposed in said communicating means intermediate said compressor unit and said turbine unit,
(7") said compressor unit comprising,
a toroidal air passage of spiral cross-section extending coaxially about said shaft,
first and second compressor wheels arranged in back-to-back relation with their respective vanes spaced along said passage,
an air inlet at the outer end of said spiral encompassing said shaft and delivering air to said first wheel, said first wheel then delivering the air to said second wheel, and
an air outlet at the inner end of said spiral receiving air from said second wheel and delivering it to said communicating means.
2. The gas turbine engine as defined in claim 1 and in which said turbine unit comprises (a) a toroidal passage of spiral cross-section extending coaxially about said shaft,
(b) an exhaust gas inlet at the inner end of said spiral receiving gases from said communicating means,
() first and second turbine wheels carried on said shaft and having their respective vanes spaced along said last passage, said first turbine wheel receiving exhaust gases from said last inlet and delivering it to said second turbine wheel, and
(d) an exhaust gas outlet at the outer end of said spiral receiving exhaust gases from said second turbine wheel.
3. The engine as defined in claim 1 and in which said shaft is further provided with power takeoff means disposed intermediate said compressor and said turbine unit.
4. A gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft,
(b) a compressor unit carried at one end of said shaft,
(0) a turbine unit carried at the other end of said shaft,
(d) means providing communication between said compressor unit and said turbine unit,
(e) a burner disposed in said communicating means' intermediate said compressor unit and said turbine unit, (f) said turbine unit comprising a toroidal passage of spiral cross-section extending coaxially about said shaft, an exhaust gas inlet at the inner end of said spiral receiving gases from said communicating means, first and second turbine Wheels carried on said shaft and having their respective vanes spaced along said last passage, said first turbine wheel receiving exhaust gases from said last inlet and delivering it to said second turbine wheel, an exhaust gas outlet at the outer end of said spiral receiving exhaust gases from said second turbine wheel. 5. A gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft, (b) a compressor unit carried at one end of said shaft,
4 (c) a turbine unit carried at the other end of said shaft, (d) means providing communication between said compressor unit and said turbine unit,
5 (e) a burner disposed in said communicating means intermediate said compressor unit and said turbine unit,
(f) said compressor unit comprising a toroidal air passage of spiral cross section extending coaxially about said shaft,
first and second axial-inflow radial-outflow compressor wheels arranged in back-to-back relation with their respective vanes spaced along said passage,
an air inlet at the outer end of said spiral encompassing said shaft and delivering air axially to said first wheel, said first wheel then delivering the air axially to said second wheel, and
an air outlet at the inner end of said spiral receiving air radially from said said second wheel and delivering it to said communicating means.
6. The gas turbine engine as defined in claim 5 and including power takeoff means fixed to said shaft intermediate said compressor unit and said turbine unit.
7. The gas turbine engine as defined in claim 5 and in which said turbine unit comprises (a) a toroidal passage of spiral cross-section extending coaxially about said shaft, (b) an exhaust gas inlet at the inner end of said spiral receiving gases from said communicating means,
(0) first and second radial-inflow axial-outflow turbine wheels carried on said shaft and having their respective vanes spaced along said last passage, said first turbine wheel receiving exhaust gases radially from said last inlet and delivering it radially to said second turbine wheel, and
(d) an exhaust gas outlet at the outer end of said spiral receiving exhaust gases axially from said second turbine wheel.
8. A gas turbine engine comprising (a) a shaft and means rotatably supporting said shaft,
(b) a compressor unit carried at one end of said shaft,
H (c) a turbine unit carried at the other end of said shaft,
' (d) means providing communication between said compressor unit and said turbine unit,
(e) a burner disposed in said communicating means intermediate said compressor unit and said turbine unit,
(f) said turbine unit comprising a toroidal passage of spiral cross-section extending coaxially about said shaft, an exhaust gas inlet at the inner end of said spiral receiving gases from said communicating means,
first and second radial-inflow axial-outflow turbine wheels carried on said shaft and having their respective vanes spaced along said last passage, said first turbine wheel receiving exhaust gases radially from said last inlet and delivering it radially to said second turbine Wheel, and
an exhaust gas outlet at the outer end of said spiral receiving exhaust gases axially from saidsecond turbine wheel. 9. In a gas turbine engine including a two-stage compressor, an air inlet for said compressor, and an outlet delivering air from said compressor to a burner, the improvement comprising:
(a) a toroidal air passage of spiral cross-section in said compressor,
(b) each of the compressor wheels having vanes re- 5 spectively spaced in said passage,
(c) said air inlet being disposed at one end of said spiralling passage, and
(d) said air outlet being disposed at the opposite end of said spiralling passage.
10. In a gass turbine engine including a two-stage tur- 5 blue, a gas-exhaust inlet for said turbine, and a gasexhaust outlet for said turbine, the improvement comprising:
(a) a toroidal passage of spiral cross-section in said turbine, 10
(b) each of the turbine wheels having vanes respectively spaced in said passage,
(0) said inlet being disposed at one end of said spiralling passage, and
(d) said outlet being disposed at the opposite end of 15 said spiralling passage.
References Cited in the file of this patent UNITED STATES PATENTS Birmann June 22, 1948 Birmann June 14, 1949 einhardt Ian. 16, 1951 Pavlecka Oct. 15, 1957 Hill Jan. 28, 1958 Sampietro July 18, 1961 FOREIGN PATENTS Germany Feb. 23, 1923 Great Britain Dec. 10, 1948

Claims (1)

1. A GAS TURBINE ENGINE COMPRISING (A) A SHAFT AND MEANS ROTATABLY SUPPORTING SAID SHAFT, (B) A COMPRESSOR UNIT CARRIED AT ONE END OF SAID SHAFT, (C) A TURBINE UNIT CARRIED AT THE OTHER END OF SAID SHAFT, (D) MEANS PROVIDING COMMUNICATION BETWEEN SAID COMPRESSOR UNIT AND SAID TURBINE UNIT, (E) A BURNER DISPOSED IN SAID COMMUNICATING MEANS INTERMEDIATE SAID COMPRESSOR UNIT AND SAID TURBINE UNIT, (F) SAID COMPRESSOR UNIT COMPRISING, A TOROIDAL AIR PASSAGE OF SPIRAL CROSS-SECTION EXTENDING COAXIALLY ABOUT SAID SHAFT, FIRST AND SECOND COMPRESSOR WHEELS ARRANGED IN BACK-TO-BACK RELATION WITH THEIR RESPECTIVE VANES SPACED ALONG SAID PASSAGE, AN AIR INLE AT THE OUTER END OF SAID SPIRAL ENCOMPASSING SAID SHAFT AND DELIVERING AIR TO SAID FIRST WHEEL, SAID FIRST WHEEL THEN DELIVERING THE AIR TO SAID SECOND WHEEL, AND AN AIR OUTLET AT THE INNER END OF SAID SPIRAL RECEIVING AIR FROM SAID SECOND WHEEL AND DELIVERING IT TO SAID COMMUNICATING MEANS.
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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3761205A (en) * 1972-03-20 1973-09-25 Avco Corp Easily maintainable gas turbine engine
US4640153A (en) * 1983-12-09 1987-02-03 Teledyne Industries, Inc. Accessory drive for a turbine engine
US11415046B1 (en) * 2019-06-04 2022-08-16 United States Of America As Represented By The Secretary Of The Air Force Disk engine with circumferential swirl radial combustor

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE369837C (en) * 1921-01-06 1923-02-23 Georg Stauber Dr Ing Rotary compressor
US2443717A (en) * 1942-05-02 1948-06-22 Turbo Engineering Corp Exhaust gas and hot air turbine system
GB614160A (en) * 1945-03-05 1948-12-10 Power Jets Res & Dev Ltd Improvements relating to combustion turbine power plant
US2473356A (en) * 1942-04-18 1949-06-14 Turbo Engineering Corp Combustion gas turbine arrangement
US2538179A (en) * 1945-09-04 1951-01-16 Grace K Weinhardt Rotary power generator
US2809493A (en) * 1951-03-19 1957-10-15 American Mach & Foundry Centrifugal flow compressor and gas turbine power plant with a centrifugal flow compressor, toroidal combustion chamber, and centripetal flow turbine
US2821067A (en) * 1956-05-28 1958-01-28 Boeing Co Combustion chamber construction in a gas turbine engine
US2992529A (en) * 1956-08-23 1961-07-18 Thompson Ramo Wooldridge Inc Turbine blade cooling

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE369837C (en) * 1921-01-06 1923-02-23 Georg Stauber Dr Ing Rotary compressor
US2473356A (en) * 1942-04-18 1949-06-14 Turbo Engineering Corp Combustion gas turbine arrangement
US2443717A (en) * 1942-05-02 1948-06-22 Turbo Engineering Corp Exhaust gas and hot air turbine system
GB614160A (en) * 1945-03-05 1948-12-10 Power Jets Res & Dev Ltd Improvements relating to combustion turbine power plant
US2538179A (en) * 1945-09-04 1951-01-16 Grace K Weinhardt Rotary power generator
US2809493A (en) * 1951-03-19 1957-10-15 American Mach & Foundry Centrifugal flow compressor and gas turbine power plant with a centrifugal flow compressor, toroidal combustion chamber, and centripetal flow turbine
US2821067A (en) * 1956-05-28 1958-01-28 Boeing Co Combustion chamber construction in a gas turbine engine
US2992529A (en) * 1956-08-23 1961-07-18 Thompson Ramo Wooldridge Inc Turbine blade cooling

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3761205A (en) * 1972-03-20 1973-09-25 Avco Corp Easily maintainable gas turbine engine
US4640153A (en) * 1983-12-09 1987-02-03 Teledyne Industries, Inc. Accessory drive for a turbine engine
US11415046B1 (en) * 2019-06-04 2022-08-16 United States Of America As Represented By The Secretary Of The Air Force Disk engine with circumferential swirl radial combustor
US12018613B1 (en) * 2019-06-04 2024-06-25 United States Of America As Represented By The Secretary Of The Air Force Disk engine with circumferential swirl radial combustor

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