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US3709629A - Integrated flow gas turbine - Google Patents

Integrated flow gas turbine Download PDF

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US3709629A
US3709629A US3709629DA US3709629A US 3709629 A US3709629 A US 3709629A US 3709629D A US3709629D A US 3709629DA US 3709629 A US3709629 A US 3709629A
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blades
combustion
gases
rotor
hot
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E Traut
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E Traut
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    • 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
    • F02C3/00Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/045Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor having compressor and turbine passages in a single rotor-module

Abstract

A gas turbine having a rotor serving as both compressor and turbine, and utilizing a plurality of non-rotating arcuate members disposed in spaced relation about the periphery of the rotor. These arcuate members are involved in the directing of the flow of combustion products into proximity of the blading of the turbine, to cause its rotation, and by virtue of their advantageous design, these arcuate members not only help establish a cool air boundary against which the combustion products react and thus minimize heating of the blades, but also form passages for the subsequent exhausting of the combustion products.

Description

United States Patent 1191 Traut 1451 Jan. 9, 1973 s41 INTEGRATED FLOW GAS TURBINE 2,658,338 11/1953 Leduc ..416/95 Inventor: Earl w. Tram PIO. Box 2,873,945 1/1959 Kuhn ..415/178 Fort Lauderdale, 33307 FOREIGN PATENTS 0R APPLICATIONS Filed: y 1970 941,397 4/1956 Germany ..60/39.43 [21] Appl. No.: 40,633

Primary Examiner-Douglas l-lart Related US. Application Data Assistant Examiner-Warren Olsen [63] Continuation-in-part of Ser. No. 741,623, July 1, Attorney-Julian Renfro 1968, abandoned. ABSTRACT [52] US. Cl AIS/56, 60/3943 A gas turbine having a rotor serving as both compre [51] Int. Cl ..F0ld l/22 and turbine, and utilizing a plurality of non-rotat [58] Field of Search .....60/39.43; 415/54, 58, 56, 57, ing arcuate members disposed in spaced relation 415/175, 177, 199, 178 about the periphery of the rotor. These arcuate members are involved in the directing of the flow of com- [56] References Cited bustion products into proximity of the blading of the turbine, to cause its rotation, and by virtue of their ad- UNITED STATES PATENTS vantageous design, these arcuate members not only 1,349,487 8/1920 Bennett ..415/56 p establish a cool air boundary against which the 1,882,630 10/1932 Jarvis 415/56 combustion products react and thus minimize heating 2,537,344 1/1951 Gruss 60/3943 of the blades, but also form passages for the sub- 3,964 3/1957 Theimer I 9 3 sequent exhausting of the combustion products. 3,310,940 3/1967 Oetliker 60/3943 3,283,509 11/1966 Nitsch ..4l5/143 6 Claims, 12 Drawing Figures PATENTEBJAI 9191s 3; 709,629

SHEET 1 BF 4 INVENTOR EARL WQTRAUT ATTORNEY PATENTED JAN 9 I973 SHEET 2 BF 4 INVENTOR EARL W. TRAUT ATTORNEY PATENTEDJAN 197 3. 709.629

SHEET 3 OF 4 2 INVENTOR- EARL W. TRALJT BY%K@7% ATTORNEY INTEGRATED FLOW GAS TURBINE RELATIONSHIP TO PRIOR APPLICATION This invention is a Continuation-in-Part of my earlier patent application entitled Centripetal Flow Gas Turbine," filed July 1, 1968, Ser. No. 741,623, now'abandoned.

SUMMARY This invention relates to an integrated flow gas turbine, and more particularly to a novel engine in which hot gases created by the combustion of fuel in a combustion chamber are caused to impinge upon blading of a dual purpose rotor arranged to rotate at high speed and to deliver air under pressure to the combustion chamber, my engine serving to deliver a useful amount of shaft power, or as a gas generator.

Most gas turbine engines today are equipped with a separate compressor, often driven from the same shaft as the turbine, for it is necessary to have a considerable amount of air flowing into the engine in order for combustion to take place on a continuous basis in the combustion chamber,

The present invention differs substantially from known prior art engines by utilizing a rotor containing only one set of blades, but with these blades being configured and arranged to perform not only the function of compressing the incoming air and delivering it into the combustion area of the engine, but also the function of receiving the reaction or thrust from the high temperature gases, with the reaction of the gases against the blades serving to perpetuate the rotation thereof. Thus, my engine utilizes well known compression, combustion and reaction cycles in a novel and useful arrangement.

All facets of my invention involve the dual purpose rotor arrangement in which one portion of each blade of the rotor receives the thrust from burning hot gases, such serving to cause further and continued rotation of the blading, and with another portion of each blade of the rotor serving to accomplish the compression of incoming air for the combustion to continue in the intended manner, and for cooling of the blades. However, one embodiment of my invention involves a centripetal flow arrangement in which the rotor is disposed radially outwardly with respect to the combustion chamber, with the relatively cool incoming air flowing centripetally along the blades, and then entering the combustion chamber, with the products from the continuous combustion then flowing outwardly through guide nozzles so as to react against the radially inner portions of the blades, thus to cause the continued rotation of the rotor.

Another embodiment of my invention involves a centrifugal flow arrangement in which the rotor is disposed radially inwardly of the combustion chamber, with air compressed by the rotor flowing centrifugally into the combustion chamber, with the products of combustion thereafter flowing past the radially outer portions of the blades. Still another embodiment involves an axial flow arrangement, with the relatively cool air from the final compressor stage entering the combustion chamber axially, and then flowing in the reverse direction through guide nozzles and reacting against the trailing portions of the final compressor stage blading, thus causing the continued rotation of the rotor, with the configuration in each embodiment being such that uncombusted compressed air separates the blades from the combustion products to such an extent that heating of the blades is minimized, thus permitting the use of much less expensive blades than are required in conventional gas turbines, where no such cooling effect is present.

It is therefore a principal object of my invention to provide an integrated flow gas turbine in which a single rotor is utilized.

It is another object of my invention to provide a gas turbine in which air compressed by a rotor is caused to flow along the blading of the rotor in such a manner as to cool the blading, even in the presence of combustion products.

It is still another object of my invention to provide novel blade configurations for a dual purpose rotor, thus to extract a maximum amount of thrust from the combustion products, and at the same time to derive the maximum cooling for the blades.

It is still another object of my invention to provide a gas turbine in which the products of combustion react against an air boundary that automatically adapts its shape to changes in turbine operating parameters.

These and other objects, features and advantages of my invention will be more apparent from a study of the appended drawings in which:

FIG. 1 is a side elevational view of my integrated flow gas turbine in the centripetal flow embodiment, with some parts in section to reveal internal detail;

FIG. 2 is an end view of the device shown in FIG. 1, also being partly in section;

FIG. 3 is a perspective view of my gas turbine to a smaller scale;

FIG. 4 is a view taken along lines 4-4 in FIG. 1 to reveal gearing;

FIG. 5 is a view taken along lines 55 in FIG. 1 to reveal further blading and nozzle details of the centripetal flow embodiment;

FIG. 5a is an enlarged view ofa portion of FIG. 5, but revealing blading ofa different shape;

FIG. 6 is a fragmentary perspective view of one of the members shown in FIG. 5;

FIG. 7 is a cross-sectional view of a centrifugal flow embodiment of my invention;

FIG. 8 is a view taken along lines 8-8 in FIG. 7;

FIG. 9 is a plan view of an axial flow embodiment of my invention;

FIG. 10 is a cross-sectional view of the axial flow embodiment at approximately the mid portion; and

FIG. 11 is a view taken along lines 11-11 in FIG. 9.

DETAILED DESCRIPTION Turning now to FIG. 1, it will be seen that I have there shown a side elevational view of an exemplary version of my integrated centripetal flow gas turbine I0, with portions of this figure presented in section to reveal internal detail. A rotor 11 is arranged to rotate about a stationary combustion chamber 17, as perhaps best seen in FIGS. 2 and 5. FIG. 3 shows to a reduced scale, the external appearance of the engine.

The rotor 11 comprises a plurality of essentially straight blades 12 arranged in a circular combination, with the blades at one end of the device as seen in FIG. I joined to a support ring 34 that is rotatable on a bearing 35. On the other end, the blades are joined to an internal gear 28 that is in mesh with a plurality of gears 29 mounted on shafts 32, from which power can be delivered for accessories or the like. Engine torque is transmitted to a drive shaft 27 via small gears 30, which are also mounted on the shafts 32. These small gears mesh with large gear 31, which is mounted upon shaft 27. Note FIG. 4.

It will be noted that the blades 12 are essentially arcuate in cross section, having a concave side 14 and a convex side 15, and being spaced essentially equidistant so as to provide space 16 between the blades. It will further be noted that the blades are of comparatively thick construction, with the tips thinner than the roots. As will be seen in greater detail hereinafter, the blades 12 in these figures are thicker at the root location 12a to provide a reaction surface against which gases can react to cause the rotor to turn at a high rate of speed.

Referring principally to FIGS. 2 and 5, it will be noted that the centrally disposed combustion chamber is defined by a plurality of stationary wall components 18 located adjacent the inner periphery of the blades of the rotor 11. As will be seen in FIG. 1, these wall components are of substantially arcuate configuration, being supported by stationary end plates 13, with a non-rotating shaft 37 extending between these end plates to maintain them in the desired relationship. A shroud 38 surrounds the shaft 37 so as to protect it from the heat of the combustion process. Shaft 37 is hollow to permit cooling air to flow through.

FIGS. 2 and reveal that the arcuate wall components 18 are spaced apart so as to form at most locations, guide nozzles 19 which communicate with the combustion chamber 17. It is through these nozzles that exhaust gases flow in order to react against the base of blades 12. Wall component 24 is different from the generally arcuate wall components 18 in that it contains at least one fuel nozzle 21 and igniter 26; see FIG. 6.

Adjacent the wall component 24 is defined an intake duct 20 through which air compressed by the rotation of the blades 12 is caused to flow centripetally so as to enter the combustion chamber; see FIGS. 2 and 5. Fuel nozzle 21 sprays fuel to mix with this incoming; air, with the fuel to air ratio being such that a continuous combustion process can take place in the volume 17 enclosed by the stationary wall components 18. FIG. 1 reveals that more than one fuel nozzle and more than one igniter can be utilized. Each of the wall components 18 is provided with a'concave side 22 facing away from the combustion chamber 17, so as to form a recess 23 between the combustion chamber and the rotor, which recess substantially faces the rotor. The several recesses 23 in effect form exhaust ducts, which in turn connect to the exhaust opening 25 revealed in FIGS. 1 and 2 to be disposed in one of the end plates 13.

Combustion takes place substantially within the central chamber defined by the arcuate members 18, with the combustion products leaving the nozzles 19 at great speed and impinging upon the radially inner ends 120 of the rotor blades 12. During steady state operation, the pressure of the air compressed by the rotation of the blades 12 is only slightly higher than the pressure of the combustion products flowing outwardly through the nozzles 19, but both of these pressures are much higher than the pressure in the recesses 23 and the exhaust ducts. Thus, the hot combustion gases deflect off of the lower surfaces 12a of the blades 12 and are then drawn inwardly with unburned air into the recesses 23 defined in the interior of the stationary wall components 18. Gas in the recesses 23 flows from right to left as viewed in FIG. 1, and then flows outwardly through the exhaust openings 25.

FIG. 5a reveals in general the phenomenon just discussed wherein the hot gases leaving an exhaust nozzle 19 in fact flows around the hooked portion 18a of the member 18 and thence flows into the interior portion 23 of the stationary wall component. It should be noted that unburned gases flowing centripetally between the blades 12 deflect these combustion products and thus prevent the overheating of the blades. A cool air boundary may be regarded as existing between a location adjacent the point 33 of each member 18, and a location slightly radially outwardly of the hooked portion 18a of the adjacent stationary wall component. This hot-cold boundary is identified by a short curved dashed line in FIG. 5a. This boundary will tend to remain in the approximate position just described during steady state operation of my turbine, although it is continually interrupted by the radially inner portions of the blades as they continue to rotate. Engine acceleration increases the pressure of the hot gases flowing through the exhaust nozzles 19 and causes the cool air boundary to bend away from the portion 18a of the stationary wall component until such time as acceleration has ceased, at which time the boundary will be restored to essentially the original position.

As will be discussed hereinafter, FIG. 5a depicts a slightly different blade configuration from that involved in FIGS. 2 and 5, for in FIG. 5a, the blades have a rounded base portion 12b.

I have noted that the unburned compressed air flowing through spaces 16 will increase in pressure as it proceeds radially inward towards base or 12b of each blade, at which location it will expand slightly due to the additional space available. In FIG. 5, after a given blade has moved past a given nozzle 19 to a position essentially adjacent the hook portion 18a, the base 12a of the blade is then reacted upon by the hot combustion gas. The slight expansion of the compressed gas, the change of direction of the combustion gases and the difference in velocities of the burned and unburned gases causes a certain amount of turbulence within the boundary previously described to exist from the point 33 along the space between the hooked portion 18a and the radially inner portion 120 of the nearest blade 12.

It is important to note with regard to FIG. 5 that the only section of the blades reacted upon by the hot combustion gases is the base portion 12a, and significantly, even this portion is intermittently cooled by the air compressed by the blades that travels past the nozzles.

Returning to FIG. 5a, it will be noted that the lower surface 12b of the blades has been shaped, with the entire trailing surface of the blade now being convex, with a smaller radius at the root of the blade than at the opposite edge of the blade. In the configuration in accordance with FIG. 5a, the cool air boundary tends to be defined by the blade roots as the blades pass by the nozzle, and most importantly, the hot gas only approaches the radially inner tip of the blade, with the turbine reaction taking place against the cool air boundary. This arrangement makes it possible for the first time to design a simple air boundary blade of variable shape that will efficiently adapt itself to radical changes in operating parameters, such as absolute pressure, absolute temperature, velocity, accelerations and different fuels. This is of course in contrast with conventional turbines, which must be designed for fixed shape solid blades.

Turning now to FIG. 7, it will be seen that many of the relationships involve in this integrated centrifugal flow gas turbine are quite similar to those described in conjunction with the centripetal flow design discussed in conjunction with FIGS. 1 through 6. In FIG. 7, the rotation of the blades 112 causes a substantial amount of flow of unburned air to enter the air intake ducts 120, flowing in each instance past a nozzle 121 and an igniter 126. Fuel is sprayed into the chambers 117 in the proper ratio in order that an effective combustion process can take place in the combustion chambers. Combustion products then flow at substantial speed radially inwardly through the exhaust nozzles 119 so as to impinge upon the tips of the blades 112.

As will be noted from the upper portion of FIG. 7, there is a substantial amount of mixing taking place between the burned and unburned gases, with the unburned gases serving to protect the tips of the blades 112. Inasmuch as pressure in the recesses 123 defined in the stationary wall components 118 opposite the combustion chamber 117 is less than the pressure of either the burned or the unburned gases, the mixture flows into these recesses and thence in a substantially axial direction to an overboard location.

FIG. 8 reveals a section taken along lines 8-8 in FIG. 7 to reveal the manner in which the exhaust gases leave the engine.

Turning now to FIG. 10, and to related FIG. 9, it will be noted that I have there shown an integrated axial flow gas turbine in accordance with my invention, in which the combustion chamber 217 is defined by stationary end plate 213 and radially oriented stationary wall components 218 adjacent which axial flow rotor 211 is disposed. The blades 212 of this rotor are mounted upon a shaft 227 at spaced locations, with the rotation of this shaft causing air to be delivered into intake ducts 220, which connect into the combustion chamber 217. Several fuel nozzles 22] inject fuel into the combustion chamber 217 so as to achieve the proper fuel to air ratio necessary for desirable combustion. Hot combustion gases leave the combustion chamber through exhaust nozzles 219 so as to react upon the adjacent tips of the blades 212 in the manner shown in FIG. 11. As before, there is a cool air boundary to deflect the combustion products, with the result being that the blades 212 are protected from being overheated. Thereafter, the burned and unburned gases flow outwardly through ducts 223 that connect to exhaust openings 225.

As should now be apparent, there is sufficient reaction of the combustion products against the near side of the blades 212 to cause the rotor 211 to rotate and provide useful power.

As will now be apparent, I have described several embodiments of my invention that I regard as being primary, with different ones of these embodiments being suitable to meet a wide variety of needs.

However, I am not to be limited to the embodiments shown and described herein, for if desired, the compressed air or the hot combustion products could be generated elsewhere and then directed into a machine in accordance with any of the primary embodiments of this invention, or compressed air from a separate source could be used for blade cooling instead of being generated by the blades of this invention.

As a further point, a portion of the blading can be utilized only for compression, involving a structural modification different than the foregoing, and involving use of a flow divider such that part of the air compressed by the rotor is delivered for combustion into an adjacent surrounding combustion chamber, whereas the remainder of the flow from the compressor is utilized only for cooling this other portion of the blading.

As a further point, one edge of a blade can be used for generating pressure and an essentially perpendicular edge can be used for obtaining thrust and yet be self-cooling. For instance, the outward tips of a centrifugal compressor can be used for conventional generation of pressure, while portions of the radial edges can be beveled, shaped, or otherwise configured and used as a turbine reaction surface.

As a further point, almost any conventional or other turbine blade can be utilized for obtaining thrust and yet be self-cooling by utilizing the previously described cool air boundary for hot gas reaction.

Iclaim:

1. A turbine adapted to be operated by hot gases comprising a housing, a bladed rotor disposed in said housing and adapted to spin at high speed, a plurality of generally arcuate members disposed in said housing adjacent the blades of said rotor, said arcuate members being independent, two-sided members spaced apart so as to define passages therebetween, through which passages hot gases can pass so as to act upon the blades of said rotor, said arcuate members being individually configured so as each to define a recess into which hot gases pass after acting upon said blades, with the passages and the recesses thus being disposed in a single alternating array on only one side of said blades, and exhaust means to which said recesses are connected so that hot gases can be carried away.

2. The turbine as recited in claim 1 in which a radial inflow-outflow device is defined, in which the hot gases are directed radially inwardly to act against said blades, with the flow leaving the blades flowing substantially radially outwardly in order to enter said recesses.

3. The turbine as recited in claim 1 in which a centripetal flow arrangement is defined, with the combustion products flowing between said arcuate members flowing radially outwardly so as to impinge upon the blades of said turbine, with the combustion products thereafter turning and then flowing substantially radially inwardly in order to enter said recesses.

4. The turbine as recited in claim 1 in which an axial flow device is defined, in which the combustion products are directed substantially axially in order to act upon the blades of said rotor and then thereafter turning to flow substantially axially away from said rotor in order to enter said recesses.

5. A turbine adapted to be operated by hot gases comprising a housing, a bladed rotor disposed in said housing and adapted to spin at high speed, a plurality of generally arcuate members disposed in said housing adjacent the blades of said rotor, said arcuate members being spaced apart so as to define passages therebetween, through which hot gases can pass so as to act upon the blades of said rotor, said arcuate members being individually configured so as each to define a recess into which hot gases pass after acting upon said blades, with the passages and the recesses thus being disposed in an alternating array on only one side of said blades, and exhaust means to which said recesses are connected so that hot gases can be carried away, and means defining a hot-cold boundary layer between said arcuate members and said blades, by which the hot gases are substantially prevented from impinging directly upon said turbine blades, said boundary layer resulting from a substantial flow of relatively cold gas under pressure from between said blades, each blade having an edge adjacent said arcuate members and another edge distant therefrom, said blades, during their rapid rotation, conducting cooling gas from between said distant edges to said boundary layer, and

thence into said recesses.

6. A turbine adapted to be operated by hot gases comprising a housing, a bladed rotor disposed in said housing and adapted to spin at high speed, a plurality of generally arcuate members disposed in said housing adjacent the blades of said rotor, said arcuate members being i spaced apart so as to define passages therebetween, through which hot gases can pass so as to act upon the blades of said rotor, said arcuate members being individually configured so as each to define a recess into which hot gases pass after acting upon said blades, with the passages and the recesses thus being disposed in an alternating array on only one side of said blades, and exhaust means to which said recesses are connected so that hot gases can be carried away, the hot gases moving toward said blades being met with flow of compressed gas delivered as a result of the rapid rotation of said blades, such that the hot gases are substantially prevented by the presence of the compressed gas from impinging directly upon the turbine blades, thus serving to substantially prolong the life of said blades.

Claims (6)

1. A turbine adapted to be operated by hot gases comprising a housing, a bladed rotor disposed in said housing and adapted to spin at high speed, a plurality of generally arcuate members disposed in said housing adjacent the blades of said rotor, said arcuate members being independent, two-sided members spaced apart so as to define passages therebetween, through which passages hot gases can pass so as to act upon the blades of said rotor, said arcuate members being individually configured so as each to define a recess into which hot gases pass after acting upon said blades, with the passages and the recesses thus being disposed in a single alternating array on only one side of said blades, and exhaust means to which said recesses are connected so that hot gases can be carried away.
2. The turbine as recited in claim 1 in which a radial inflow-outflow device is defined, in which the hot gases are directed radially inwardly to act against said blades, with the flow leaving the blades flowing substantially radially outwardly in order to enter said recesses.
3. The turbine as recited in claim 1 in which a centripetal flow arrangement is defined, with the combustion products flowing between said arcuate members flowing radially outwardly so as to impinge upon the blades of said turbine, with the combustion products thereafter turning and then flowing substantially radially inwardly in order to enter said recesses.
4. The turbine as recited in claim 1 in which an axial flow device is defined, in which the combustion products are directed substantially axially in order to act upon the blades of said rotor and then thereafter turning to flow substantially axially away from said rotor in order to enter said recesses.
5. A turbine adapted to be operated by hot gases comprising a housing, a bladed rotor disposed in said housing and adapted to spin at high speed, a plurality of generally arcuate members disposed in said housing adjacent the blades of said rotor, said arcuate members being spaced apart so as to define passages therebetween, through which hot gases can pass so as to act upon the blades of said rotor, said arcuate members being individually configured so as each to define a recess into which hot gases pass after acting upon said blades, with the passages and the recesses thus being disposed in an alternating array on only one side of said blades, and exhaust means to which said recesses are connected so that hot gases can be carried away, and means defining a hot-cold boundary layer between said arcuate members and said blades, by which the hot gases are substantially prevented from impinging directly upon said turbine blades, said boundary layer resulting from a substantial flow of relatively cold gas under pressure from between said blades, each blade having an edge adjacent said arcuate members and another edge distant therefrom, said blades, during their rapid rotation, conducting cooling gas from between said distant edges to said boundary layer, and thence into said recesses.
6. A turbine adapted to be operated by hot gases comprising a housing, a bladed rotor disposed in said housing and adapted to spin at high speed, a plurality of generally arcuate members disposed in said housing adjacent the blades of said rotor, said arcuate members being spaced apart so as to define passages therebetween, through which hot gases can pass so as to act upon the blades of said rotor, said arcuate members being individually configured so as each to define a recess into which hot gases pass after acting upon said blades, with the passages and the recesses thus being disposed in an alternating array on only one side of said blades, and exhaust means to which said recesses are connected so that hot gases can be carried away, the hot gases moving toward said blades being met with flow of compressed gas delivered as a result of the rapid rotation of said blades, such that the hot gases are substantially prevented by the presence of the compressed gas from impinging directly upon the turbine blades, thus serving to substantially prolong the life of said blades.
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Cited By (6)

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US4503669A (en) * 1983-02-25 1985-03-12 Henri Geoffroy Gas turbine thrust system
US5754637A (en) * 1995-04-17 1998-05-19 Samsung Electronics Co., Ltd. Collective house interphone system and method therefore
US6430917B1 (en) 2001-02-09 2002-08-13 The Regents Of The University Of California Single rotor turbine engine
US7044718B1 (en) 2003-07-08 2006-05-16 The Regents Of The University Of California Radial-radial single rotor turbine
US20090199812A1 (en) * 2003-03-21 2009-08-13 Jung Kuang Chou Structure of the rotary engine
US8839599B1 (en) * 2013-10-07 2014-09-23 Juan Pedro Mesa, Jr. Axial combustion engine

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US4397146A (en) * 1979-05-29 1983-08-09 Zepco, Inc. Gas turbine
GB9026748D0 (en) * 1990-12-08 1991-01-30 Tayler Colin A M Improvements in or relating to gas turbines
US20030210980A1 (en) * 2002-01-29 2003-11-13 Ramgen Power Systems, Inc. Supersonic compressor
US7334990B2 (en) * 2002-01-29 2008-02-26 Ramgen Power Systems, Inc. Supersonic compressor
US7434400B2 (en) * 2002-09-26 2008-10-14 Lawlor Shawn P Gas turbine power plant with supersonic shock compression ramps
US7293955B2 (en) * 2002-09-26 2007-11-13 Ramgen Power Systrms, Inc. Supersonic gas compressor
JP5705945B1 (en) * 2013-10-28 2015-04-22 ミネベア株式会社 Centrifugal fan

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US2783964A (en) * 1945-07-11 1957-03-05 Theimer Oscar Turbines
US2873945A (en) * 1952-11-06 1959-02-17 Garrett Corp Radial wheel construction
US3283509A (en) * 1963-02-21 1966-11-08 Messerschmitt Boelkow Blohm Lifting engine for vtol aircraft
US3310940A (en) * 1965-10-07 1967-03-28 Stalker Corp Gas turbines

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US1349487A (en) * 1917-05-31 1920-08-10 Erastus S Bennett Turbine-engine
US1882630A (en) * 1929-09-30 1932-10-11 Jarvis Christopher Turbine
US2783964A (en) * 1945-07-11 1957-03-05 Theimer Oscar Turbines
US2537344A (en) * 1945-08-06 1951-01-09 Francis K Gruss Turbine compressor
US2658338A (en) * 1946-09-06 1953-11-10 Leduc Rene Gas turbine housing
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US2873945A (en) * 1952-11-06 1959-02-17 Garrett Corp Radial wheel construction
US3283509A (en) * 1963-02-21 1966-11-08 Messerschmitt Boelkow Blohm Lifting engine for vtol aircraft
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Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4503669A (en) * 1983-02-25 1985-03-12 Henri Geoffroy Gas turbine thrust system
US5754637A (en) * 1995-04-17 1998-05-19 Samsung Electronics Co., Ltd. Collective house interphone system and method therefore
US6430917B1 (en) 2001-02-09 2002-08-13 The Regents Of The University Of California Single rotor turbine engine
US20090199812A1 (en) * 2003-03-21 2009-08-13 Jung Kuang Chou Structure of the rotary engine
US7044718B1 (en) 2003-07-08 2006-05-16 The Regents Of The University Of California Radial-radial single rotor turbine
US8839599B1 (en) * 2013-10-07 2014-09-23 Juan Pedro Mesa, Jr. Axial combustion engine

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