US3037742A - Compressor turbine - Google Patents

Compressor turbine Download PDF

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Publication number
US3037742A
US3037742A US84192159A US3037742A US 3037742 A US3037742 A US 3037742A US 84192159 A US84192159 A US 84192159A US 3037742 A US3037742 A US 3037742A
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Prior art keywords
shroud
blades
rotor
band
compressor
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Expired - Lifetime
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Dent Eugene
Robert E Weiser
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Motors Liquidation Co
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Motors Liquidation Co
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02Blade-carrying members, e.g. rotors
    • F01D5/022Blade-carrying members, e.g. rotors with concentric rows of axial blades

Description

4 Sheets-Sheet 1 Filed Sept. 17, 195 9 June 5, 1962 E. DENT ETAL 3,037,742

COMPRESSOR TURBINE Filed Sept. 17, 1959 4 Sheets-Sheet 2 IN V! N TOPS ATTORNEY June 5, 1962 T ETAL 3,037,742

COMPRESSOR TURBINE Filed Sept. 17, 1959 4 Sheets-Sheet 3 INVENTORS Y A M ATTORNEY June 5, 1962 E. DENT ETAL COMPRESSOR TURBINE 4 Sheets-$heet 4 Filed Sept. 17, v1959 United States Patent 3,037,742 COMPRESEOR TURBINE Eugene Dent, Indianapolis, Ind., and Robert E. Weiser,

San Diego, Calif., assignors to General Motors Corporation, Detroit, Mich., a corporation of Delaware Filed dept. 17, 1959, Ser. No. 841,921 7 Claims. ((11. 253-77) This invention relates to the rotors of turbomachinery, more particularly to a two-tier coaxial blade rotor construction for a gas turbine power plant wherein the outer tier comprises turbine blading, and more specifically to an improved shroud construction intermediate the two tiers of blading.

In a gas turbine power plant of the type described, the shroud intermediate the two tiers of blades not only serves to support the Widely spaced outer tips of the compressor blading and the peripheral turbine blading, but further improves the efiiciency of the axial flow compressor, acts as a sealing means between the gas turbine and the compressor chambers and further constitutes a thermal barrier therebetween. The provision of such a shroud involvcs considerable diifculty in the design and assembly of such a rotor, however. The spacing of the compressor blade tips and the turbine blades must be accurately maintained with respect to such a shroud and the associated structure. As the rotational speed of such a rotor is relatively high, the shroud must also be designed to withstand the necessarily high centrifugal forces acting thereon. In most cases, however, it is particularly important that the structure of the rotor be kept as light as possible. If the shroud is heavy, it not only contributes unnecessary weight directly to the rotor structure but it may also require a heavier disk to withstand the additional centrifugal force imposed thereon by the shroud mass. It is therefore desirable that the rotor shroud be made as light as possible but have suflicient strength to withstand the centrifugal forces acting thereon.

The invention contemplates an improved lightweight shroud structure which may be easily and economically fabricated and to which the compressor blade tips and the turbine blades may be easily and securely attached to provide a relatively light coaxial blade rotor structure of extremely high strength.

The foregoing and other objects, advantages and features of the invention will be apparent to those skilled in the art from the following detailed description of two preferred embodiments thereof having reference to the accompanying drawings, in which:

FIGURE 1 is a fragmentary rear view of a portion of a two-tier, coaxial turbine-compressor rotor embodying a first form of the invention, the view being taken in a plane perpendicular to the axis of the rotor;

FIGURE 2 is a fragmentary perspective view of a portion of the rotor shown in FIGURE 1 and shows details of the shroud and blade structure in broken-away section;

FIGURE 3 is a fragmentary perspective view showing the tip of one of the compressor blades as used in the form of the invention shown in FIGURE 1;

FIGURE 4 is a fragmentary view showing the mounting of the turbine blades on the shroud structure in a peripheral elevation or development of the rotor of FIG- URE 1;

FIGURE 5 is a fragmentary sectional view showing a second form of two-tier, coaxial turbine-compressor rotor stage embodying the invention with portions thereof broken away and in section to show the details of construction, the primary section of the view being taken in a plane including the axis of the rotor; and

FIGURE 6 is a fragmentary perspective view of a portion of the rotor shown in FIGURE 5 showing further details of this form of the invention in broken-away section.

Referring first to FIGURE 1, which sufiiciently illustrates a compressor rotor structure incorporating the in vention, the rotor comprises a rotor disk or Wheel 10, only a portion of the periphery of which is illustrated. The periphery of the wheel includes a rim 11 having a plurality of equiangularly spaced dovetail slots 12 therein extending longitudinally across the rim or, in other words, axially of the rotor. The dovetail slots 12 serve to mount a ring or row of compressor blades 13 which extend radially from the rim of the rotor. Each blade comprises an airfoil or blade portion 14, a platform 15, a stalk 16 and a root 17 which is configured to fit one of the dovetail slots 12. Since the platform and blade are of greater axial extent than the root, webs 18 extend obliquely from the faces of the root and stalk to the underside of the platform. The blade platforms engage along their edges to provide a frusto-conical ring defining the inner boundary of the air flow path through the compressor. An interrupted flange 19 extends radially outwardly from the rim 11 between the stalks to the inner surfaces delined by the platforms 15 and provides a sealing means preventing air leakage between the inlet and discharge sides of the rotor.

It will be noted that the compressor blades are relatively long and widely spaced at their tips. The compressor blades are also disposed at a considerable angle or skewed to the rotor axis, which is perpendicular to the plane of FIGURE 1. As seen in FIGURES 2 and 3, the tips of the compressor blades are notched to provide a plurality of prongs or dovetail projections 20 which define a plurality of equally spaced parallel side face portions which are normal to the rotor axis. These pronged ends of the compressor blades mount a shroud ring 21.

In accordance with the invention, the shroud ring 21 is of a brazed fabricated sheet metal construction comprising an inner sheet metal band 22, the outer sheet metal band 25, and an intermediate corrugated sheet metal band 27 sandwiched there'oetween. The inner band 22 is of channeled cross-section facing outwardly and providing two axially spaced, radially extending seal flanges 23 interconnected by an annular base section 24. The outer sheet metal band 25 is of oppositely disposed channeled cross-section and is nested within the channel of the inner band 22, having two inclined end fiages 26 which extend obliquely and inwardly from an outer base section 30 and are secured by brazing to the base portion 24 of the inner band immediately adjacent the seal flanges 23. The intermediate corrugated band 27 is of rectangularly sinuous cross-section defining a number of intermediate radial webs 28 and two outwardly extending radial webs 28 at opposite ends thereof. The inner and outer ends of the webs 28 and 28 are joined by axially extending webs 29 and 29' which are brazed to the base sections 24 and 30 of the inner and outer shroud bands, respectively. The outer peripheries of the two end webs 28 are also brazed annularly to the outer band flanges 26 thus reinforcing these flanges and increasing the overall stiffness of the shroud assembly.

The annular base section of the inner band 22 is provided with a number of openings or cutouts 31. These cutouts mate with the inwardly facing channels defined by the radial webs 28 and are disposed to receive the dovetail prongs 20 of the several compressor blade tips during assembly of the rotor. When thus assembled, the oppositely disposed side face portions of the several prongs engage the mating faces of the radial webs 28 of the shroud assembly. The dovetail joints thus formed between the pronged tips of the several compressor blades and the shroud ring are then preferably brazed securing the prongs 20 to the radial webs 23 and securing the tips of the several compressor blades to the inner band 22. It will be seen that this shroud construction is of extremely light weight but when integrated by brazing to the several compressor blade tips is of sufiicient strength and rigidity to withstand the centrifugal forces acting thereon and to maintain the angular spacing of the several compressor blade tips.

The lightweight shroud construction described above is further reinforced by mounting an outer tier of hollow turbine blades 32 circumferentially thereon. The turbine blades 32 may be suitably fabricated from sheet metal stock or may be hollow cast and subsequently machined or otherwise formed to the desired configuration. In the form of the invention shown in FIGURES 1-4, each turbine blade is secured to the shroud assembly by means of a flanged root element 33. Each root element 33 is suitably formed from sheet metal stock to provide a first base flange 33 mateable with and brazed to the circumferential surface of the shroud section and a second flange 33" extending outwardly therefrom. The outwardly extending flange of each root or blade mounting element is embraced by and brazed to the radially inner end of its associated turbine blade. As best seen in FIG- URE 4, the several turbine blades 32 and their respective flanged root elements 33 overlap circumferentially of the shroud 21 and serve to further increase the overall strength and rigidity of the shroud construction while while minimizing the additional mass carried thereby.

The rotor illustrated in FIGURES S and 6 differs from that of the previously described embodiment in that it is of somewhat greater compressor capacity and consequently includes two rows of outer blading which are axially spaced and cooperate with stator blading or nozzle means to provide a two-stage turbine. With respect to the invention, however, this embodiment differs primarily in the structure of the shroud ring and the mounting of the several turbine blades thereon.

In this form of the inventon, the shroud ring indicated generally at 34 comprises a fabricated annular sheet metal portion 35 which extends axially between two spaced end seal rings 36. These end rings may be cast or otherwise suitably formed. The fabricated sheet metal portion of the shroud 34 comprises three radially spaced concentric cylindrical sheet metal bands 37, 38 and 39. Two corrugated sheet metal bands and 41 are sandwitched between the cylindrical bands 37 and 38 and between the bands 38 and 39, respectively. As in the preceding embodiment, these corrugated bands 40 and 41 are of rectangularly sinuous cross-section and thus define a number of radial webs 42 and 43, respectively, which are joined by axially extending Webs 44, 44' and 45, 45, respectively. The axially extending webs 44 and 44 are brazed to the inner and intermediate circular bands 37 and 38, respectively, and the axial Webs 45, 45 of the outer corrugated band 41 are similarly brazed to the intermediate and outer circular bands 38 and 39, respectively. The end rings 36 are each provided with two concentric flanges 46 and 47. The opposite ends of the inner circular band 37 and the adjacent radial end webs 42 of the corrugated band 40 are brazed to the inner end ring flanges 46. The outer end ring flanges 47 are similarly brazed to the adjacent ends of the outer circular band 39 and of the radial end webs 43 of the corrugated band 41. The opposite ends of the intermediate circular band are also brazed to the end rings 36 intermediate the inner and outer flanges 46 and 47 thereof.

A in the preceding embodiment, the inner band 37 is provided with a number of cutouts 48 which open on the inwardly facing channels defined by the inner corrugated band 40. These cutouts are disposed and adapted to receive the dovetailed pronged tips 49 of compressor blades 50. Upon assembly of the shroud on the tips of the compressor blades, as in the previous embodiment, the pronged tips 49 of the compresor blades are secured by brazing to the radial webs 42 of the inner corrugated band 40 and to the inner cylindrical band 37 to provide an integrated light Weight shrouded rotor structure of relatively high strength.

In this embodiment of the invention, the turbine blades 51 are secured to the shroud assembly in the same manner as the compressor blade tips, the radially inner ends of the turbine blades being provided with a plurality of dovetail prongs 52. These prongs are insertable through cutouts 53 which are provided therefor in the outer cylindrical band 39 and open on the outwardly facing channels defined by the radial webs 43 of the outer corrugated band 41. When the turbine blades have been thus assembled on the periphery of the shroud, the prongs 52 and the radially inner ends of the turbine blades are secured by brazing to the intermediate and outer circular bands 38 and 39 and the intermediate corrugated band 41 thereby further reinforcing the relatively lightweight shroud structure.

From the foregoing it will be seen that the invention accomplishes its several objectives of providing an improved rotor having a relatively lightweight shroud of high strength and rigidity fabricated wholly or in the main of sheet metal components, the strength and rigidity of the sheet metal shroud structure being further fortified by the brazed dovetail mounting of the shroud on the several compressor blade tips and by the alternative forms of turbine blade mounting on the outer periphery of the shroud.

While only two embodiments of the invention have been shown and described for the purpose of explaining the principals thereof, it will be apparent to those skilled in the art that many modifications may be made therein without departing from the scope of the invention, as defined in the following claims.

We claim:

1. A turbomachine comprising, in combination, a rotor, a first row of blades mounted on said rotor and extending radially outwardly therefrom, said blades being skewed to the axis of the rotor and each having a number of aligned spaced projections extending outwardly from the outer end thereof in parallel relation normal to the axis of said rotor, a fabricated annular shroud assembly embracing and securing the outer ends of said first blades in equiangular spaced relation to each other, said shroud assembly comprising an outer band, an inner band, and a rectangularly sinuous band intermediate and bonded to said outer and inner bands, said intermediate band having radially and circumferentially extending webs with the ends thereof extending transversely of the shroud, and said inner band having a plurality of openings therein disposed to receive the projections of the several blades intermediate the radial webs of said intermediate band, said projections being bonded to said inner and intermediate bands thereby forming a plurality of shroud reinforcing dovetail joints therewith, and a second row of blades mounted on and secured to the periphery of said shroud, said second row having a plurality of blades secured to and reinforcing said shroud intermediate the ends of said first blades.

2. A rotor element for a turbornachine comprising a rotor disk, a first row of blades mounted on said rotor disk and extending radially outwardly therefrom, said blades being skewed to the axis of the rotor and having a number of aligned spaced projections extending outwardly from the outer end thereof in parallel relation normal to the axis of said rotor, a fabricated sheet metal shroud element of corrugated ply construction embracing and securing the outer ends of said first blades in equiangularly spaced relation to each other, said shroud having a plurality of openings therein disposed to receive the projections of the several blades intermediate the corrugations of said shroud, and the corrugations of said shroud and said projections being bonded to form a plurality of shroud reinforcing dovetail joints therebetween.

3. A rotor element as set forth in claim 2, including a second row of blades mounted on and secured to the periphery of said shroud, said second row having a plurality of blades secured to and reinforcing said fabricated shroud intermediate the ends of said first blades.

4. A rotor element for a turbomachine comprising a rotor disk, a row of compressor blades mounted on said rotor disk and extending radially outwardly therefrom, said blades being skewed to the axis of the rotor and having a number of aligned spaced projections extending outwardly from the outer end thereof in parallel relation normal to the axis of said rotor, a fabricated shroud assembly annularly embracing and securing the outer ends of said compressor blades in equiangular spaced relation to each other, said shroud assembly comprising three concentric radially spaced sheet metal bands and two rectangularly sinuous sheet metal bands intermediate and bonded to said radially spaced bands, said sinuous bands having radially and circumferentially extending Webs With the ends thereof extending transversely of the shroud, the innermost of said concentric bands having a plurality of openings therein disposed to receive the projections of the compressor blades intermediate the radial Webs of the innermost sinuous band, a row of hollow turbine blades mounted on and secured to the periphery of said shroud, said turbine blades having a number of aligned spaced projections extending radially inwardly therefrom, the outermost of said concentric bands having a plurality of openings disposed to receive the projections of the turbine blades intermediate the radial Webs of the outermost sinuous band, and said compressor and turbine blade projections engaging and being bonded to said sinuous bands thereby forming a plurality of shroud reinforcing dovetail joints therewith.

5. A rotor element as set forth in claim 4, wherein said shroud assembly includes two annular end members secured to and supporting the opposite axial ends of said concentric and sinuous bands.

6. A rotor eiement for a turbomachine comprising a rotor disk, a row of compressor blades mounted on said rotor disk and extending radially outwardly therefrom, said blades being skewed to the axis of the rotor and each having a number of aligned spaced projections extending outwardly from the outer end thereof in parallel relation normal to the axis of said rotor, a fabricated annular shroud assembly embracing and securing the outer ends of said first blades in equiangular spaced relation to each other, said shroud assembly comprising an outer band, an inner band, and a substantially rectangular sinuous band intermediate and bonded to said outer and inner bands, said sinuous band having radially and circumferentially extending webs with the ends thereof extending transversely of the shroud, said inner band having a plurality of openings disposed to receive the projections of the several blades intermediate the radial Webs of said intermediate band, and said projections being bonded to said inner and intermediate bands thereby forming a plurality of shroud reinforcing dovetail joints therewith.

7. A rotor element as set forth in claim 6, including a row of turbine blades mounted on and reinforcing the periphery of said shroud intermediate the ends of said first blades, each of said turbine blades comprising a hollow blade portion and a flanged root element bonded to said blade portion and to said outer band.

References Cited in the file of this patent UNITED STATES PATENTS 2,398,113 Parrish Apr. 9, 1946 2,771,622 Thorp Nov. 27, 1956 2,801,789 Moss Aug. 6, 1957 2,912,222 Wilkes Nov. 10, 1959 2,971,745 Warren et a1 Feb. 14, 1961 FOREIGN PATENTS 548,433 Italy Sept. 25, 1956

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Cited By (47)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3269120A (en) * 1964-07-16 1966-08-30 Curtiss Wright Corp Gas turbine engine with compressor and turbine passages in a single rotor element
US3494539A (en) * 1967-04-03 1970-02-10 Rolls Royce Fluid flow machine
US4648802A (en) * 1984-09-06 1987-03-10 Pda Engineering Radial flow rotor with inserts and turbine utilizing the same
US20070022738A1 (en) * 2005-07-27 2007-02-01 United Technologies Corporation Reinforcement rings for a tip turbine engine fan-turbine rotor assembly
US20070295011A1 (en) * 2004-12-01 2007-12-27 United Technologies Corporation Regenerative Turbine Blade and Vane Cooling for a Tip Turbine Engine
US20080014078A1 (en) * 2004-12-01 2008-01-17 Suciu Gabriel L Ejector Cooling of Outer Case for Tip Turbine Engine
US20080093174A1 (en) * 2004-12-01 2008-04-24 Suciu Gabriel L Tip Turbine Engine with a Heat Exchanger
US20090071162A1 (en) * 2004-12-01 2009-03-19 Suciu Gabriel L Peripheral combustor for tip turbine engine
US20090142184A1 (en) * 2004-12-01 2009-06-04 Roberge Gary D Vectoring transition duct for turbine engine
US20090148273A1 (en) * 2004-12-01 2009-06-11 Suciu Gabriel L Compressor inlet guide vane for tip turbine engine and corresponding control method
US20090145136A1 (en) * 2004-12-01 2009-06-11 Norris James W Tip turbine engine with multiple fan and turbine stages
US20090155079A1 (en) * 2004-12-01 2009-06-18 Suciu Gabriel L Stacked annular components for turbine engines
US20090232650A1 (en) * 2004-12-01 2009-09-17 Gabriel Suciu Tip turbine engine and corresponding operating method
US7631480B2 (en) 2004-12-01 2009-12-15 United Technologies Corporation Modular tip turbine engine
US7845157B2 (en) 2004-12-01 2010-12-07 United Technologies Corporation Axial compressor for tip turbine engine
US7874163B2 (en) 2004-12-01 2011-01-25 United Technologies Corporation Starter generator system for a tip turbine engine
US7874802B2 (en) 2004-12-01 2011-01-25 United Technologies Corporation Tip turbine engine comprising turbine blade clusters and method of assembly
US7878762B2 (en) 2004-12-01 2011-02-01 United Technologies Corporation Tip turbine engine comprising turbine clusters and radial attachment lock arrangement therefor
US7883315B2 (en) 2004-12-01 2011-02-08 United Technologies Corporation Seal assembly for a fan rotor of a tip turbine engine
US7882694B2 (en) 2004-12-01 2011-02-08 United Technologies Corporation Variable fan inlet guide vane assembly for gas turbine engine
US7882695B2 (en) 2004-12-01 2011-02-08 United Technologies Corporation Turbine blow down starter for turbine engine
US7883314B2 (en) 2004-12-01 2011-02-08 United Technologies Corporation Seal assembly for a fan-turbine rotor of a tip turbine engine
US7887296B2 (en) 2004-12-01 2011-02-15 United Technologies Corporation Fan blade with integral diffuser section and tip turbine blade section for a tip turbine engine
US7927075B2 (en) 2004-12-01 2011-04-19 United Technologies Corporation Fan-turbine rotor assembly for a tip turbine engine
US7934902B2 (en) 2004-12-01 2011-05-03 United Technologies Corporation Compressor variable stage remote actuation for turbine engine
US7937927B2 (en) 2004-12-01 2011-05-10 United Technologies Corporation Counter-rotating gearbox for tip turbine engine
US7959532B2 (en) 2004-12-01 2011-06-14 United Technologies Corporation Hydraulic seal for a gearbox of a tip turbine engine
US7959406B2 (en) 2004-12-01 2011-06-14 United Technologies Corporation Close coupled gearbox assembly for a tip turbine engine
US7976273B2 (en) 2004-12-01 2011-07-12 United Technologies Corporation Tip turbine engine support structure
US7976272B2 (en) 2004-12-01 2011-07-12 United Technologies Corporation Inflatable bleed valve for a turbine engine
US8024931B2 (en) 2004-12-01 2011-09-27 United Technologies Corporation Combustor for turbine engine
US8033092B2 (en) 2004-12-01 2011-10-11 United Technologies Corporation Tip turbine engine integral fan, combustor, and turbine case
US8033094B2 (en) 2004-12-01 2011-10-11 United Technologies Corporation Cantilevered tip turbine engine
US8061968B2 (en) 2004-12-01 2011-11-22 United Technologies Corporation Counter-rotating compressor case and assembly method for tip turbine engine
US8083030B2 (en) 2004-12-01 2011-12-27 United Technologies Corporation Gearbox lubrication supply system for a tip engine
US8096753B2 (en) 2004-12-01 2012-01-17 United Technologies Corporation Tip turbine engine and operating method with reverse core airflow
US8152469B2 (en) 2004-12-01 2012-04-10 United Technologies Corporation Annular turbine ring rotor
US8365511B2 (en) 2004-12-01 2013-02-05 United Technologies Corporation Tip turbine engine integral case, vane, mount and mixer
US8468795B2 (en) 2004-12-01 2013-06-25 United Technologies Corporation Diffuser aspiration for a tip turbine engine
US8561383B2 (en) 2004-12-01 2013-10-22 United Technologies Corporation Turbine engine with differential gear driven fan and compressor
US8641367B2 (en) 2004-12-01 2014-02-04 United Technologies Corporation Plurality of individually controlled inlet guide vanes in a turbofan engine and corresponding controlling method
US8757959B2 (en) 2004-12-01 2014-06-24 United Technologies Corporation Tip turbine engine comprising a nonrotable compartment
US8807936B2 (en) 2004-12-01 2014-08-19 United Technologies Corporation Balanced turbine rotor fan blade for a tip turbine engine
US8967945B2 (en) 2007-05-22 2015-03-03 United Technologies Corporation Individual inlet guide vane control for tip turbine engine
US9003759B2 (en) 2004-12-01 2015-04-14 United Technologies Corporation Particle separator for tip turbine engine
US9109537B2 (en) 2004-12-04 2015-08-18 United Technologies Corporation Tip turbine single plane mount
US9845727B2 (en) 2004-12-01 2017-12-19 United Technologies Corporation Tip turbine engine composite tailcone

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3269120A (en) * 1964-07-16 1966-08-30 Curtiss Wright Corp Gas turbine engine with compressor and turbine passages in a single rotor element
US3494539A (en) * 1967-04-03 1970-02-10 Rolls Royce Fluid flow machine
US4648802A (en) * 1984-09-06 1987-03-10 Pda Engineering Radial flow rotor with inserts and turbine utilizing the same
US20080093174A1 (en) * 2004-12-01 2008-04-24 Suciu Gabriel L Tip Turbine Engine with a Heat Exchanger
US9003759B2 (en) 2004-12-01 2015-04-14 United Technologies Corporation Particle separator for tip turbine engine
US9003768B2 (en) 2004-12-01 2015-04-14 United Technologies Corporation Variable fan inlet guide vane assembly, turbine engine with such an assembly and corresponding controlling method
US20070295011A1 (en) * 2004-12-01 2007-12-27 United Technologies Corporation Regenerative Turbine Blade and Vane Cooling for a Tip Turbine Engine
US20080014078A1 (en) * 2004-12-01 2008-01-17 Suciu Gabriel L Ejector Cooling of Outer Case for Tip Turbine Engine
US9541092B2 (en) 2004-12-01 2017-01-10 United Technologies Corporation Tip turbine engine with reverse core airflow
US20090071162A1 (en) * 2004-12-01 2009-03-19 Suciu Gabriel L Peripheral combustor for tip turbine engine
US20090142184A1 (en) * 2004-12-01 2009-06-04 Roberge Gary D Vectoring transition duct for turbine engine
US20090148273A1 (en) * 2004-12-01 2009-06-11 Suciu Gabriel L Compressor inlet guide vane for tip turbine engine and corresponding control method
US20090145136A1 (en) * 2004-12-01 2009-06-11 Norris James W Tip turbine engine with multiple fan and turbine stages
US20090155079A1 (en) * 2004-12-01 2009-06-18 Suciu Gabriel L Stacked annular components for turbine engines
US20090232650A1 (en) * 2004-12-01 2009-09-17 Gabriel Suciu Tip turbine engine and corresponding operating method
US7607286B2 (en) 2004-12-01 2009-10-27 United Technologies Corporation Regenerative turbine blade and vane cooling for a tip turbine engine
US7631485B2 (en) 2004-12-01 2009-12-15 United Technologies Corporation Tip turbine engine with a heat exchanger
US7631480B2 (en) 2004-12-01 2009-12-15 United Technologies Corporation Modular tip turbine engine
US7845157B2 (en) 2004-12-01 2010-12-07 United Technologies Corporation Axial compressor for tip turbine engine
US7854112B2 (en) 2004-12-01 2010-12-21 United Technologies Corporation Vectoring transition duct for turbine engine
US7874163B2 (en) 2004-12-01 2011-01-25 United Technologies Corporation Starter generator system for a tip turbine engine
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US7882695B2 (en) 2004-12-01 2011-02-08 United Technologies Corporation Turbine blow down starter for turbine engine
US7883314B2 (en) 2004-12-01 2011-02-08 United Technologies Corporation Seal assembly for a fan-turbine rotor of a tip turbine engine
US7887296B2 (en) 2004-12-01 2011-02-15 United Technologies Corporation Fan blade with integral diffuser section and tip turbine blade section for a tip turbine engine
US7921636B2 (en) 2004-12-01 2011-04-12 United Technologies Corporation Tip turbine engine and corresponding operating method
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US7927075B2 (en) 2004-12-01 2011-04-19 United Technologies Corporation Fan-turbine rotor assembly for a tip turbine engine
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US20110142601A1 (en) * 2004-12-01 2011-06-16 Suciu Gabriel L Variable fan inlet guide vane assembly, turbine engine with such an assembly and corresponding controlling method
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