US3804335A - Vaneless supersonic nozzle - Google Patents

Vaneless supersonic nozzle Download PDF

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Publication number
US3804335A
US3804335A US00362403A US36240373A US3804335A US 3804335 A US3804335 A US 3804335A US 00362403 A US00362403 A US 00362403A US 36240373 A US36240373 A US 36240373A US 3804335 A US3804335 A US 3804335A
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Prior art keywords
converging
diverging
nozzle unit
chamber
common
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US00362403A
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J Sohre
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Individual
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Individual
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Priority to US00362403A priority Critical patent/US3804335A/en
Priority to CA195,114A priority patent/CA999527A/en
Priority to AU67072/74A priority patent/AU6707274A/en
Priority to GB1408874A priority patent/GB1457634A/en
Priority to NL7404887A priority patent/NL7404887A/xx
Priority to IT50338/74A priority patent/IT1004228B/it
Priority to DE19742417880 priority patent/DE2417880A1/de
Application granted granted Critical
Priority to JP49043232A priority patent/JPS508906A/ja
Publication of US3804335A publication Critical patent/US3804335A/en
Priority to US05/649,478 priority patent/USRE29128E/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D1/00—Non-positive-displacement machines or engines, e.g. steam turbines
    • F01D1/02—Non-positive-displacement machines or engines, e.g. steam turbines with stationary working-fluid guiding means and bladed or like rotor, e.g. multi-bladed impulse steam turbines
    • F01D1/04—Non-positive-displacement machines or engines, e.g. steam turbines with stationary working-fluid guiding means and bladed or like rotor, e.g. multi-bladed impulse steam turbines traversed by the working-fluid substantially axially
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades

Definitions

  • ABSTRACT A converging-diverging nozzle unit for supersonic expansion in turbines is presented wherein a plurality of converging nozzle segments open into a common diverging chamber at or near the throat of the converging nozzles. Convergence and divergence may occur in different directions in the nozzle unit; with reference to a rotating turbine wheel, convergence occurs in a direction substantially tangential with respect thereto, and divergence occurs substantially radial with respect thereto.
  • This invention relates to the field of turbomachinery. More particularly, this invention relates to the field of convergent-divergent nozzle structure for creation and expansion of supersonic flow of a compressible fluid for turbine motive fluid.
  • Jet deflection is a problem particularly associated with pressure ratios above design level, and thus there is a direct relationship between jet deflection and Mach number.
  • Pressure ratio refers, of course, to the ratio of the pressure at the inlet of the nozzle to the pressure in the area into which the nozzle discharges, commonly referred to as thetback pressure.
  • the exit angle of the fluid from the nozzles is also related to and affected by the pressure ratio.
  • the exit angle is the actue angle of the axis of the flow stream with the exit plane of the nozzles.
  • a turbine is typically designed to operate at a particular pressure ratio and exit angle; and increases in the pressure ratio willcause a deflection of the nozzle jet in a direction to increase the exit angle.
  • This jet deflection can result in shock waves, expansion waves, and pressure gradients with resultant stresses on the buckets, i.e., the rotating turbine blades.
  • blade failures can occur, axial thrust on the rotating structure can increase and fluctuate unpredictably, and an overall highlyinefficient operation results.
  • Jet deflection also results in an adverse effect on entry of the jet stream into the bucket passages.
  • the cross-sectional area of the jet stream entering the buckets is a direct function of the nozzle exit angle. Jet deflection from increased pressure ratio operation results in an increase in the cross-sectional area of the jet stream entering a bucket passage, and the buckets will not function properly if the cross-sectional area of the entering st ream exceeds a design multiple of the throat area of the bucket passage.
  • One type has convergent-divergent profiles machined into opposed side surfaces of adjacent nozzles so as to define, in one dimension, convergent-divergent passages between these profiled side surfaces.
  • the top and bottom surfaces of such passages are parallel to each other to define a constant height for the convergent-divergent passage or the height may vary linear between nozzle inlet and exit.
  • the other type of conventional nozzle is in the form of round nozzle passages, as used in drilled and reamed nozzle blocks. These nozzle blocks may be circular in cross-section and thus two-dimensionally convergent-divergent. In any case, the divergence and the throat of the passage are not defined by the inner and outer circumferential walls of the passage.
  • the present invention overcomes or reduces the above discussed and other problems of the prior art with its significantly improved and novel nozzle unit.
  • the nozzle unit configuration of the present invention provides a common expansion chamber in conjunction with a plurality of converging nozzle segments.
  • the in dividual diverging passages of conventional converging-diverging nozzles are eliminated and, in their place, a common expansion chamber is provided.
  • the individual converging nozzle segments are connected at or near their throats to the common expansion chamber so that the steam or other turbine propulsive fluid undergoes expansion to the velocity of sound (Mach 1 in the converging portions of the individual nozzle segments and is then delivered to the common chamber for supersonic expansion to the final nozzle exit pressure.
  • the common expansion chamber is a diverging chamber,'but its direction of divergence may be angled, preferably at with respect to the direction of convergence of the individual converging nozzle passages.
  • subsonic expansion is in a direction substantially tangential to the turbine wheel and supersonic expansion utilizes space in a substantially radial direction.
  • convergence and divergence in the nozzle occur in different directions with respect to the direction of flow through the nozzle.
  • a deflection control wall or, a series of such walls can be incorporated in the common expansion chamber, inclined in a direction parallel to or substantially parallel to the axis of the desired jet flow at design conditions, and hence inclined oppositely to the direction of jet deflection normally experienced at increased pressure ratios.
  • This deflection control surface will function to prevent the normal jet deflection usually encountered at pressure ratios above the design point and will force the gas to expand and flow in the general direction for design conditions.
  • the usually experienced jet deflection with its attendant inefficiencies and other problems can be eliminated.
  • the bucket entry angle and crosssectional area of the entry stream will actually decrease, thus enabling the buckets to function properly over a wider range of conditions and without an increase of stage reaction (pressure drop across the buckets).
  • the individual nozzle segments can be extended into the common expansion chamber to provide a better defined throat and thus a more efficient configuration.
  • the major supersonic expansion is still accomplished by the increase in the radial dimension of the common expansion chamber in accordance with the general concept of the invention.
  • the nozzle configuration of the present invention with its common expansion chamber provides several important advantages over the prior art.
  • the amount of circumferential space and size required to handle a given amount of steam is reduced, and, similarly, the size increase needed to accommodate an increased volume is reduced, because the expansion is accomplished in the direction of the height of the nozzles, i.e., radial inward and/or outward with respect to the turbine wheel, .rather than in the width separating the nozzles orin the circular configuration of the round nozzles.
  • turbines incorporating the present invention can pass more steam at higher pressure ratios for a given are of circumference than in the prior art, and as a result high volumes can be handled at greater pressure ratios than presently obtainable. Also, machines can be built with fewer stages and thus can be smaller and lighter. Overall operating efficiency is greatly improved because flow interruptions are greatly reduced or eliminated, and force fluctuations on the buckets are reduced by virtue of the expansion which takes place inthe common passage thereby dissipating wakes and shock effects. It is important to note that the common expansion passages provide a separation between the indivdual nozzle elements and the buckets much greater than the usual separation of approximately 1/16 inch to 41 inch found in typical prior art turbines.
  • the common expansion chamber and the radial expansion configuration of the present invention can also be incorporated in the rotating bucket structure of certain turbines. This may be done where supersonic expansion is desirable to achieve high horsepower output from a minimum number of turbine wheels.
  • one object of the present invention is to provide a novel and improved convergent-divergent nozzle structure for a turbine.
  • Another object of the present invention is to provide a novel and improved convergent-divergent nozzle unit for a turbine wherein a plurality of nozzles have a common diverging expansion chamber.
  • Another object of the present invention is to provide a novel and improved convergent-divergent nozzle unit for a turbine wherein convergence and divergence take place in different directions with respect to the direction of flow through the nozzle.
  • Still another object of the present invention is to provide a novel and improved convergent-divergent nozzle structure for turbines wherein problems of jet deflection, jet separation, vibratory loading and loss of efficiency, commonly encountered in many prior art turbines, are significantly reduced.
  • Still another object of the present invention is to provide a novel and improved convergent-divergent bucket structure for rotating turbine wheels.
  • FIG. 1 is a generalized showing, partly in section, of a turbomachine and motive fluid entry passageway;
  • FIG. 2 is a view taken along line 2-2 of FIG. 1 showing an elevation view of the supersonic convergentdivergent nozzle unit of the present invention as applied to stationary nozzles;
  • FIG. 3 is a view taken along line 3-3 of FIG. 1 showing an arc of nozzles with their common expansion chamber in accordance with the present invention
  • FIG. 4 is a view taken along line 44 of FIG. 3;
  • FIG. 5 is a view similar to FIG. 4 showing a modified wall configuration
  • FIG. 5A is a cross-sectional view similar to part of FIG. 2 showing the wall contours in the modified configuration of FIG. 5;
  • FIG. 6 is a view similar to FIG. 3 showing a modified version of the invention incorporating backup wall;
  • FIG. 6A is a view showing part of FIG. 6 in detail
  • FIG. 7 is a partial view similar to FIG. 3 showing another modification of the present invention wherein vane elements extend into the common expansion chamber;
  • FIG. 8 is a view along line 8--8 of FIG. 7;
  • FIG. 9 is a view showing the common expansion chamber of thepresent invention incorporated in the rotating buckets of a reaction turbine wheel
  • FIG. 10 is a view along line l0--10 of FIG. 9;
  • FIG. 11 is a view similar to FIG. 10 showing a plurality of backup wallsj
  • FIG. 12 is a view along line 12-12 of FIG. 11 showing an enlarged and modified partial view similar to FIG. 9;
  • FIG. 13 is a view showing the present invention applied to the stationary nozzles and rotating buckets of a turbomachine.
  • FIGS. 1 and 2 the general environment of the present invention is shown in the nature of a turbomachin ery unit 10.
  • Turbomachinery unit 10 has an inlet chamber 12 to deliver steam or other motive fluid to the unit and a turbine casing 14 for housing a rotating turbine wheel 16 which is mounted for rotation about an axis 18.
  • Many other elements and details normally found in turbomachinery units are unnecessary to a proper understanding of the present invention, and thus these other details have been omitted from the simplified environment shown in FIGS. 1 and 2.
  • Inlet chamber 12 is in flow communication with an array of nozzle segments 20 which are arranged in an arcuate array on an arc corresponding to the radius of the blades or buckets 24 of turbine 16 with respect to axis 18.
  • the individual nozzle segments 20 form part of a convergent-divergent nozzle unit 22 which serves to direct the steam or the motive fluid from inlet chamber 12 to the blades or buckets 24 of turbine wheel 16.
  • Convergent-divergent nozzle unit 22 is comprised of the individual nozzle segments 20 which cooperate with each other to form a plurality of convergent flow paths 21, and a common diverging expansion chamber 26 which is in flow communication with the convergent paths 21 between the nozzles 20.
  • the steam or other motive fluid flows through inlet chamber 12 into the convergent flow paths between adjacent nozzles 20, and that flow may, if desired, be regulated by appropriate control valving.
  • the motive fluid is subsonically expanded in traversing the converging passageways between adjacent nozzles, and it is then supersonically expanded in chamber 26 and then delivered to the buckets 24 to drive turbine wheel 16.
  • the width between adjacent nozzles 20 diminishes, and hence the passageways 21 converge, in a direction generally tangential with respect to turbine wheel 16.
  • at least one of those walls, i.e., the right portion of wall 28 in FIG. 2 is inclined with re spect to a radius of the turbine wheel so that chamber 26 canbe said to expand radially with respect to the turbine wheel.
  • wall 30 could, if desired, also be inclined so that chamber 26 would expand both toward axis 18 as well as in the direction away from axis 18.
  • Wall 28 terminates in a segment 32 which is perpendicular with respect to the radius of wheel 16, and section 32 cooperates with a similar length of wall 28 to form a nozzle exit passage with an exit edge 33 immediately adjacent to the rotating wheel 16 and blades 24.
  • the nozzles 20 are arranged in an arcuate array, and that walls 28 and 30 are similarly arcuate to conform with the annular array of turbine blades 24.
  • FIG. 3 a view of the nozzle unit is presented taken along 33 of FIG. 1.
  • the array of individual nozzles 20 forming converging flow paths 21 therebetween can be seen, and these converging flow paths 21 are seen communicating with the common expansion chamber 26.
  • a throat 34 is defined in each passage 21 at the plane of smallest cross-section area of the passage, and the throat 34 of each convergent passage 21 communicates directly with the common expansion chamber 26.
  • the width of flow passages 21 is defined by the separation between adjacent nozzles, and the height of flow passages 21 is defined by the separation between the left portion of walls 28 and 30.
  • the necessary convergent-divergent structure is present in nozzle unit 22 for supersonic expansion of a compressible medium.
  • the steam flows through inlet chamber 12 and then into convergent passageways 21 defined by the confronting walls of adjacent nozzle segments 20.
  • the throats of passageways 21 are connected to expansion chamber 26, and the fluid in each of the passageways is delivered into the common diverging expansion chamber which, as indicated above, expands in a direction angled, preferably perpendicular with respect to the direction of convergence effected by nozzle segments 20.
  • nozzle segments 20 terminate a substantial distance from turbine blades 24. While that distance will, of course, vary for different designs, it will typically be on the order of several inches, such as from V2 to 10 or more inches, so that chamber 26 is of substantial length normal to turbine wheel 16.
  • chamber 26 provides the necessary volume to accommodate the superonic expansion of the fluid being discharged from each of the passageways 21 to transform the high energy levels of the fluid into a high-velocity jet for deliva ery to the turbine buckets 24. Since the trailing edges of the nozzles may be located in a subsonic region, this will eliminate all supersonic shock waves in the entire supersonic expansion passage 26.
  • the length of chamber 26 provides a long vaneless passage in which the superonic expansion is taking place, and this vaneless passage accomplishes flow adjustment and dissipation of wakes, shock waves, expansion waves and pressure gradients, thereby significantly improving the overall efficiency of the machinery and significantly reducing the problems of vibratory loading, blade and nozzle failures and rotor thrust loading which have accompanied jet deflection and flow separation in the prior art.
  • the problems previously encountered with flow interruption are also reduced or eliminated since the long vaneless passage 26 produces a substantially uniform continuous combined flow stream for delivery to the buckets 24.
  • the walls 28 and 30 appear to be curved as seen in FIG. 4 because of the location of line 44 in FIG. 3.
  • FIGS. 5 and 5A a modified configuration is shown wherein walls 28 and 30 are slightly curved as seen in FIG. 5A, thus resulting in an apparently straight configuration when viewed as in FIG. 5.
  • This modified shape has Prandtl corners at the junctions of walls 28 and 30 with throat 34. These are some times considered desirable because they illustrates the exact point of onset of supersonic expansion. Other configurations can be obtained in a similar manner.
  • This example illustrated the capability of the invention to accommodate a wide range of thermodynamic requirements with only minor changes of hardware configuration. A wide variety of conditions can thus be accommodated without a change in nozzle segments 20.
  • FIG. 6 a view similar to FIG. 3 is shown incorporating a modification in the form of a backup deflection control wall 36 in the common expansion chamber 26.
  • Wall 36 functions to control jet deflection at pressure ratios above design level.
  • the angle or shown in FIG. 6 represents the angle between the flow axis of a discharge stream from a passage 21 and the discharge plane from the nozzles.
  • the angle a thus adjusts as a function of the ratio P /P in a direction opposite to that previously encountered, the commonly experienced shock and flow separation problems are avoided or reduced for off design conditions, and thus the nozzle unit of the present invention results in a significant improvement in nozzle efficiency.
  • a major problem heretofor encountered in supersonic bucket entry relates to a relationship between the area f, of the entering stream (determined in the plane normal to the direction of the flow stream from the nozzle discharge to the bucket) and the cross-sectional area f of the flow passage between adjacent buckets. If the ratio f exceeds a certain level, the bucket will not function properly. As can be seen in FIG. 6A, an increase in a a condition normally encounted in the prior art, results in an increase in f,, and the buckets will not function properly (or choke) if that increase becomes excessive.
  • FIGS. 7 and 8 still another modification of the present invention is shown.
  • the trailing edges 20(a) of the nozzles 20 are continued beyond throat 34 into divergent chamber 26, and they may extend partway into the chamber.
  • the passage 40 formed between the trailing edges may be parallel sided, converging or diverging, straight or curved.
  • the major part of total divergence, and hence gas expansion, between throat 34 and exit edge 38 is still accomplished in the radial direction by the presence of sloping wall 28, and the passage throat is also defined by walls 28 and/or 30.
  • wall 30 can also be sloping in the direction opposite to the slope of wall 28.
  • trailing edges 20(a) into the common radially extending expansion chamber 26 provides a somewhat better defined throat for the passageways 21 thus contributing to increased efficiency.
  • the major expansion in chamber 26 still occurs in a direction perpendicular to the passage formed by the nozzle-profiles 20 from the entry of the nozzles to throat 34.
  • the location of the throat i.e., the smallest crosssectional flow area, can be defined by contouring one or both of the wall segments 28(a) and 30(a) upstream of the throat.
  • the throat indicated at 34(a) in FIG. 8 can be defined in the radial plane with respect to the turbine wheel by varying the passage walls upstream of the throat rather than by varying the separation between the vanes as is now done in the art.
  • This wall contouring to define the throat is shown in the dotted lines in FIG. 8, but it will be understood it can be incorporated in any of the embodiments of the invention.
  • total supersonic expansion can be adjusted and determined by varying the height of the expansion chamber, i.e., by varying the distance between walls 28 and 30.
  • the contouring of wall 28 in the solid line will result in one expansion ratio, while contouring of wall 28 in the dotted line will result in a different expansion ratio.
  • the present invention makes it possible to vary expansion ratios by varying the height of the expansion chamber rather than by varying the separation between nozzles as is presently done in the art. It thus becomes possible with the present invention to stock standard nozzle blocks and use those standard nozzle blocks in installations where different expansion ratios are required and are realized by contouring the radial height of the expansion chamber. This ability to use standard nozzle blocks in installations of different expansion ratios is a very significant economic advantage since it greatly reduces the cost experienced in the prior art of customizing nozzle blocks.
  • FIGS. 9 and 10 the present invention is'shown incorporated in the rotating wheel of a pure reaction turbine which has no stationary nozzles.
  • the turbine wheel 50 is mounted forrotation on a shaft 52 and it carries anannular array of spaced buckets 54.
  • the buckets 54 arerotated in wheel 50 at their inner ends, and the outer ends of the buckets are joined to a common annular shroud 56 which forms part of the rotating assembly.
  • Shaft 52 and shroud 56 are sealed by appropriate seals 58 and 60, respectively, so high pressure steam or other motive fluid can be delivered to and expanded through the buckets in the flow direction with respect to the previous embodiments, it will be understood that the buckets may terminate at the throat or may extend any desired distance into the common expansion chamber.
  • FIGS. :9 has a common expansion chamber; i.e., chamber 70, into which all of the buckets discharge.
  • a common expansion chamber i.e., chamber 70
  • the walls 64 and 66 which define common expansion chamber 70 are inclined with respect to a radius of a turbine wheel so as to define a radially expanding chamber wherein the supersonic expansion oocurs in a direction which can be considered radial with respect to the turbine wheel.
  • the common expansion chamber is an annular chamber around the entire turbine wheel, only part of which has been shown in FIGS. 9 and 10.
  • the turbine wheel may be a pure reaction turbine having no stationary nozzles, as shown in FIGS. 9 and 10, or it may also be used in conjunction with a nozzle assembly of either the vaneless supersonic nozzle type of the present invention or a conventional prior art nozzle configuration. While both of walls 64 and 66 are shown diverging to form chamber 70, it will be understood, as discussed previously, that the radial divergence of chamber 70 can also be realized with only one of these walls inclined.
  • FIG. 1 l a modification of FIG. 10 is shown wherein backup walls 71 are incorporated. These backup walls are comparable to the backup walls described above with regard to FIG. 6.
  • the backup walls which define vaneless supersonic expansion chambers therebetween, may be distributed periodically or aperiodically about the circumference of chamber 70, with the number of backup walls varying from as few as two for the entire circumference to as many as one for each two buckets.
  • Two of the backup walls 70 are shown in FIG. l l representative of an arrangement wherein a backup wall is positioned at every third nozzle. In that arrangement for backup walls, each two adjacent backup walls would define a common radially diverging expansion chamber for the three flow paths 62 contained therebetween.
  • FIG. 12 an arrangement is shown wherein the blades 54 cooperate with a stationary shroud 72, the wall 74 of which cooperates with wall 64 to define the contour of flow passage through the buckets and also to define the contour of the common supersonic expansion chamber 70.
  • the left portion of wall 74 is inclined toward or converging with respect to wall 64
  • the right portion of wall 74 is inclined away from or diverging with respect to wall 64.
  • This converging-diverging shape of wall 74 can be used to define the throat 68 of the passageway at any desired point by adjusting the contour of wall 74 (or wall 64) to form the narrowest cross-sectional flow area of the passage at the desired location of the throat.
  • the location of the throat of the bucket or nozzle can be defined in the radial plane, i.e., radially with respect to the turbine wheel, by contouring either the top or bottom wall which defines the flow passage rather than by the contour of the side walls.
  • this modification affords the advantage that alarge number of buckets or nozzles may, if desired, be arranged in an array with a smaller than usual spacing therebetween, and the additional necessary space for the flow passage volume can be obtained by the radial contouring of the co verging partof the passage wherein the subsonic expansion takes place.
  • the turbomachinery unit of FIG. 13 has an inlet generally designated at 76, a first rotating stage 78, a stationary stage 80, a second rotating stage 82, and an exit 84.
  • the rotating stages 78 and 82 may be of the type shown in FIG.
  • each of the wheels carrying buckets 54 and cooperating with contoured shrouds 72 which serve to define the supersonic expansion passages 70 and which may, if desired, also define the contour of the subsonic expansion passages.
  • stationary stage 80 has stationary nozzles and the common supersonic expansion chamber 26. Backup walls may be included as indicated in the dotted lines.
  • a fluid entering inlet 76 at 1,000 psia can be expanded through rotating stage 78 to 100 psia, and can then be expanded through stationary stage 80 to 10 psia and can then be expanded through rotating stage 82 to l psia.
  • a 10:1 expansion can be accommodated in each of three stages to accommodate an overall pressure ratio of 1,000: l and this can be accomplished using only two rotating rows and one stationary row.
  • FIG. 13 is for illustrative purposes only, and many other and varied arrangements can be accomplished depending on the design requirements for particular installations.
  • a common diverging expansion chamber connected L to receive flow from said converging passages, said common diverging chamber extending from the vicinity of the throats of said converging passages downstream toward the turbine wheel.
  • a converging-diverging nozzle unit as in claim 1 including:
  • said backup wall is inclined at an angle to reduce the angle between the flow axis of a discharge stream from a flow passage and the exit plane of the passage upon an increase in the pressure ratio across the nozzle unit.
  • a converging-diverging nozzle unit as in claim 8 including:
  • any two adjacent backup walls being spaced apart to include at least two flow passages therebetween.
  • said common diverging expansion chamber diverges in a second direction with respect to said turbine wheel and at an angle with respect to said first direction.
  • said second direction is at an angle of approximately 90 with respect to said first direction.
  • nozzle means arranged in an arcuate array with converging flow passages defined between adjacent nozzle means for subsonic expansion of a motive fluid
  • common expansion chamber means between said array of nozzle means and the turbine wheel for su- 13 14 personic expansion of motive fluid from said nozzle vergence of said converging chamber intersect at means and delivery to the turbine wheel, said coman angle of approximately 90.
  • mon expansion chamber means having inner and 16.
  • a converging-diverging nozzle unit as in claim 12 outer arcuate walls spaced apart in a direction subwherein: stantially radial with respect to the turbine wheel, adjacent nozzles extend downstream of the throat deat least one of said inner and outer arcuate walls fined therebetween and extend into said common being inclined away from the other along at least diverging chamber. part of its length whereby said common expansion 17.
  • a converging-diverging nozzle unit as in claim 12 chamber diverges between said array of nozzle including: means and said turbine wheel. at least one backup wall in said common expansion 13.
  • a converging-diverging nozzle unit as in claim 13 sage upon an increase in the pressure ratio across wherein: the nozzle unit.
  • a converging-diverging nozzle unit as in claim 18 generally perpendicular with respect to a radius of 20 including:

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US00362403A 1973-05-21 1973-05-21 Vaneless supersonic nozzle Expired - Lifetime US3804335A (en)

Priority Applications (9)

Application Number Priority Date Filing Date Title
US00362403A US3804335A (en) 1973-05-21 1973-05-21 Vaneless supersonic nozzle
CA195,114A CA999527A (en) 1973-05-21 1974-03-15 Supersonic nozzle unit
AU67072/74A AU6707274A (en) 1973-05-21 1974-03-25 Vaneless supersonic nozzle
GB1408874A GB1457634A (en) 1973-05-21 1974-03-29 Converging-diverging supersonic nozzles
NL7404887A NL7404887A (it) 1973-05-21 1974-04-10
IT50338/74A IT1004228B (it) 1973-05-21 1974-04-11 Gruppo effusore supersonico convergente divergente
DE19742417880 DE2417880A1 (de) 1973-05-21 1974-04-11 Konvergierende/divergierende dueseneinheit zur erzeugung und expansion einer ueberschallstroemung
JP49043232A JPS508906A (it) 1973-05-21 1974-04-16
US05/649,478 USRE29128E (en) 1973-05-21 1976-01-15 Vaneless supersonic nozzle

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Application Number Priority Date Filing Date Title
US00362403A US3804335A (en) 1973-05-21 1973-05-21 Vaneless supersonic nozzle

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US05/649,478 Reissue USRE29128E (en) 1973-05-21 1976-01-15 Vaneless supersonic nozzle

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US3804335A true US3804335A (en) 1974-04-16

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US00362403A Expired - Lifetime US3804335A (en) 1973-05-21 1973-05-21 Vaneless supersonic nozzle

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US (1) US3804335A (it)
JP (1) JPS508906A (it)
AU (1) AU6707274A (it)
CA (1) CA999527A (it)
DE (1) DE2417880A1 (it)
GB (1) GB1457634A (it)
IT (1) IT1004228B (it)
NL (1) NL7404887A (it)

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3956887A (en) * 1973-11-15 1976-05-18 Rolls-Royce (1971) Limited Gas turbine engines
US3968935A (en) * 1973-05-21 1976-07-13 Sohre John S Contoured supersonic nozzle
US4027996A (en) * 1974-07-22 1977-06-07 Kraftwerk Union Aktiengesellschaft Turbomachine, such as a steam turbine with high steam inlet temperature, especially
USRE30720E (en) * 1978-07-12 1981-08-25 Contoured supersonic nozzle
US4336039A (en) * 1977-10-13 1982-06-22 Sohre John S Geothermal turbine
WO1982002418A1 (fr) * 1981-01-05 1982-07-22 Bessay Raymond Etage de turbine
EP0942150A3 (en) * 1998-03-11 2000-12-20 Rolls-Royce Plc A stator vane assembly for a turbomachine
US6196793B1 (en) * 1999-01-11 2001-03-06 General Electric Company Nozzle box
US20030194313A1 (en) * 1999-11-25 2003-10-16 Delta Electronics, Inc. Serial fan with a plurality of rotor vanes
US20040184916A1 (en) * 2003-01-30 2004-09-23 Bernhard Schmitt Turbine wheel for driving rapidly rotating tools
US20040219012A1 (en) * 2003-04-17 2004-11-04 Alstom Technology Ltd Steam turbine including a high-pressure body having a single stage of blades
EP1790826A1 (en) * 2005-11-24 2007-05-30 Siemens Aktiengesellschaft Turbine vane for a turbine of a thermal power plant
US20100077719A1 (en) * 2008-09-29 2010-04-01 Siemens Energy, Inc. Modular Transvane Assembly
US20110203282A1 (en) * 2008-09-29 2011-08-25 Charron Richard C Assembly for directing combustion gas
WO2011085835A3 (de) * 2009-12-22 2011-10-27 Robert Bosch Gmbh Lavaldüse
EP2407247A1 (de) * 2002-07-22 2012-01-18 Dürr Systems GmbH Turbinenmotor eines Rotationszerstäubers
US20140017071A1 (en) * 2012-07-11 2014-01-16 Alstom Technology Ltd Static vane assembly for an axial flow turbine
EP2851508A3 (de) * 2013-09-20 2015-04-01 Mahle International GmbH Lavaldüse
US9322335B2 (en) 2013-03-15 2016-04-26 Siemens Energy, Inc. Gas turbine combustor exit piece with hinged connections
US20180252112A1 (en) * 2017-03-01 2018-09-06 Panasonic Corporation Turbine nozzle and radial turbine including the same
BE1028108B1 (fr) * 2020-02-28 2021-09-28 Safran Aero Boosters Compresseur transsonique de turbomachine

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JPS5314205A (en) * 1976-07-23 1978-02-08 Hitachi Ltd Step structure for axial-flow fluid machine
JPS58152106A (ja) * 1982-03-05 1983-09-09 Nissan Motor Co Ltd 軸流タ−ビン
JPH01259607A (ja) * 1988-04-08 1989-10-17 Sharp Corp ホワイトノイズ波形発生回路
JP6866187B2 (ja) * 2017-03-01 2021-04-28 パナソニック株式会社 タービンノズル及びそれを備えたラジアルタービン
JP6826574B2 (ja) * 2018-09-26 2021-02-03 三菱重工マリンマシナリ株式会社 部分送入タービン

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GB342289A (en) * 1929-10-18 1931-01-19 Gen Electric Co Ltd Improvements in elastic fluid turbines
US2378372A (en) * 1937-12-15 1945-06-12 Whittle Frank Turbine and compressor
US2258793A (en) * 1940-03-19 1941-10-14 Westinghouse Electric & Mfg Co Elastic-fluid turbine
GB615219A (en) * 1945-12-13 1949-01-04 Power Jets Res & Dev Ltd Improvements relating to rotary power conversion machines
US2702157A (en) * 1949-09-28 1955-02-15 Edward A Stalker Compressor employing radial diffusion
US2839239A (en) * 1954-06-02 1958-06-17 Edward A Stalker Supersonic axial flow compressors
US2953295A (en) * 1954-10-22 1960-09-20 Edward A Stalker Supersonic compressor with axially transverse discharge
US3744724A (en) * 1972-03-24 1973-07-10 Sulzer Ag Air distributing channel

Cited By (33)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3968935A (en) * 1973-05-21 1976-07-13 Sohre John S Contoured supersonic nozzle
US3956887A (en) * 1973-11-15 1976-05-18 Rolls-Royce (1971) Limited Gas turbine engines
US4027996A (en) * 1974-07-22 1977-06-07 Kraftwerk Union Aktiengesellschaft Turbomachine, such as a steam turbine with high steam inlet temperature, especially
US4336039A (en) * 1977-10-13 1982-06-22 Sohre John S Geothermal turbine
USRE30720E (en) * 1978-07-12 1981-08-25 Contoured supersonic nozzle
WO1982002418A1 (fr) * 1981-01-05 1982-07-22 Bessay Raymond Etage de turbine
EP0942150A3 (en) * 1998-03-11 2000-12-20 Rolls-Royce Plc A stator vane assembly for a turbomachine
US6196793B1 (en) * 1999-01-11 2001-03-06 General Electric Company Nozzle box
US20030194313A1 (en) * 1999-11-25 2003-10-16 Delta Electronics, Inc. Serial fan with a plurality of rotor vanes
US20070128023A1 (en) * 1999-11-25 2007-06-07 Kuo-Cheng Lin Serial fan with a plurality of rotor vanes
US7238004B2 (en) * 1999-11-25 2007-07-03 Delta Electronics, Inc. Serial fan with a plurality of rotor vanes
US7740446B2 (en) 1999-11-25 2010-06-22 Delta Electronics, Inc. Serial fan with a plurality of rotor vanes
EP2407247A1 (de) * 2002-07-22 2012-01-18 Dürr Systems GmbH Turbinenmotor eines Rotationszerstäubers
US20040184916A1 (en) * 2003-01-30 2004-09-23 Bernhard Schmitt Turbine wheel for driving rapidly rotating tools
US7223075B2 (en) 2003-01-30 2007-05-29 GAT Gesellschaft für Antriebstechnik mbH Turbine wheel for driving rapidly rotating tools
US20040219012A1 (en) * 2003-04-17 2004-11-04 Alstom Technology Ltd Steam turbine including a high-pressure body having a single stage of blades
US7056086B2 (en) * 2003-04-17 2006-06-06 Alstom Technology Ltd Steam turbine including a high-pressure body having a single state of blades
EP1790826A1 (en) * 2005-11-24 2007-05-30 Siemens Aktiengesellschaft Turbine vane for a turbine of a thermal power plant
US20100077719A1 (en) * 2008-09-29 2010-04-01 Siemens Energy, Inc. Modular Transvane Assembly
US8276389B2 (en) 2008-09-29 2012-10-02 Siemens Energy, Inc. Assembly for directing combustion gas
US20110203282A1 (en) * 2008-09-29 2011-08-25 Charron Richard C Assembly for directing combustion gas
US8230688B2 (en) 2008-09-29 2012-07-31 Siemens Energy, Inc. Modular transvane assembly
US9206691B2 (en) 2009-12-22 2015-12-08 Robert Bosch Gmbh Laval nozzle
WO2011085835A3 (de) * 2009-12-22 2011-10-27 Robert Bosch Gmbh Lavaldüse
US20140017071A1 (en) * 2012-07-11 2014-01-16 Alstom Technology Ltd Static vane assembly for an axial flow turbine
US9316107B2 (en) * 2012-07-11 2016-04-19 Alstom Technology Ltd Static vane assembly for an axial flow turbine
US9322335B2 (en) 2013-03-15 2016-04-26 Siemens Energy, Inc. Gas turbine combustor exit piece with hinged connections
EP2851508A3 (de) * 2013-09-20 2015-04-01 Mahle International GmbH Lavaldüse
US20180252112A1 (en) * 2017-03-01 2018-09-06 Panasonic Corporation Turbine nozzle and radial turbine including the same
CN108533332A (zh) * 2017-03-01 2018-09-14 松下电器产业株式会社 涡轮喷嘴以及具备涡轮喷嘴的径流式涡轮机
US10677077B2 (en) * 2017-03-01 2020-06-09 Panasonic Corporation Turbine nozzle and radial turbine including the same
CN108533332B (zh) * 2017-03-01 2022-03-08 松下电器产业株式会社 涡轮喷嘴以及具备涡轮喷嘴的径流式涡轮机
BE1028108B1 (fr) * 2020-02-28 2021-09-28 Safran Aero Boosters Compresseur transsonique de turbomachine

Also Published As

Publication number Publication date
CA999527A (en) 1976-11-09
AU6707274A (en) 1975-09-25
DE2417880A1 (de) 1974-12-12
NL7404887A (it) 1974-11-25
GB1457634A (en) 1976-12-08
IT1004228B (it) 1976-07-10
JPS508906A (it) 1975-01-29

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