US3871791A - Blade for fluid flow machines - Google Patents

Blade for fluid flow machines Download PDF

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Publication number
US3871791A
US3871791A US333659A US33365973A US3871791A US 3871791 A US3871791 A US 3871791A US 333659 A US333659 A US 333659A US 33365973 A US33365973 A US 33365973A US 3871791 A US3871791 A US 3871791A
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US
United States
Prior art keywords
shank
aerofoil
inner end
radially
disc
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US333659A
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English (en)
Inventor
Kenneth Ronald Guy
Robert Burns Hood
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
Original Assignee
Rolls Royce 1971 Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce 1971 Ltd filed Critical Rolls Royce 1971 Ltd
Application granted granted Critical
Publication of US3871791A publication Critical patent/US3871791A/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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the blade has in succession an aerofoil, a shank and a root, and the shape of the aerofoil is continued into the shank but at progressively reducing camber so that, whereas adjacent the aerofoil the camber of the shank is the same as that of the aero-foil, adjacent the root the camber of the shank is zero.
  • Skill 1 BF 2 BLADE FOR FLUID FLOW MACHINES This invention relates to blades for fluid flow machines.
  • lt is known for such blades to comprise an aerofoil, a shank and a root.
  • the aerofoil manifests the stagger, camber and cross-sectional shape required by the aerodynamics of the blade.
  • the root is a portion of the blade shaped to engage a recess in a rotor body thereby to support the blade inter alia against centrifugal force.
  • the shank is a portion of the blade connecting the root to the aerofoil and providing a degree of flexibility between root and aerofoil.
  • the shank is still of substantially greater cross-section than the aerofoil so that the junction of aerofoil and shank is a locality of relatively high stress concentration.
  • the stagger of the aerofoil is of progressively reducing magnitude in the radially inward direction, this change of stagger being referred to as twist.”
  • the twist of the blade gives rise to a torsion couple in the sense tending to reduce the twist.
  • the torsion couple gives rise to a diminution in the centrifugal stress at the leading and trailing edges of the aerofoil and an increase in the centrifugal stress at the medial parts of the aerofoil crosssection. The stresses so produced tend to produce outof-plane deformation of the cross-section.
  • a blade for a fluid flow machine comprising in succession an aerofoil, a shank and a root, wherein the shank has at its radially outer end a stagger, camber and crosssectional shape similar to that of the adjacent end of the aerofoil, and wherein the shank is shaped for the camber thereof to be progressively reducing towards the root end of the shank.
  • the reduction of the camber leads to a corresponding reduction in warping restraint stresses, the latter now occurring at the root end of the shank.
  • the shank is made sufficiently long to make it possible to reduce the camber to zero at the root end thereof. Thereby the warping restraint stresses are reduced to a minimum.
  • the stagger of the shank may also be reduced progressively towards the root end of the shank to make it possible to improve the stress distribution in the shank and to reduce stress concentrations at the root recess in the rotor body.
  • FIG. 1 is a side elevation of a fan for a gas turbine engtne emobodymg a blades according to the invention.
  • FIG. 2 is a view in the direction of arrow II in FIG. 1.
  • FIG. 3 is a plan view of FIG. 2.
  • FIG. 4 is a section on the line C-C in FIG. 1 but having the same orientation as FIG. 3.
  • FIG. 5 is a view similar to FIG. 2 but embodying a modification.
  • FIG. 6 is a view similar to FIG. 4 and embodying the modification shown in FIG. 5.
  • a rotor 10 comprising a disc 11 and blades 12 (only one shown). Each blade comprises in succession an aerofoil 13, a shank l4 and a root 15. At the junction between the shank and the aerofoil the blade comprises a platform 16 which is a wall part of the fluid flow passage controlled by the aerofoil.
  • the blade is connected to the disc by interdigitation (FIG. 2) between the rootand a recess 18 in the disc.
  • the axis of rotation of the rotor is denoted 19 (FIGS. 3,4).
  • the direction of rotation of the blade is shown by an arrow 20 (FIGS. 2,3).
  • a line 21 is the locus of the centres of gravity of successive cross-sections of the blade transverse to the line 21.
  • the stag ger is defined by an angle a between the chord line, 22, of the section and the axis 19.
  • the camber is defined as the curvature (i.e., rate of change of slope) of the socalled camber line 23A of the section.
  • the shape of the section is given by the contour 24A of the section as shown.
  • the stagger of the aerofoil progressively reduces between the sections Al, B1, the stagger of the latter section being given by an angle B.
  • the change in stagger defines the twist of the blade.
  • the camber of section B1 is greater than that of section A1 and is given by a line 238.
  • the shape of section BI is given by the contour 248.
  • the radially outermost end, i.e., the section C1 of the shank is formed (FIG. 4) to have a stagger angle B, camber line 23C and cross-sectional shape 24C similar to those of the adjacent end, i.e., the section B], of the aerofoil, and the shank is shaped so that the camber of successive sections thereof is of progressively reducing magnitude.
  • the camber of the radially in nermost section D1 ofthe shank is zero, i.e., the section is symmetrical about a line 25.
  • the line 25 is also the axis of symmetry of the root 15.
  • a shank shaped in this way has leading and trailing edges 26 and 27 respectively which extend from the section Cl, where the edges 26, 27 are remote from the plane, referred to as the plane of symmetry of the root 15, of the lines 21, 25 to a position at the section D1 where the edges 26, 27 intersect that plane.
  • the edges 26, 27 are curved to have a common tangent with leading and trailing edges 28 and 29 respectively of the aerofoil so that good continuity of form is preserved at the junction of aerofoil and shank with a view of minimising stress concentrations.
  • the warping restraint stresses occur at the junction of shank and root i.e., at the section D1.
  • the warping restraint stresses are therefore only those created by the twist of the blade itself.
  • FIGS. 5 and 6 show a blade in which the stagger of the root is zero, i.e. the root has a plane of symmetry aligned to the axis 19 (FIG. 6).
  • the curvature of the edges 26,27 is now more nearly the same, the curvature of the trailing edge 27 being less acute at the expense of the curvature of the leading edge 26 being tolerably more acute.
  • a fan rotor for gas turbine engines comprising:
  • each of said blades comprising an aerofoil of given camber and twist extending radially outwardly of said wall and having a radially inner end adjacent said wall, the chord of said aerofoil generally increasing radially outwardly,
  • shank extending between the inner end of the aerofoil and a location adjacent the periphery of the disc, said shank having a radially outer end of a cross-section similar in stagger, camber and aerofoil shape to that of the inner end of the aerofoil, and the camber of the shank gradually progressively decreasing from the outer end to the inner end thereof so as to be zero at the inner end of the shank,
  • a fan rotor for gas turbine engines comprising a root portion directly adjacent the radially inner end of the shank and radially inwardly thereof, and the disc including, with respect to each root portion, a recess engageable by the root portion to connect the blade to the disc.
  • a fan rotor for gas turbine engines wherein the root portion has a plane of symmetry, the shank and aerofoil each having a trailing edge, and the trailing edge of the shank, at the radially outer end thereof, has a common tangent with the trailing edge of the aerofoil at the radially inner end thereof, and the trailing edge of the shank intersects the plane of symmetry of the root portion at the radially inner end of the shank.
  • a fan rotor for gas turbine engines having a leading edge, the leading edge of the shank at the radially outer end thereof has a common tangent with the leading edge of the aerofoil at the radially inner end thereof, and the leading edge of the shank intersects the plane of symmetry of the root portion at the radially inner end of the shank.

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US333659A 1972-03-09 1973-02-20 Blade for fluid flow machines Expired - Lifetime US3871791A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB1107872A GB1419381A (en) 1972-03-09 1972-03-09 Fan for gas turbine engines

Publications (1)

Publication Number Publication Date
US3871791A true US3871791A (en) 1975-03-18

Family

ID=9979625

Family Applications (1)

Application Number Title Priority Date Filing Date
US333659A Expired - Lifetime US3871791A (en) 1972-03-09 1973-02-20 Blade for fluid flow machines

Country Status (6)

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US (1) US3871791A (US20050271598A1-20051208-C00001.png)
JP (1) JPS5430123B2 (US20050271598A1-20051208-C00001.png)
DE (1) DE2309404C2 (US20050271598A1-20051208-C00001.png)
FR (1) FR2175429A5 (US20050271598A1-20051208-C00001.png)
GB (1) GB1419381A (US20050271598A1-20051208-C00001.png)
IT (1) IT977572B (US20050271598A1-20051208-C00001.png)

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4120607A (en) * 1976-03-26 1978-10-17 Rolls-Royce Limited Rotor blade for a gas turbine engine
US4326836A (en) * 1979-12-13 1982-04-27 United Technologies Corporation Shroud for a rotor blade
US4595340A (en) * 1984-07-30 1986-06-17 General Electric Company Gas turbine bladed disk assembly
US4682935A (en) * 1983-12-12 1987-07-28 General Electric Company Bowed turbine blade
US4957411A (en) * 1987-05-13 1990-09-18 Societe Nationale D'etude Et De Construction De Moteurs D'aviaton S.N.E.C.M.A. Turbojet engine with fan rotor blades having tip clearance
US5044885A (en) * 1989-03-01 1991-09-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Mobile blade for gas turbine engines providing compensation for bending moments
US5067876A (en) * 1990-03-29 1991-11-26 General Electric Company Gas turbine bladed disk
US5108261A (en) * 1991-07-11 1992-04-28 United Technologies Corporation Compressor disk assembly
US5435694A (en) * 1993-11-19 1995-07-25 General Electric Company Stress relieving mount for an axial blade
US5480284A (en) * 1993-12-20 1996-01-02 General Electric Company Self bleeding rotor blade
US6299412B1 (en) * 1999-12-06 2001-10-09 General Electric Company Bowed compressor airfoil
US6375419B1 (en) 1995-06-02 2002-04-23 United Technologies Corporation Flow directing element for a turbine engine
US10598033B2 (en) 2014-09-08 2020-03-24 Safran Aircraft Engines Vane with spoiler

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4621979A (en) * 1979-11-30 1986-11-11 United Technologies Corporation Fan rotor blades of turbofan engines

Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1772876A (en) * 1927-11-17 1930-08-12 Parsons Billet or blank suitable for the production of turbine blades
US1981392A (en) * 1932-12-03 1934-11-20 Manganese Bronze & Brass Compa Propeller and the like
US2193616A (en) * 1937-07-10 1940-03-12 Baumann Werner Screw propeller
US2327453A (en) * 1941-10-04 1943-08-24 Eric A F Presser Helical marine propeller
US2391623A (en) * 1943-12-08 1945-12-25 Armstrong Siddeley Motors Ltd Bladed rotor
US2421890A (en) * 1944-11-27 1947-06-10 Goetaverken Ab Turbine blade
US2974728A (en) * 1957-10-21 1961-03-14 Lennox Ind Inc Fan construction
US2999668A (en) * 1958-08-28 1961-09-12 Curtiss Wright Corp Self-balanced rotor blade
US3012709A (en) * 1955-05-18 1961-12-12 Daimler Benz Ag Blade for axial compressors
US3173490A (en) * 1962-07-25 1965-03-16 Hiller Aircraft Company Inc Propeller blade for vtol aircraft
US3477795A (en) * 1967-05-01 1969-11-11 Rolls Royce Bladed rotor for a fluid flow machine
US3490852A (en) * 1967-12-21 1970-01-20 Gen Electric Gas turbine rotor bucket cooling and sealing arrangement
US3576377A (en) * 1967-12-22 1971-04-27 Rolls Royce Blades for fluid flow machines

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3294364A (en) * 1962-01-02 1966-12-27 Gen Electric Rotor assembly
US3334685A (en) * 1965-08-18 1967-08-08 Gen Electric Fluid boiling and condensing heat transfer system

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1772876A (en) * 1927-11-17 1930-08-12 Parsons Billet or blank suitable for the production of turbine blades
US1981392A (en) * 1932-12-03 1934-11-20 Manganese Bronze & Brass Compa Propeller and the like
US2193616A (en) * 1937-07-10 1940-03-12 Baumann Werner Screw propeller
US2327453A (en) * 1941-10-04 1943-08-24 Eric A F Presser Helical marine propeller
US2391623A (en) * 1943-12-08 1945-12-25 Armstrong Siddeley Motors Ltd Bladed rotor
US2421890A (en) * 1944-11-27 1947-06-10 Goetaverken Ab Turbine blade
US3012709A (en) * 1955-05-18 1961-12-12 Daimler Benz Ag Blade for axial compressors
US2974728A (en) * 1957-10-21 1961-03-14 Lennox Ind Inc Fan construction
US2999668A (en) * 1958-08-28 1961-09-12 Curtiss Wright Corp Self-balanced rotor blade
US3173490A (en) * 1962-07-25 1965-03-16 Hiller Aircraft Company Inc Propeller blade for vtol aircraft
US3477795A (en) * 1967-05-01 1969-11-11 Rolls Royce Bladed rotor for a fluid flow machine
US3490852A (en) * 1967-12-21 1970-01-20 Gen Electric Gas turbine rotor bucket cooling and sealing arrangement
US3576377A (en) * 1967-12-22 1971-04-27 Rolls Royce Blades for fluid flow machines

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4120607A (en) * 1976-03-26 1978-10-17 Rolls-Royce Limited Rotor blade for a gas turbine engine
US4326836A (en) * 1979-12-13 1982-04-27 United Technologies Corporation Shroud for a rotor blade
US4682935A (en) * 1983-12-12 1987-07-28 General Electric Company Bowed turbine blade
US4595340A (en) * 1984-07-30 1986-06-17 General Electric Company Gas turbine bladed disk assembly
US4957411A (en) * 1987-05-13 1990-09-18 Societe Nationale D'etude Et De Construction De Moteurs D'aviaton S.N.E.C.M.A. Turbojet engine with fan rotor blades having tip clearance
US5044885A (en) * 1989-03-01 1991-09-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Mobile blade for gas turbine engines providing compensation for bending moments
US5067876A (en) * 1990-03-29 1991-11-26 General Electric Company Gas turbine bladed disk
US5108261A (en) * 1991-07-11 1992-04-28 United Technologies Corporation Compressor disk assembly
US5435694A (en) * 1993-11-19 1995-07-25 General Electric Company Stress relieving mount for an axial blade
US5480284A (en) * 1993-12-20 1996-01-02 General Electric Company Self bleeding rotor blade
US6375419B1 (en) 1995-06-02 2002-04-23 United Technologies Corporation Flow directing element for a turbine engine
US6299412B1 (en) * 1999-12-06 2001-10-09 General Electric Company Bowed compressor airfoil
US10598033B2 (en) 2014-09-08 2020-03-24 Safran Aircraft Engines Vane with spoiler

Also Published As

Publication number Publication date
GB1419381A (en) 1975-12-31
FR2175429A5 (US20050271598A1-20051208-C00001.png) 1973-10-19
IT977572B (it) 1974-09-20
DE2309404A1 (de) 1973-09-13
JPS5430123B2 (US20050271598A1-20051208-C00001.png) 1979-09-28
JPS48101603A (US20050271598A1-20051208-C00001.png) 1973-12-21
DE2309404C2 (de) 1982-07-22

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