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Publication numberUS3871791 A
Publication typeGrant
Publication dateMar 18, 1975
Filing dateFeb 20, 1973
Priority dateMar 9, 1972
Also published asDE2309404A1, DE2309404C2
Publication numberUS 3871791 A, US 3871791A, US-A-3871791, US3871791 A, US3871791A
InventorsGuy Kenneth Ronald, Hood Robert Burns
Original AssigneeRolls Royce 1971 Ltd
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Blade for fluid flow machines
US 3871791 A
Abstract
The disclosure of this invention relates to a rotor blade for a fan for a gas turbine engine. 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. This leads to a reduction in the stresses produced by a torsion couple arising in operation and acting in the sense tending to reduce the twist of the blade.
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Description  (OCR text may contain errors)

[ Mar. 18, 1975 BLADE FOR FLUID FLOW MACHINES [75] Inventors: Kenneth Ronald Guy; Robert Burns Hood, both of Bristol, England [73] Assignee: Rolls-Royce (1971) Limited,

London, England 22 Filed: Feb. 20, 1973 21 App1.No.:333,659

[30] Foreign Application Priority Data Mar. 9. 1972 Great Britain 11078/72 [521 US. Cl 416/193, 416/196, 416/234, 416/242 [51] Int. Cl. FOld 5/14, FOld 5/22 [58] Field of Search 416/193, 223, 239, 234, 416/242, 248, 196, 191,236 A [56] References Cited UNITED STATES PATENTS 1,772,876 8/1930 Parsons et a1..... 416/234 UX 1,981,392 11/1934 Selman 416/223 2,193,616 3/1940 Baumann.. 416/236 X 2.327.453 8/1943 Presser 416/239 X 2.391.623 12/1945 Heppner i 416/191 2.421.890 6/1947 .lohansson 416/193 2,974,728 3/1961 Culp 416/239 2.999.668 9/1961 Howald et a1 416/97 3.012.709 12/1961 Schne11 416/236v 3,173,490 3/1965 Stuart 416/223 3,477,795 11/1969 Beesley 416/193 X 3,490,852 l/l970 Carlstrom et a1. 416/95 3,576,377 4/1971 Beanland et a1 416/191 FOREIGN PATENTS OR APPLlCATIONS 932,045 11/1947 France 416/95 1,263,677 5/1961 France 416/193 1,330,690 5/1963 France 416/223 229,266 10/1943 Switzerland 416/234 8,524 2/1965 Japan 416/193 1,187,166 3/1959 France 416/236 A Primary ExaminerEverette A. Powell, Jr. Attorney, Agent, or Firm-Stevens, Davis, Miller & Mosher [57] ABSTRACT The disclosure of this invention relates to a rotor blade for a fan for a gas turbine engine. 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. This leads to a reduction in the stresses produced by a torsion couple arising in operation and acting in the sense tending to reduce the twist of the blade.

4 Claims, 6 Drawing Figures pmmggum 8197-5 9.871. .791

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. However, in known blades 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.

For aerodynamic reasons the stagger of the aerofoil is of progressively reducing magnitude in the radially inward direction, this change of stagger being referred to as twist." Under centrifugal force the twist of the blade gives rise to a torsion couple in the sense tending to reduce the twist. At sections of the aerofoil remote from any end restraint 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. At the junction between aerofoil and shank the relatively greater stiffness of the latter restrains the out-of-plane deformation of the adjacent aerofoil and gives rise to socalled warping restraint stresses. It can be shown that these stresses increase with increasing camber.

In relatively highly cambered and twisted rotor blades for aero-engine fans the warping restraint stresses dominate the stress pattern at the inner end of the aerofoil and this can be critical for the design of the blade. It is an object of the invention to provide a construction of blade in which the warping restraint stresses are reduced.

According to the invention there is provided 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. Preferably 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.

Examples of blades according to this invention will now be described with reference to the accompanying drawings wherein:

FIG. 1 is a side elevation ofa 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.

Referring to FIGS. 1 to 4, there is shown 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 cross-sections of the blade at lines AA, B-B, CC, D-D (FIG. 1) are denoted A1, B1, C1, D1 respectively in FIGS. 2, 3,4.

The change in cross section between the radially inner and outer ends of the aerofoil is determined in accordance with known principles and need not be discussed in detail, but attention is drawn to the change in stagger and to the presence of camber.

Taking section A1 as an example (FIG. 2), 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.

In the past it has been the practice to make the crosssection of the shank substantially larger than that of the radially innermost section B1 of the aerofoil, and the cross-section of the shank was more nearly the same as the parallelogrammic shape of the root as shown in FIG. 3. In accordance with the invention, 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. As shown in FIG. 4 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.

As shown in FIGS. 1, 4 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.

Similarity between the sections B1, C l is affected by the angles at which the edges 28, 29 enter the platform 16 and by the thickness of the platform. This results, in this example, in that the section Cl is of shorter span and of lesser stagger angle that the section B1, and the term similar applied to thesections B1, C1 is intended to include these differences.

In use, when the blade is subject to centrifugal force, the warping restraint stresses occur at the junction of shank and root i.e., at the section D1. By virtue of the absence of any camber at the section D1 the warping restraint stresses are therefore only those created by the twist of the blade itself.

It will be noted from FIG. 2 that in order to preserve continuity of line between the edges 29, 27, the latter has to be curved fairly acutely, and in the opposite sense to that of the edge 29, in order to fit the condition of having at its one end a common tangent with the edge 29 and at its other end lying on line 25. This acute curvature can be a disadvantage if the shank is to be relatively short and the twist and camber are relatively high. This situation can be avoided by changing the stagger of the root in the sense of reducing it, i.e., bringing the root more nearly into alignment with the axis 19.

The result of such a change in root stagger are shown in FIGS. 5 and 6 which 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. This results not only in a better stress balance in the shank but also makes it possible for the recess 18 in the disc to be parallel to the axis 19. This simplifies the machining of the recess and avoids the stress concentrations encountered in recesses having a pronounced stagger.

We claim:

I. A fan rotor for gas turbine engines comprising:

a disc having an axis of rotation,

fan blades connected to the disc and extending radially from the periphery thereof,

means defining a wall extending circumferentially about said axis in spaced apart relationship to the periphery of the disc,

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,

a 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,

and means provided at the radially inner end of the shank for connecting the blade to the disc.

2. A fan rotor for gas turbine engines according to claim 1, said connecting means 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.

3. A fan rotor for gas turbine engines according to claim 2, 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.

4. A fan rotor for gas turbine engines according to claim 3, the shank and aerofoil each 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.

Patent Citations
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Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US4120607 *Mar 16, 1977Oct 17, 1978Rolls-Royce LimitedRotor blade for a gas turbine engine
US4326836 *Dec 13, 1979Apr 27, 1982United Technologies CorporationShroud for a rotor blade
US4595340 *Jul 30, 1984Jun 17, 1986General Electric CompanyGas turbine bladed disk assembly
US4682935 *Dec 12, 1983Jul 28, 1987General Electric CompanyBowed turbine blade
US4957411 *May 11, 1988Sep 18, 1990Societe 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
US5044885 *Mar 1, 1990Sep 3, 1991Societe 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
US5067876 *Mar 29, 1990Nov 26, 1991General Electric CompanyGas turbine bladed disk
US5108261 *Jul 11, 1991Apr 28, 1992United Technologies CorporationCompressor disk assembly
US5435694 *Nov 19, 1993Jul 25, 1995General Electric CompanyStress relieving mount for an axial blade
US5480284 *Dec 20, 1993Jan 2, 1996General Electric CompanySelf bleeding rotor blade
US6299412 *Dec 6, 1999Oct 9, 2001General Electric CompanyBowed compressor airfoil
US6375419Jun 2, 1995Apr 23, 2002United Technologies CorporationFlow directing element for a turbine engine
Classifications
U.S. Classification416/193.00R, 416/193.00A, 416/196.00R, 416/242, 416/223.00A, 416/234
International ClassificationF04D29/34, F02K3/00, F01D5/14, F04D29/32, F02K3/04
Cooperative ClassificationF01D5/141
European ClassificationF01D5/14B