WO2007045815A1 - A blade mounting - Google Patents
A blade mounting Download PDFInfo
- Publication number
- WO2007045815A1 WO2007045815A1 PCT/GB2006/003544 GB2006003544W WO2007045815A1 WO 2007045815 A1 WO2007045815 A1 WO 2007045815A1 GB 2006003544 W GB2006003544 W GB 2006003544W WO 2007045815 A1 WO2007045815 A1 WO 2007045815A1
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- blade
- root
- aerofoil
- corners
- axis
- Prior art date
Links
Classifications
-
- 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
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/34—Blade mountings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/30—Arrangement of components
- F05D2250/31—Arrangement of components according to the direction of their main axis or their axis of rotation
- F05D2250/313—Arrangement of components according to the direction of their main axis or their axis of rotation the axes being perpendicular to each other
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/30—Arrangement of components
- F05D2250/31—Arrangement of components according to the direction of their main axis or their axis of rotation
- F05D2250/314—Arrangement of components according to the direction of their main axis or their axis of rotation the axes being inclined in relation to each other
Definitions
- the present invention relates to blade mountings and blades utilised in gas turbine engines. Operation of gas turbine engines is relatively well known and includes a number of aerofoil blades secured in mountings in different stages of the gas turbine engine. These blades are generally secured through root mountings which may take the form of dovetail root sections which enter a reciprocally shaped slot in order to secure the blade to a rotor disc. Normally the airfoils forming the blade are curved. In such circumstances front and rear edges of the blade root are cut to provide an orthogonal flat face for consistency with the rotor disc edge surfaces at the front and rear edges of the blade. Thus, these blade root edges include relatively sharp corners and angular parts.
- the acute corners if they impact against a casing when a fan blade fails can cause problems.
- the acute corners concentrate impact load from the relatively heavy root section of the blade upon disintegration.
- additional thickness to a fan casing adds considerably to the necessary weight of the fan casing with detrimental effects upon engine operational efficiency. It is found that a lmm increase in thickness in a large fan casing can add approximately 16kg to overall weight.
- a blade for a gas turbine engine the blade rotatable in use about an axis of rotation, the blade comprising a root for securing the blade, the root curved for alignment in use with a mounting slot, characterised in that an end of the root is substantially perpendicular to the axis of curvature of the root.
- both ends of the root may be substantially perpendicular to the axis of curvature of the root .
- a blade for a gas turbine engine the blade rotatable in use about an axis of rotation, the blade comprising a root for securing the blade, the root curved for alignment in use with a mounting slot, characterised in that an end of the root makes an angle of substantially 25 degrees to the plane perpendicular to the axis of rotation.
- both ends of the root may make an angle of substantially 25 degrees to the plane perpendicular to the axis of rotation.
- the blade may comprise an aerofoil which extends from the root and a portion of the aerofoil adjacent to the end of the root may be rounded so as to provide a smooth transition between the aerofoil and the end of the root.
- At least one end of the root may be chamfered or truncated, so as to reduce the axial length of the root.
- the blade may be part of a gas turbine engine.
- Fig. 1 is a schematic cross section of a prior blade
- Fig. 2 is a schematic cross section of a blade in accordance with the present invention
- Fig. 3 is a schematic end view of a prior blade
- Fig. 4 is a schematic end view of a blade in accordance with the present invention.
- Fig. 5 is a schematic view of a blade mounting.
- the fan blade root is generally in the form of as indicated a dovetail which locates in a corresponding reciprocal slot in the fan rotor disc. End surfaces of the root are formed so that they are parallel with the front and rear surfaces of the fan rotor disc and so orthogonal to the engine rotational axis.
- Such orthogonal presentation of the ends of the root sections for the blades creates sharp and angular portions which as indicated may concentrate imparted load upon impact with a fan casing should the blade become detached.
- Such potentially heavier impact forces require a thicker casing which in turn adds significantly to overall engine weight and therefore reduces efficiency.
- Figure 1 provides an illustration of a prior blade profile cross section.
- the blade 1 has a root 2 which incorporates angular corners 3 on the pressure side of an aerofoil 4. These acute corners as indicated previously will act through the potentially narrow impact zone of the acute point 5 of each corner 3 to impart relatively high impact loads . Such impact loads necessitate thicker and more robust casing profiles and therefore add to overall weight.
- the corners 3 are created by desire to have end faces 6 of the root 2 which are orthogonal to the axis of rotation for an engine incorporating the blade 1 in a blade assembly. It will be understood that the root end surfaces 6 will generally be continuous and aligned with front and rear faces of a fan rotor disc as will be described later.
- Fig. 2 illustrates a blade mounting arrangement in accordance with the present invention.
- a blade 21 with a similar aerofoil 24 to that described with regard to Fig. 1 is provided.
- the aerofoil 24 extends from a root 22 which is utilised to secure the blade in a mounting arrangement comprising a number of aerofoil blades 24 secured to a fan rotor disc.
- the root 22 now includes end surfaces 26 which are substantially perpendicular to an axis of curvature 27 for the blade 21 or at least turned towards that orientation.
- the ends 26 of the root 22 as indicated previously have a dovetail shaping and are relatively square ended so that the corners 23 of the roots 22 are less acute and of a smaller dimension.
- the impact forces should these corners 23 strike upon a fan casing as a result of failure of the blade 21 will be less severe.
- By such angling of the end faces 26 there is a reduction of the severity of the angle 23 in comparison with angle 3 in lever 1.
- the faces 26 will be at an angle 28 in the order of 25 degrees to the orthogonal plane, that is to say the plane 29 perpendicular to the axis of rotation for an engine incorporating the blades 21 and typically consistent with the front and rear surfaces of a mounting rotor disc in which the blades are secured through the roots 22.
- the end face 6 of the root 2 will generally be in a plane 9 which extends orthogonally upwards from the end edges 10 of the blade 1.
- the end plane 29a in Figure 2 is consistent with the plane 9 depicted in Figure 1 that is to say orthogonal to the end edges 20 of the blade aerofoil 24.
- the axial length of the blade 21 is increased.
- FIGS. 3 and 4 respectively illustrate schematically end views of a prior blade (Fig. 3) and a present blade
- a blade 31 again incorporates an aerofoil 34 extending from a root 32.
- the root 32 is of a dovetail nature and is slid into a reciprocally shaped slot in a rotor mounting disc in order to secure the blade 31 in use.
- the root 32 at a front edge 35 is generally flat and as described above creates the angular corner 35 for the blade 31 which may cause disproportionate impact damage with a fan casing should the blade 31 fail and disintegrate. It is the severity of the angle at 35 which results in acute corners with sharp points (5 in figure 1) causing high impact loads.
- a blade 41 in accordance with the present invention is illustrated with an end face 36 to the root 42.
- the blade 41 as previously incorporates an aerofoil blade 44 which develops from the root 42.
- the sharp edge 35 depicted in Fig.3 it will be noted that there is profiling of the transition between the root 42 and in particular the face 46 to the adjacent portions 43 of the aerofoil 44.
- these transition portions 43 there is less acute angling of the aerofoil 44 which in combination with the produced angular corners in the root 42 further limits the potential for sharp point impacts with a fan casing which may as indicated cause greater damage.
- the actual profile cross section in the transition portions 43 may be chosen in accordance with operation requirements and it will be understood is dependant upon operational flow stressing on the aerofoil 44 in use. It will be understood that the aerofoil 44 may be stressed such that the transitional profile in portions 43 should not reduce the overall operational efficiency or fatigue life of the blade 41 in use.
- the present blade will generally reduce the potential for the blade root in particular to cause damage to incident portions of the fan casing within which the fan blade of the present mounting arrangement and a fan blade assembly is secured. It will be understood if there is greater control of potential impact damage it is possible to more confidently use thinner fan casing thicknesses giving a reduction in overall weight.
- Figure 5 provides a perspective view of a prior blade mounting 50.
- a blade 51 includes an aerofoil 54 and a root portion 52.
- the root 52 and therefore the blade 51 are secured in a fan rotor disc 60 through a slot 61.
- the root 52 has a dovetail cross section and is slid along the length of the slot 51 during assembly such that an end face 56 is generally consistent with end surfaces 62 of the fan disc 60 about the slot 61.
- the root 52 is secured in the slot 51 through a slide end retainer assembly 63 to ensure appropriate presentation of the blade 51 in use.
- the fan rotor disc 61 incorporates annulus filler fixings for location of filler mountings between the blades 51 in use.
- edges of the root 52 create angular corners (3 in figure 1) and it is the potential for these angular corners to impinge upon fan casings which is avoided by the present configuration for the root (22 in figure 2) and the transition to the blade overall.
- the present blade and blade mounting arrangement achieves an overall reduction in weight by reducing the necessary thickness of fan casing to ensure that there is no penetration of that casing if the blade should fail.
- This advantage is achieved through altering the angle of the ends of the root portion along with additional profiling of the blade adjacent of these root ends. In short, by reducing the acuteness of the points the impact load area of any fragments from blade failure is broadened and therefore the impact force spread over a great area of the fan casing.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
Claims
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US11/992,303 US20090136356A1 (en) | 2005-10-19 | 2006-09-22 | Blade Mounting |
GB0802967A GB2442695A (en) | 2005-10-19 | 2006-09-22 | A blade mounting |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB0521242.8 | 2005-10-19 | ||
GBGB0521242.8A GB0521242D0 (en) | 2005-10-19 | 2005-10-19 | A blade mounting |
Publications (1)
Publication Number | Publication Date |
---|---|
WO2007045815A1 true WO2007045815A1 (en) | 2007-04-26 |
Family
ID=35452013
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/GB2006/003544 WO2007045815A1 (en) | 2005-10-19 | 2006-09-22 | A blade mounting |
Country Status (3)
Country | Link |
---|---|
US (1) | US20090136356A1 (en) |
GB (2) | GB0521242D0 (en) |
WO (1) | WO2007045815A1 (en) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20090185910A1 (en) * | 2007-10-30 | 2009-07-23 | Mclaughlan James | Gas-turbine blade root |
Families Citing this family (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9239062B2 (en) * | 2012-09-10 | 2016-01-19 | General Electric Company | Low radius ratio fan for a gas turbine engine |
US9677421B2 (en) * | 2012-10-24 | 2017-06-13 | United Technologies Corporation | Gas turbine engine rotor drain feature |
KR101882109B1 (en) * | 2016-12-23 | 2018-07-25 | 두산중공업 주식회사 | Gas turbine |
Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR1095392A (en) * | 1953-12-01 | 1955-06-01 | Csf | Gas turbine blades |
US3471127A (en) * | 1966-12-08 | 1969-10-07 | Gen Motors Corp | Turbomachine rotor |
US4050134A (en) * | 1974-10-29 | 1977-09-27 | Westinghouse Electric Corporation | Method for removing rotatable blades without removing the casting of a turbine |
GB2148404A (en) * | 1983-10-19 | 1985-05-30 | Gen Motors Corp | End seal for turbine blade base |
DE19705323A1 (en) * | 1997-02-12 | 1998-08-27 | Siemens Ag | Turbo-machine blade |
EP1219782A2 (en) * | 2000-12-21 | 2002-07-03 | United Technologies Corporation | Bladed rotor assembly |
US20030049130A1 (en) * | 2001-09-13 | 2003-03-13 | Miller Harold Edward | Method and system for replacing a compressor blade |
EP1355045A2 (en) * | 2002-04-16 | 2003-10-22 | United Technologies Corporation | Bladed rotor for a gas turbine engine with a tiered blade to hub interface |
EP1495829A1 (en) * | 2003-07-10 | 2005-01-12 | ROLLS-ROYCE plc | Method of linear friction welding of blades to aerofoil blisks and blade having a root with a taper ratio less than 2 |
Family Cites Families (12)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US1719415A (en) * | 1927-09-14 | 1929-07-02 | Westinghouse Electric & Mfg Co | Turbine-blade attachment |
BE755608A (en) * | 1969-09-04 | 1971-02-15 | Gen Electric | COMPRESSOR BLADES |
US3986793A (en) * | 1974-10-29 | 1976-10-19 | Westinghouse Electric Corporation | Turbine rotating blade |
US5067876A (en) * | 1990-03-29 | 1991-11-26 | General Electric Company | Gas turbine bladed disk |
US5088894A (en) * | 1990-05-02 | 1992-02-18 | Westinghouse Electric Corp. | Turbomachine blade fastening |
US5067877A (en) * | 1990-09-11 | 1991-11-26 | United Technologies Corporation | Fan blade axial retention device |
US5431542A (en) * | 1994-04-29 | 1995-07-11 | United Technologies Corporation | Ramped dovetail rails for rotor blade assembly |
US5599190A (en) * | 1995-01-27 | 1997-02-04 | The Whitaker Corporation | Communication wiring system including a reconfigurable outlet assembly |
US5836744A (en) * | 1997-04-24 | 1998-11-17 | United Technologies Corporation | Frangible fan blade |
JP4316168B2 (en) * | 2001-08-30 | 2009-08-19 | 株式会社東芝 | Method for selecting blade material and shape of steam turbine blade and steam turbine |
US6846159B2 (en) * | 2002-04-16 | 2005-01-25 | United Technologies Corporation | Chamfered attachment for a bladed rotor |
GB0302116D0 (en) * | 2003-01-30 | 2003-03-05 | Rolls Royce Plc | A rotor |
-
2005
- 2005-10-19 GB GBGB0521242.8A patent/GB0521242D0/en not_active Ceased
-
2006
- 2006-09-22 WO PCT/GB2006/003544 patent/WO2007045815A1/en active Application Filing
- 2006-09-22 GB GB0802967A patent/GB2442695A/en not_active Withdrawn
- 2006-09-22 US US11/992,303 patent/US20090136356A1/en not_active Abandoned
Patent Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR1095392A (en) * | 1953-12-01 | 1955-06-01 | Csf | Gas turbine blades |
US3471127A (en) * | 1966-12-08 | 1969-10-07 | Gen Motors Corp | Turbomachine rotor |
US4050134A (en) * | 1974-10-29 | 1977-09-27 | Westinghouse Electric Corporation | Method for removing rotatable blades without removing the casting of a turbine |
GB2148404A (en) * | 1983-10-19 | 1985-05-30 | Gen Motors Corp | End seal for turbine blade base |
DE19705323A1 (en) * | 1997-02-12 | 1998-08-27 | Siemens Ag | Turbo-machine blade |
EP1219782A2 (en) * | 2000-12-21 | 2002-07-03 | United Technologies Corporation | Bladed rotor assembly |
US20030049130A1 (en) * | 2001-09-13 | 2003-03-13 | Miller Harold Edward | Method and system for replacing a compressor blade |
EP1355045A2 (en) * | 2002-04-16 | 2003-10-22 | United Technologies Corporation | Bladed rotor for a gas turbine engine with a tiered blade to hub interface |
EP1495829A1 (en) * | 2003-07-10 | 2005-01-12 | ROLLS-ROYCE plc | Method of linear friction welding of blades to aerofoil blisks and blade having a root with a taper ratio less than 2 |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20090185910A1 (en) * | 2007-10-30 | 2009-07-23 | Mclaughlan James | Gas-turbine blade root |
US8721292B2 (en) * | 2007-10-30 | 2014-05-13 | Rolls-Royce Deutschland Ltd & Co Kg | Gas-turbine blade root |
Also Published As
Publication number | Publication date |
---|---|
GB2442695A (en) | 2008-04-09 |
GB0521242D0 (en) | 2005-11-23 |
US20090136356A1 (en) | 2009-05-28 |
GB0802967D0 (en) | 2008-03-26 |
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