US3412978A - Radial flow turbine or compressor rotor - Google Patents

Radial flow turbine or compressor rotor Download PDF

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Publication number
US3412978A
US3412978A US566910A US56691066A US3412978A US 3412978 A US3412978 A US 3412978A US 566910 A US566910 A US 566910A US 56691066 A US56691066 A US 56691066A US 3412978 A US3412978 A US 3412978A
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rotor
flow
radial flow
vanes
turbine
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US566910A
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Keenan John Gregory
Scott John Alexander Henderson
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Rolls Royce PLC
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Rolls Royce PLC
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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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/04—Blade-carrying members, e.g. rotors for radial-flow machines or engines
    • F01D5/043—Blade-carrying members, e.g. rotors for radial-flow machines or engines of the axial inlet- radial outlet, or vice versa, type
    • F01D5/045—Blade-carrying members, e.g. rotors for radial-flow machines or engines of the axial inlet- radial outlet, or vice versa, type the wheel comprising two adjacent bladed wheel portions, e.g. with interengaging blades for damping vibrations

Definitions

  • the radially inner portion of the vanes are thicker than the radially outer portions and the opposite axial ends of the rotor are of different diameters.
  • Each of the vanes also has a thickness measured at a constant radius with respect to the axis of rotation which increases with increasing distance from the end of the rotor of smaller diameter.
  • This invention concerns radial flow turbine or compressor rotors and, although not so restricted, it will hereinafter be described with reference to its use on a radial flow turbine for a gas turbine engine.
  • the vanes are substantially constant in cross-sectional area, and it has been found that the effective flow cross-sectional area of the flow passage defined between adjacent vanes varies substantially along the fiow path through the rotor.
  • the fiow cross-sectional area of the flow passages may increase by as much as 50% and then decrease by somewhat more than the initial increase to a final cross-sectional area somewhat less than the inlet cross-sectional area. It will be appreciated that such variation in cross-sectional area hinders smooth flow of the fluid through the rotor, thus causing losses due to turblence.
  • a radial flow turbine or compressor rotor having radially extending vanes defining flow passages therebetween, the radially inner portions of each vane being thicker than the radially outer portions thereof so that the effective flow cross sectional area of each flow passage varies at a substantially constant rate from the inlet to the outlet thereof.
  • the rotor has opposite axial ends of different diameter and the thickness of each vane, measured at a constant radius with respect to the axis of rotation of the rotor, increases, at least in the radially inner part of the rotor, with increasing distance from the smaller diameter end of the rotor.
  • the rotor may form part of a radial flow turbine, the flow cross-sectional area of each fiow passage decreasing at a substantially constant rate in the downstream direction.
  • Said turbine may be that of a gas turbine engine.
  • FIGURE 1 is a broken-away elevation of a gas turbine engine including a radial flow turbine in accordance with the present invention
  • FIGURES 2 to are cross-sectional views on the lines 2-2, 33, 44 and 5--5 of FIGURE 1,
  • FIGURE 6 is a cross-sectional view to a larger scale of part of the radial flow turbine shown in FIGURE 1, and
  • FIGURE 7 is a graph illustrating the flow characteristics of the rotor shown in FIGURE 6.
  • a gas turbine engine 10 comprising a radial flow compressor 11, combustiOn equipment 12 and a radial fiow turbine 13.
  • Radial flow turbine 13 includes a radial flow rotor 14, comprising a hub 15 on which there are mounted a plurality of radially extending blades or vanes 16. As more clearly seen in FIGURE 6, the rotor is formed on two portions connected together by four dowel pins 17. r
  • FIG. 1 The cross-sectional form of the blades or vanes 16 is shown in detail in FIGURES 2 to 5.
  • the full line figures illustrate the cross sectional shape of the modified rotor in accordance with the present invention and the broken line indicates the form of the blades on a conventional radial filow turbine rotor. It will be noted that the rotor blades are the same at the radially outer portion of the rotor, but, as the radially inner portion of the rotor is approached, the rotor blades thicken compared with those of a conventional turbine rotor, whereby the cross-sectional area of the flow passage defined between them is decreased.
  • FIG. 6 Indicated on FIGURE 6 is an area C, defined within a broken outline, this area indicating that in which the vanes 16 are thickened to provide the decrease in cross-sectional area which enables a constant decrease in area to be obtained.
  • the thickening of the vanes shown in FIG- URES 2 to 5 is achieved by thickening the vanes in the area C, the remaining parts of the vanes being left at their original thickness, i.e. that shown, for example, in FIG- URE 2.
  • the curved graph A indicates the variation in flow area of the flow passages defined between adjacent vanes on a typical conventional turbine, i.e. that having vanes in accordance with the broken outline shown in FIGURES 2-5. It will be noted that the flow passages commence with an area of approximately 2.5 square inches which increases to nearly 3.5 square inches and then decreases to somewhat less than 2 inches adjacent the outlet to the rotor. It will be appreciated that such a variation in cross-sectional area discourages smooth flow of fluid through the rotor and creates turbulence and thus loss in efficiency.
  • the linear graph B indicates the flow passage characteristic with a rotor in accordance with the present invention, i.e. having blades 16 in accordance with the full line arrangement shown in FIGURES 2-5. It will be appreciated that the thickening of the vanes 16 adjacent the radially inner portion of the rotor decreases the cross-sectional area of the flow passages adjacent the radially inner portion of the rotor, and thus the substantial increase in crosssectional area indicated by graph A is reduced and the linear form of graph indicated at B is produced. With such a form it will be appreciated that the decrease in crosssectional area of each flow passage in a downstream direction is constant, and thus smooth flow of the fluid is obtained.
  • the invention is not restricted to use in a radial fiow turbine, but may, for example, be employed in a radial flow compressor.
  • the fluid in this case will fiow radially outwardly and thus the flow passages between adjacent vanes will effectively increase at a substantially constant rate in a downstream direction.
  • a radial flow rotor having an inlet and an outlet adjacent opposite axial ends and vanes extending from the inlet to the outlet and defining flow passages therebetween, said axial ends of said radial flow rotor having different diameters; each of said vanes having radially inner portions thicker than radially outer portions thereof; and each of said vanes, when measured at a constant radius with respect to the axis of rotation of the rotor, increasing in thickness with increasing distance from the end of the smaller diameter of said rotor so that each of said flow passages has an effective flow cross-sectional area varying at a substantially constant rate from the inlet to the outlet of said rotor,
  • a radial flow rotor as claimed in claim 1 in which said ellective flow cross-sectional area of each of said flow passages decreases at a substantially constant rate from the inlet to the outlet of said rotor.
  • a radial flow rotor as claimed in claim 1 in which the effective flow cross-sectional area of each of said flow passages increases at a substantially constant rate from the inlet to the outlet of said rotor.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
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Description

1968 J. G. KEENAN ETAL 3,
RADIAL FLOW TURBINE OR COMPRESSOR ROTOR Filed July 21. 1966 2 Sheets-Sheet 1 I nvenlors Nov. 26, 1968 J. G. KEENAN ETAL 3,412,973
RADIAL FLOW TURBINE OR'COMPRESSOR ROTOR Filed July 2L, 1966 2 Sheets-Sheet 2 3'0 [101/ ARM $00M! M058 )5 new ill/G7 0/ mm; 1mm
Allomiys United States Patent Office Patented Nov. 26, 1968 3,412,978 RADIAL FLOW TURBINE OR COMPRESSOR ROTOR John Gregory Keenan and John Alexander Henderson Scott, Derby, England, assiguors to Rolls-Royce Limited, Derby, England, a British company Filed July 21, 1966, Ser. No. 566,910 Claims priority, application Great Britain, Aug. 6, 1965, 33,872/ 65 3 Claims. (Cl. 253-39) ABSTRACT OF THE DISCLOSURE A radial flow compressor or turbine rotor having an inlet and an outlet with vanes extending from the inlet to the outlet and defining flow passages therebetween. The radially inner portion of the vanes are thicker than the radially outer portions and the opposite axial ends of the rotor are of different diameters. Each of the vanes also has a thickness measured at a constant radius with respect to the axis of rotation which increases with increasing distance from the end of the rotor of smaller diameter. These features of the rotor provide each flow passage with an effective flow cross-sectional area varying at a substantially constant rate from the inlet to the outlet of the rotor.
This invention concerns radial flow turbine or compressor rotors and, although not so restricted, it will hereinafter be described with reference to its use on a radial flow turbine for a gas turbine engine.
In conventional radial flow turbine rotors, the vanes are substantially constant in cross-sectional area, and it has been found that the effective flow cross-sectional area of the flow passage defined between adjacent vanes varies substantially along the fiow path through the rotor. Thus, in a conventional radial flow turbine, it has been found that, from the inlet, the fiow cross-sectional area of the flow passages may increase by as much as 50% and then decrease by somewhat more than the initial increase to a final cross-sectional area somewhat less than the inlet cross-sectional area. It will be appreciated that such variation in cross-sectional area hinders smooth flow of the fluid through the rotor, thus causing losses due to turblence.
According, therefore, to the present invention there is provided a radial flow turbine or compressor rotor having radially extending vanes defining flow passages therebetween, the radially inner portions of each vane being thicker than the radially outer portions thereof so that the effective flow cross sectional area of each flow passage varies at a substantially constant rate from the inlet to the outlet thereof.
In a preferred embodiment, the rotor has opposite axial ends of different diameter and the thickness of each vane, measured at a constant radius with respect to the axis of rotation of the rotor, increases, at least in the radially inner part of the rotor, with increasing distance from the smaller diameter end of the rotor.
The rotor may form part of a radial flow turbine, the flow cross-sectional area of each fiow passage decreasing at a substantially constant rate in the downstream direction. Said turbine may be that of a gas turbine engine.
The invention is illustrated, merely by way of example, in the accompanying drawings, in which:
FIGURE 1 is a broken-away elevation of a gas turbine engine including a radial flow turbine in accordance with the present invention,
FIGURES 2 to are cross-sectional views on the lines 2-2, 33, 44 and 5--5 of FIGURE 1,
FIGURE 6 is a cross-sectional view to a larger scale of part of the radial flow turbine shown in FIGURE 1, and
FIGURE 7 is a graph illustrating the flow characteristics of the rotor shown in FIGURE 6.
Referring to the drawings, there is shown a gas turbine engine 10 comprising a radial flow compressor 11, combustiOn equipment 12 and a radial fiow turbine 13. Radial flow turbine 13 includes a radial flow rotor 14, comprising a hub 15 on which there are mounted a plurality of radially extending blades or vanes 16. As more clearly seen in FIGURE 6, the rotor is formed on two portions connected together by four dowel pins 17. r
The cross-sectional form of the blades or vanes 16 is shown in detail in FIGURES 2 to 5. The full line figures illustrate the cross sectional shape of the modified rotor in accordance with the present invention and the broken line indicates the form of the blades on a conventional radial filow turbine rotor. It will be noted that the rotor blades are the same at the radially outer portion of the rotor, but, as the radially inner portion of the rotor is approached, the rotor blades thicken compared with those of a conventional turbine rotor, whereby the cross-sectional area of the flow passage defined between them is decreased.
Indicated on FIGURE 6 is an area C, defined within a broken outline, this area indicating that in which the vanes 16 are thickened to provide the decrease in cross-sectional area which enables a constant decrease in area to be obtained. Thus, the thickening of the vanes shown in FIG- URES 2 to 5 is achieved by thickening the vanes in the area C, the remaining parts of the vanes being left at their original thickness, i.e. that shown, for example, in FIG- URE 2.
Referring now to FIGURE 7, the curved graph A indicates the variation in flow area of the flow passages defined between adjacent vanes on a typical conventional turbine, i.e. that having vanes in accordance with the broken outline shown in FIGURES 2-5. It will be noted that the flow passages commence with an area of approximately 2.5 square inches which increases to nearly 3.5 square inches and then decreases to somewhat less than 2 inches adjacent the outlet to the rotor. It will be appreciated that such a variation in cross-sectional area discourages smooth flow of fluid through the rotor and creates turbulence and thus loss in efficiency.
The linear graph B indicates the flow passage characteristic with a rotor in accordance with the present invention, i.e. having blades 16 in accordance with the full line arrangement shown in FIGURES 2-5. It will be appreciated that the thickening of the vanes 16 adjacent the radially inner portion of the rotor decreases the cross-sectional area of the flow passages adjacent the radially inner portion of the rotor, and thus the substantial increase in crosssectional area indicated by graph A is reduced and the linear form of graph indicated at B is produced. With such a form it will be appreciated that the decrease in crosssectional area of each flow passage in a downstream direction is constant, and thus smooth flow of the fluid is obtained.
It will be appreciated that the flow characteristics of a radial flow rotor in accordance with the present invention are greatly enhanced compared with those of a conventional radial flow rotor.
It will be appreciated that the invention is not restricted to use in a radial fiow turbine, but may, for example, be employed in a radial flow compressor. The fluid in this case will fiow radially outwardly and thus the flow passages between adjacent vanes will effectively increase at a substantially constant rate in a downstream direction.
We claim:
1. A radial flow rotor having an inlet and an outlet adjacent opposite axial ends and vanes extending from the inlet to the outlet and defining flow passages therebetween, said axial ends of said radial flow rotor having different diameters; each of said vanes having radially inner portions thicker than radially outer portions thereof; and each of said vanes, when measured at a constant radius with respect to the axis of rotation of the rotor, increasing in thickness with increasing distance from the end of the smaller diameter of said rotor so that each of said flow passages has an effective flow cross-sectional area varying at a substantially constant rate from the inlet to the outlet of said rotor,
2. A radial flow rotor as claimed in claim 1 in which said ellective flow cross-sectional area of each of said flow passages decreases at a substantially constant rate from the inlet to the outlet of said rotor.
3. A radial flow rotor as claimed in claim 1 in which the effective flow cross-sectional area of each of said flow passages increases at a substantially constant rate from the inlet to the outlet of said rotor.
References Cited UNITED STATES PATENTS EVERETTE A. POWELL, JR., Primary Examiner.
US566910A 1965-08-06 1966-07-21 Radial flow turbine or compressor rotor Expired - Lifetime US3412978A (en)

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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20100215508A1 (en) * 2009-02-25 2010-08-26 Behzad Hagshenas Axially Segmented Impeller
JP2014001712A (en) * 2012-06-20 2014-01-09 Toyota Central R&D Labs Inc Radial turbine rotor, and variable geometry turbocharger including the same
JP2015021397A (en) * 2013-07-16 2015-02-02 株式会社豊田中央研究所 Turbine rotor, turbine unit and turbocharger
US20150247409A1 (en) * 2012-04-11 2015-09-03 Honeywell International Inc. Axially-split radial turbines
US11421702B2 (en) 2019-08-21 2022-08-23 Pratt & Whitney Canada Corp. Impeller with chordwise vane thickness variation

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2621851A (en) * 1945-05-31 1952-12-16 Power Jets Res & Dev Ltd Rotary impeller and the like
US2819012A (en) * 1950-12-22 1958-01-07 Gen Motors Corp Centrifugal compressor
FR1150392A (en) * 1955-06-18 1958-01-10 Gas turbine rotor
US3032315A (en) * 1955-08-16 1962-05-01 Laval Steam Turbine Co Turbine blading

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2621851A (en) * 1945-05-31 1952-12-16 Power Jets Res & Dev Ltd Rotary impeller and the like
US2819012A (en) * 1950-12-22 1958-01-07 Gen Motors Corp Centrifugal compressor
FR1150392A (en) * 1955-06-18 1958-01-10 Gas turbine rotor
US3032315A (en) * 1955-08-16 1962-05-01 Laval Steam Turbine Co Turbine blading

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20100215508A1 (en) * 2009-02-25 2010-08-26 Behzad Hagshenas Axially Segmented Impeller
US8147208B2 (en) * 2009-02-25 2012-04-03 Hamilton Sundstrand Corporation Axially segmented impeller
US20150247409A1 (en) * 2012-04-11 2015-09-03 Honeywell International Inc. Axially-split radial turbines
US9726022B2 (en) * 2012-04-11 2017-08-08 Honeywell International Inc. Axially-split radial turbines
JP2014001712A (en) * 2012-06-20 2014-01-09 Toyota Central R&D Labs Inc Radial turbine rotor, and variable geometry turbocharger including the same
JP2015021397A (en) * 2013-07-16 2015-02-02 株式会社豊田中央研究所 Turbine rotor, turbine unit and turbocharger
US11421702B2 (en) 2019-08-21 2022-08-23 Pratt & Whitney Canada Corp. Impeller with chordwise vane thickness variation

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