US3620021A - Gas turbine engines - Google Patents

Gas turbine engines Download PDF

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Publication number
US3620021A
US3620021A US28361A US3620021DA US3620021A US 3620021 A US3620021 A US 3620021A US 28361 A US28361 A US 28361A US 3620021D A US3620021D A US 3620021DA US 3620021 A US3620021 A US 3620021A
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Prior art keywords
fan
turbine
engine
main fan
pass
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Expired - Lifetime
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US28361A
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Brian Wilfred Lawrie
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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
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D19/00—Axial-flow pumps
    • F04D19/02—Multi-stage pumps
    • F04D19/026—Multi-stage pumps with a plurality of shafts rotating at different speeds
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan

Definitions

  • a gas turbine by-pass engine has a compressor means, a combustion means and a turbine means, a freely rotatable main fan extends completely across the by-pass duct and a secondary fan is positioned upstream of the main fan and extends only partially across the by-pass duct, the secondary fan being driven by the turbine means.
  • This invention relates to gas turbine by-pass engines in which a fan is mounted in the by-pass duct of the engine.
  • a gas turbine by-pass engine having compressor means, combustion means and turbine means, a freely rotatable main fan extending completely across the by-pass duct and a rotatable secondary fan positioned upstream of the main fan and extending only partially across the by-pass duct, the secondary fan being driven by the turbine means and being arranged to drive the main fan.
  • the tip speed of both the main fan and sec ondary fan is arranged to give a blade relative inlet Mach number always less than unity.
  • the main fan may be arranged to be mounted for free rotation on a compressor casing.
  • FIG. 1 shows a diagrammatic representation of one form of gas turbine by-pass engine according to the present invention and FIG. 2 shows a modified form of the engine shown in FIG. 1.
  • a gas turbine by-pass engine consists of compressor means combustion means 12 and turbine means 14.
  • the compressor means 10 has a multistage compressor 16 driven by a turbine 18 through a shaft 20 and a conventional high pressure compressor 22 driven by a turbine 24 through a shaft 26.
  • a main fan 28 extends completely across the by-pass duct 30- of the engine and is mounted on bearings 50 which are located on a casing 32 of the compressor 16.
  • the main fan 28 is in the form of a spoked wheel having a hub, a number of radial blades or spokes 28a which act as turbine blades connected to a splitter ring 38 and a number of fan blades attached to the splitter ring.
  • a secondary fan 40 is mounted upstream of the main fan and the blades of the fan extend only as far as the splitter ring of the main fan.
  • the secondary fan 40 is mounted on a shaft 42 which is a driven by a turbine 44.
  • the shafts 26 and 42 can rotate either in the same or the opposite direction both to the shaft 20 and to each other.
  • the torque required to drive the main fan 28 is transmitted aero-dynamically through the blades 28a acting as turbine blades and sufficient torque is provided by the turbine 44 through the shaft 42, to drive the secondary fan 40 and the main fan 28.
  • Both the main fan and the secondary fan are arranged to operate at sufficiently low tip speeds to ensure that the blade tip relative inlet fioW Mach numbers are less than unity. This arrangement avoids undesirable tone noise which is associated with blade relative inlet flow Mach numbers of greater than unity.
  • the main fan 28 in order to reduce the diameter of the bearing means 50, the main fan 28 is extended inwardly and is mounted on relatively small diameter bearing means 52 which are located between the compressor 16 and the secondary fan 40.
  • a principal difiiculty with fans of low hub/ tip ratio, as in the case of a conventional high by-pass ratio engine is that the tip runs faster than necessary for maximum efficiency if the root is to provide a satisfactory pressure rise.
  • the provision of the secondary fan 40 both enables a satisfactory pressure rise to be obtained at the root of the main fan and the intake to compressor 16 to allow the tip speed of the main fan to be reduced, thus giving a reduced main fan white noise level.
  • the present arrangement offers a number of advantages.
  • a lower number of turbine stages is required; this is because the torque required to drive the low speed main fan 28 is transmitted aerodynamically through the blades 28a; e.g. a high by-pass ratio fan engine having eleven turbine stages can be reduced to an engine having five turbine stages.
  • the secondary fan requires a small amount of torque which is produced by the moderately loaded turbine 44. This ensures a sufficiently low overall turbine noise level.
  • a higher compression ratio of the by-pass flow can be obtained and this is due to the presence of the secondary fan which reduces the whirl angle of the flow onto a ring of outlet guide vanes 46 in the region of the roots of the guide vanes, thus relieving the high outlet guide vanes root diffusion loading associated with the conventional low tip speed front fan engine design.
  • a gas turbine engine having a by-pass duct and comprising:
  • compressor means combustion means and turbine means in flow series
  • said compressor means including a low pressure compressor and a high pressure compressor and said turbine means including a high pressure stage, an intermediate pressure stage, and a low pressure stage; a freely rotatable main fan extending completely across said by-pass duct; a rotatable secondary fan positioned upstream of said main fan and extending only partially across said by-pass duct; a first shaft for driving said high pressure com pressor by said high pressure stage of said turbine means; a second shaft for driving said low pressure compressor by said intermediate pressure stage of said turbine means; and a third shaft for driving said secondary fan by said low pressure stage of said turbine means.
  • An engine as claimed in claim 1 including a splitter secured to said main fan, and in which said secondary fan is rotatable within said splitter.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A GAS TURBINE BY-PASS ENGINE HAS A COMPRESSOR MEANS, A COMBUSTION MEANS AND A TURBINE MEANS, A FREELY ROTATABLE MAIN FAN EXTENDS COMPLETELY ACROSS THE BY-PASS DUCT AND A SECONDARY FAN IS POSITIONED UPSTREAM OF THE MAIN FAN AND EXTENDS ONLY PARTIALLY ACROSS THE BY-PASS DUCT, THE SECONDARY FAN BEING DRIVEN BY THE TRUBINE MEANS.

Description

Nov. 16, 1971' a. w; LAWRIE 3,620,021
Gas TURBINE mamas Fi1ed.Apr11 14, 1970 I 2 Sheets-Shoot 1 Nov. 16, 1971 a. w. LAWRIE us wmzsmn mamas 2 Sheets-Shut I Filed April l4, 1970 mwwl \lll
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I I n venlar Zena/v M/KMJflWZ/E United States Patent ()1 lice 3,620,021 Patented Nov. 16, 1971 3,620,021 GAS TURBINE ENGINES Brian Wilfred Lawrie, Allestree, England, assignor to Rolls-Royce Limited, Derby, England Filed Apr. 14, 1970, Ser. No. 28,361 Int. Cl. F02k 3/04 US. Cl. 60-226 5 Claims ABSTRACT OF THE DISCLOSURE A gas turbine by-pass engine has a compressor means, a combustion means and a turbine means, a freely rotatable main fan extends completely across the by-pass duct and a secondary fan is positioned upstream of the main fan and extends only partially across the by-pass duct, the secondary fan being driven by the turbine means.
This invention relates to gas turbine by-pass engines in which a fan is mounted in the by-pass duct of the engine.
According to the present invention there is provided a gas turbine by-pass engine having compressor means, combustion means and turbine means, a freely rotatable main fan extending completely across the by-pass duct and a rotatable secondary fan positioned upstream of the main fan and extending only partially across the by-pass duct, the secondary fan being driven by the turbine means and being arranged to drive the main fan.
Preferably the tip speed of both the main fan and sec ondary fan is arranged to give a blade relative inlet Mach number always less than unity. The main fan may be arranged to be mounted for free rotation on a compressor casing.
The invention will now be more particularly described with reference to the accompanying drawings in which:
FIG. 1 shows a diagrammatic representation of one form of gas turbine by-pass engine according to the present invention and FIG. 2 shows a modified form of the engine shown in FIG. 1.
In FIG. 1 a gas turbine by-pass engine consists of compressor means combustion means 12 and turbine means 14.
The compressor means 10 has a multistage compressor 16 driven by a turbine 18 through a shaft 20 and a conventional high pressure compressor 22 driven by a turbine 24 through a shaft 26. A main fan 28 extends completely across the by-pass duct 30- of the engine and is mounted on bearings 50 which are located on a casing 32 of the compressor 16. The main fan 28 is in the form of a spoked wheel having a hub, a number of radial blades or spokes 28a which act as turbine blades connected to a splitter ring 38 and a number of fan blades attached to the splitter ring. A secondary fan 40 is mounted upstream of the main fan and the blades of the fan extend only as far as the splitter ring of the main fan. The secondary fan 40 is mounted on a shaft 42 which is a driven by a turbine 44. The shafts 26 and 42 can rotate either in the same or the opposite direction both to the shaft 20 and to each other.
The torque required to drive the main fan 28 is transmitted aero-dynamically through the blades 28a acting as turbine blades and sufficient torque is provided by the turbine 44 through the shaft 42, to drive the secondary fan 40 and the main fan 28.
Both the main fan and the secondary fan are arranged to operate at sufficiently low tip speeds to ensure that the blade tip relative inlet fioW Mach numbers are less than unity. This arrangement avoids undesirable tone noise which is associated with blade relative inlet flow Mach numbers of greater than unity.
In FIG. 2, in order to reduce the diameter of the bearing means 50, the main fan 28 is extended inwardly and is mounted on relatively small diameter bearing means 52 which are located between the compressor 16 and the secondary fan 40.
A principal difiiculty with fans of low hub/ tip ratio, as in the case of a conventional high by-pass ratio engine is that the tip runs faster than necessary for maximum efficiency if the root is to provide a satisfactory pressure rise. The provision of the secondary fan 40 both enables a satisfactory pressure rise to be obtained at the root of the main fan and the intake to compressor 16 to allow the tip speed of the main fan to be reduced, thus giving a reduced main fan white noise level.
In comparison with a conventional front fan three shaft engine of high by-pass ratio in which a low fan tip speed is required the present arrangement offers a number of advantages. A lower number of turbine stages is required; this is because the torque required to drive the low speed main fan 28 is transmitted aerodynamically through the blades 28a; e.g. a high by-pass ratio fan engine having eleven turbine stages can be reduced to an engine having five turbine stages. The secondary fan requires a small amount of torque which is produced by the moderately loaded turbine 44. This ensures a sufficiently low overall turbine noise level.
A higher compression ratio of the by-pass flow can be obtained and this is due to the presence of the secondary fan which reduces the whirl angle of the flow onto a ring of outlet guide vanes 46 in the region of the roots of the guide vanes, thus relieving the high outlet guide vanes root diffusion loading associated with the conventional low tip speed front fan engine design.
I claim:
1. A gas turbine engine having a by-pass duct and comprising:
compressor means, combustion means and turbine means in flow series, said compressor means including a low pressure compressor and a high pressure compressor and said turbine means including a high pressure stage, an intermediate pressure stage, and a low pressure stage; a freely rotatable main fan extending completely across said by-pass duct; a rotatable secondary fan positioned upstream of said main fan and extending only partially across said by-pass duct; a first shaft for driving said high pressure com pressor by said high pressure stage of said turbine means; a second shaft for driving said low pressure compressor by said intermediate pressure stage of said turbine means; and a third shaft for driving said secondary fan by said low pressure stage of said turbine means.
2. An engine as claim in claim 1 in which said low pressure compressor has an outer casing, bearing means mounted on said outer casing, and said main fan being mounted on said bearing means.
3. An engine as claimed in claim 1 in which said low pressure compressor has an intake, and including bearing means located between said secondary fan and said intake, said main fan being mounted on said bearing means.
4. An engine as claimed in claim 1 including a splitter secured to said main fan, and in which said secondary fan is rotatable within said splitter.
5. An engine as claimed in claim 1 in which both the main fan and the secondary fan are arranged each to have a tip speed to give a blade relative inlet Mach number of less than unity.
References Cited UNITED STATES PATENTS ALLAN D. HERRMANN, Primary Examiner US. Cl. X.R.
US28361A 1970-04-14 1970-04-14 Gas turbine engines Expired - Lifetime US3620021A (en)

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Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3729957A (en) * 1971-01-08 1973-05-01 Secr Defence Fan
US3747343A (en) * 1972-02-10 1973-07-24 United Aircraft Corp Low noise prop-fan
US4790133A (en) * 1986-08-29 1988-12-13 General Electric Company High bypass ratio counterrotating turbofan engine
US4860537A (en) * 1986-08-29 1989-08-29 Brandt, Inc. High bypass ratio counterrotating gearless front fan engine
US6209311B1 (en) 1998-04-13 2001-04-03 Nikkiso Company, Ltd. Turbofan engine including fans with reduced speed
US20060016172A1 (en) * 2004-05-28 2006-01-26 Rolls-Royce Plc Gas turbine engine
US20080098717A1 (en) * 2006-10-31 2008-05-01 Robert Joseph Orlando Turbofan engine assembly and method of assembling same
US20080098714A1 (en) * 2006-10-31 2008-05-01 Robert Joseph Orlando Turbofan engine assembly and method of assembling same
US20090107109A1 (en) * 2007-10-26 2009-04-30 United Technologies Corporation Variable bypass turbine fan
US20090193785A1 (en) * 2008-01-31 2009-08-06 General Electric Company Power generating turbine systems
US20110146228A1 (en) * 2009-12-21 2011-06-23 John Lewis Baughman Power extraction system
US20110146289A1 (en) * 2009-12-21 2011-06-23 John Lewis Baughman Power extraction method
US20130287545A1 (en) * 2012-04-25 2013-10-31 Gabriel L. Suciu Geared turbofan with three turbines all counter-rotating
US9976489B2 (en) * 2013-12-16 2018-05-22 United Technologies Corporation Gas turbine engine for long range aircraft

Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3729957A (en) * 1971-01-08 1973-05-01 Secr Defence Fan
US3747343A (en) * 1972-02-10 1973-07-24 United Aircraft Corp Low noise prop-fan
US4790133A (en) * 1986-08-29 1988-12-13 General Electric Company High bypass ratio counterrotating turbofan engine
US4860537A (en) * 1986-08-29 1989-08-29 Brandt, Inc. High bypass ratio counterrotating gearless front fan engine
US6209311B1 (en) 1998-04-13 2001-04-03 Nikkiso Company, Ltd. Turbofan engine including fans with reduced speed
US20060016172A1 (en) * 2004-05-28 2006-01-26 Rolls-Royce Plc Gas turbine engine
US7500352B2 (en) * 2004-05-28 2009-03-10 Rolls-Royce Plc Gas turbine engine
US7921634B2 (en) * 2006-10-31 2011-04-12 General Electric Company Turbofan engine assembly and method of assembling same
US20080098717A1 (en) * 2006-10-31 2008-05-01 Robert Joseph Orlando Turbofan engine assembly and method of assembling same
US20080098714A1 (en) * 2006-10-31 2008-05-01 Robert Joseph Orlando Turbofan engine assembly and method of assembling same
US7926259B2 (en) * 2006-10-31 2011-04-19 General Electric Company Turbofan engine assembly and method of assembling same
US8028513B2 (en) * 2007-10-26 2011-10-04 United Technologies Corporation Variable bypass turbine fan
US20090107109A1 (en) * 2007-10-26 2009-04-30 United Technologies Corporation Variable bypass turbine fan
US20090193785A1 (en) * 2008-01-31 2009-08-06 General Electric Company Power generating turbine systems
US20110146228A1 (en) * 2009-12-21 2011-06-23 John Lewis Baughman Power extraction system
US20110146289A1 (en) * 2009-12-21 2011-06-23 John Lewis Baughman Power extraction method
US20130287545A1 (en) * 2012-04-25 2013-10-31 Gabriel L. Suciu Geared turbofan with three turbines all counter-rotating
US9074485B2 (en) * 2012-04-25 2015-07-07 United Technologies Corporation Geared turbofan with three turbines all counter-rotating
US20150361881A1 (en) * 2012-04-25 2015-12-17 United Technologies Corporation Geared turbofan with three turbines all counter-rotating
US20180128168A1 (en) * 2012-04-25 2018-05-10 United Technologies Corporation Geared turbofan with three turbines all counter-rotating
US10830130B2 (en) * 2012-04-25 2020-11-10 Raytheon Technologies Corporation Geared turbofan with three turbines all counter-rotating
US9976489B2 (en) * 2013-12-16 2018-05-22 United Technologies Corporation Gas turbine engine for long range aircraft

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