US3620020A - Gas turbine engine - Google Patents

Gas turbine engine Download PDF

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Publication number
US3620020A
US3620020A US28360A US3620020DA US3620020A US 3620020 A US3620020 A US 3620020A US 28360 A US28360 A US 28360A US 3620020D A US3620020D A US 3620020DA US 3620020 A US3620020 A US 3620020A
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United States
Prior art keywords
fan
turbine
compressor
main fan
pass
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Expired - Lifetime
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US28360A
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David George Halliwell
Michael Vaughan Outram
Martin Hume Bryan-Brown
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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
    • 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 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 is driven by the turbine means and the main fan is mounted on a casing having blades driven by the compressor 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 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.
  • the tip speed of both the main fan and secondary fan is arranged to give a blade relative inlet Mach number always less than unity.
  • the main fan is arranged to be mounted on a casing having blades driven by the compressor means.
  • a gas turbine by-pass engine consists of compressor means 10, combustion means 12 and turbine means 14.
  • the compressor means 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 a rotatable casing 32 which has rows of stator blades (which remove energy from the air and therefore act as turbine blades) positioned between rows of compressor blades 36 on the low pressure compressor 16.
  • the main fan 28 is in the form of a spoked wheel having a hub, a number of radial blades or spokes 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 driven by a turbine 44.
  • the shaft 20 and the casing 32 rotate in the same direction as shown by the arrows A and the remaining 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 aerodynamically through the low pressure compressor stators which are in fact the blades 34 acting as turbine blades. Suflicient torque is provided by the turbine Patented Nov. 16, 1971 18 through the shaft 20 and hence the low pressure compressor blades 36, to drive the compressor stators and to provide the required net pressure rise over the compressor 16.
  • Both the main fan and the secondary fan are arranged to operate at sufiiciently low tip speeds to ensure that the blade tip relative inlet flow 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.
  • 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 roof 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 stators 34 from the high speed blade rows 36; these are supplied with sufficient torque by the high speed turbine 18 to drive the low speed main fan 28 and to provide the required pressure rise over the low pressure compressor 16; 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 40 produces the pressure rise required over the inner portion of the by-pass duct in order that the torque required by the main fan balances the torque transmitted by the stators 34.
  • the secondary fan requires a small amount of torque which is produced by the moderately loaded turbine 44. This ensures a sufiiciently 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 vane root diffusion loading associated with 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; a rotatable main fan extending completely across said by-pass duct; a rotatable secondary fan extending only partially across said bypass duct and rotatable within an annular splitter secured to said main fan, said compressor means having a high pressure and a low pressure compressor, said low pressure compressor having a rotatable casing to which said main fan is secured and is thereby arranged to be driven by said low pressure compressor, said turbine means having a high pressure, an intermediate pressure and a low pressure stage, said high pressure stage being arranged to drive said 'high pressure compressor by means of a first shaft, said intermediate pressure stage being arranged to drive said low pressure compressor by means of -a second 'shaft and said low pressure stage being arranged to drive said secondary fan by means of a third shaft.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (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 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 SECOND-

ARY FAN IS DRIVEN BY THE TURBINE MEANS AND THE MAIN FAN IS MOUNTED ON A CASING HAVING BLADES DRIVEN BY THE COMPRESSOR MEANS.

Description

GAS TURBINE ENGINE Filed A ril 14. 1970 Mm mavdurmm By Merl/v 01m:
/ RYAN-Beam Attorneys United States Patent 3,620,020 GAS TURBINE ENGINE David George Halliwell, Allestree, Michael Vaughan Outram, Eggington, and Martin Hume Bryan-Brown, Etwall, England, assignors to Rolls-Royce Limited Filed Apr. 14, 1970, Ser. No. 28,360 Claims priority, application Great Britain, Apr. 16, 1969, 19,360/ 69 Int. Cl. F02k 3/04 US. Cl. 60-226 2 Claims ABSTRACT OF THE DISCLOSURE A gas turbine by-pass engine has a compressor means, a combustion means and a turbine means, a 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 is driven by the turbine means and the main fan is mounted on a casing having blades driven by the compressor 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 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.
Preferably the tip speed of both the main fan and secondary fan is arranged to give a blade relative inlet Mach number always less than unity. The main fan is arranged to be mounted on a casing having blades driven by the compressor means.
. The invention will now be more particularly described with reference to the accompanying drawing which shows a diagrammatic representation of one form of gas turbine by-pass engine according to the present invention.
In the drawing, a gas turbine by-pass engine consists of compressor means 10, combustion means 12 and turbine means 14.
The compressor means 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 a rotatable casing 32 which has rows of stator blades (which remove energy from the air and therefore act as turbine blades) positioned between rows of compressor blades 36 on the low pressure compressor 16. The main fan 28 is in the form of a spoked wheel having a hub, a number of radial blades or spokes 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 driven by a turbine 44. The shaft 20 and the casing 32 rotate in the same direction as shown by the arrows A and the remaining 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 aerodynamically through the low pressure compressor stators which are in fact the blades 34 acting as turbine blades. Suflicient torque is provided by the turbine Patented Nov. 16, 1971 18 through the shaft 20 and hence the low pressure compressor blades 36, to drive the compressor stators and to provide the required net pressure rise over the compressor 16.
Both the main fan and the secondary fan are arranged to operate at sufiiciently low tip speeds to ensure that the blade tip relative inlet flow 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.
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 roof 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 stators 34 from the high speed blade rows 36; these are supplied with sufficient torque by the high speed turbine 18 to drive the low speed main fan 28 and to provide the required pressure rise over the low pressure compressor 16; 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 40 produces the pressure rise required over the inner portion of the by-pass duct in order that the torque required by the main fan balances the torque transmitted by the stators 34. The secondary fan requires a small amount of torque which is produced by the moderately loaded turbine 44. This ensures a sufiiciently 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 vane root diffusion loading associated with conventional low tip speed front fan engine design.
We claim:
1. A gas turbine engine having a by-pass duct and comprising:
compressor means, combustion means and turbine means in flow series; a rotatable main fan extending completely across said by-pass duct; a rotatable secondary fan extending only partially across said bypass duct and rotatable within an annular splitter secured to said main fan, said compressor means having a high pressure and a low pressure compressor, said low pressure compressor having a rotatable casing to which said main fan is secured and is thereby arranged to be driven by said low pressure compressor, said turbine means having a high pressure, an intermediate pressure and a low pressure stage, said high pressure stage being arranged to drive said 'high pressure compressor by means of a first shaft, said intermediate pressure stage being arranged to drive said low pressure compressor by means of -a second 'shaft and said low pressure stage being arranged to drive said secondary fan by means of a third shaft.
2. An engine as claimed in claim 1 in which both said main fan and said secondary fan are arranged each to 3 4 have a tip speed to give a blade relative inlet Mach num- 3,385,064 5/1968 Wilde et a1. 415-79 ber of less than unity. 3,546,882 12/1970 Berkey 41579 UX References Cited ALLAN D. HERRMANN, Primary Examiner 3 448 582 S$59 ZTATES PiXTENTS 60 226 5 may at a 6O--39.16 R; 415-79 2,702,985 3/1955 Howell 6039.16 CR
US28360A 1969-04-16 1970-04-14 Gas turbine engine Expired - Lifetime US3620020A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB1936069 1969-04-16

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DE (1) DE2018077C3 (en)
FR (1) FR2039215B1 (en)
GB (1) GB1257497A (en)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS4891800A (en) * 1972-02-10 1973-11-29
US3861139A (en) * 1973-02-12 1975-01-21 Gen Electric Turbofan engine having counterrotating compressor and turbine elements and unique fan disposition
WO2010059321A3 (en) * 2008-11-21 2011-04-14 General Electric Company Gas turbine engine booster having rotatable radially inwardly extending blades and non-rotatable vanes
WO2013165862A1 (en) 2012-04-30 2013-11-07 United Technologies Corporation Geared turbofan with three turbines all co-rotating
US9739206B2 (en) 2012-01-31 2017-08-22 United Technologies Corporation Geared turbofan gas turbine engine architecture
US10030586B2 (en) 2012-01-31 2018-07-24 United Technologies Corporation Geared turbofan gas turbine engine architecture
US11608786B2 (en) 2012-04-02 2023-03-21 Raytheon Technologies Corporation Gas turbine engine with power density range
US11913349B2 (en) 2012-01-31 2024-02-27 Rtx Corporation Gas turbine engine with high speed low pressure turbine section and bearing support features

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2382382B (en) 2001-11-23 2005-08-10 Rolls Royce Plc A fan for a turbofan gas turbine engine
GB0809759D0 (en) * 2008-05-30 2008-07-09 Rolls Royce Plc Gas turbine engine
DE102014226696A1 (en) * 2014-12-19 2016-06-23 Rolls-Royce Deutschland Ltd & Co Kg Pre-compressor device, retrofit kit and aircraft engine
CA2923331A1 (en) * 2015-03-19 2016-09-19 David P. Houston Geared turbofan gas turbine engine architecture

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS4891800A (en) * 1972-02-10 1973-11-29
US3861139A (en) * 1973-02-12 1975-01-21 Gen Electric Turbofan engine having counterrotating compressor and turbine elements and unique fan disposition
WO2010059321A3 (en) * 2008-11-21 2011-04-14 General Electric Company Gas turbine engine booster having rotatable radially inwardly extending blades and non-rotatable vanes
US8166748B2 (en) 2008-11-21 2012-05-01 General Electric Company Gas turbine engine booster having rotatable radially inwardly extending blades and non-rotatable vanes
US9739206B2 (en) 2012-01-31 2017-08-22 United Technologies Corporation Geared turbofan gas turbine engine architecture
US9828944B2 (en) 2012-01-31 2017-11-28 United Technologies Corporation Geared turbofan gas turbine engine architecture
US10030586B2 (en) 2012-01-31 2018-07-24 United Technologies Corporation Geared turbofan gas turbine engine architecture
US11913349B2 (en) 2012-01-31 2024-02-27 Rtx Corporation Gas turbine engine with high speed low pressure turbine section and bearing support features
US11608786B2 (en) 2012-04-02 2023-03-21 Raytheon Technologies Corporation Gas turbine engine with power density range
US11970984B2 (en) 2012-04-02 2024-04-30 Rtx Corporation Gas turbine engine with power density range
EP2844861A4 (en) * 2012-04-30 2015-05-06 United Technologies Corp Geared turbofan with three turbines all co-rotating
WO2013165862A1 (en) 2012-04-30 2013-11-07 United Technologies Corporation Geared turbofan with three turbines all co-rotating
US10036350B2 (en) 2012-04-30 2018-07-31 United Technologies Corporation Geared turbofan with three turbines all co-rotating

Also Published As

Publication number Publication date
FR2039215A1 (en) 1971-01-15
DE2018077C3 (en) 1975-02-06
FR2039215B1 (en) 1974-11-08
GB1257497A (en) 1971-12-22
DE2018077B2 (en) 1974-06-20
DE2018077A1 (en) 1970-11-19

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