US3381475A

US3381475A - Jet propulsion engines

Info

Publication number: US3381475A
Application number: US560517A
Authority: US
Inventors: Derek Aubrey Roberts
Original assignee: Bristol Siddeley Engines Ltd
Current assignee: Bristol Siddeley Engines Ltd
Priority date: 1965-06-28
Filing date: 1966-06-27
Publication date: 1968-05-07
Anticipated expiration: 1985-05-07
Also published as: DE1956119U; DE1476948A1

Description

2 Sheets-Sheet 1 Filed June 27, 1966 INVENTQQ DEQEK AUBREY Rosems May 7, 1968 D. A. ROBERTS JET PROPULSION ENGINES 2 Sheets-Sheet :3

Filed June 27, 1966 \NvENToR DEREK AUBREY ROBERTS United States Patent 3,381,475 JET PROPULSIUN ENGllNES Derek Aubrey Roberts, Sunbury-onfiihames, England, assignor to Bristol Siddeley Engines Limited Filed June 27, 1966, Ser. No. 566,517

Claims priority, application Great Britain, June 23, 1965,

7 Claims. (Cl. 669-226) ABSTRACT GF THE DHSCLUSURE The disclosure of this invention pertains to a jet propulsion engine in which an aft-fan assembly is positioned downstream of the conventional turbine and is provided with second and third turbine rotors arranged so that the products of combustion act on the second turbine rotor and then directly on the third turbine rotor. The second and third turbine rotors being arranged to rotate in opposite directions and gearing connecting said second and third turbine rotors determine the relationship between their speeds. A ducted fan assembly is provided and includes a set of static blades and a set of rotary blades, the rotary blades being driven from said second and third turbine rotors and connected directly to one of said turbine rotors.

This invention relates to propulsion engines of the kind comprising a rotary air compressor, usually of the axial flow type, arranged to deliver air rearwardly to combustion chamber apparatus, a turbine rotor (herein called the first turbine rotor) on which the products of combustion from the combustion chamber apparatus act, arranged to drive the compressor rotor, and a so-called aft-fan assembly including a ducted fan which is arranged to be driven through further turbine apparatus on which the products of combustion act after acting on the first turbine rotor, and has its blades located in a duct through which, during operation, they impel air rearwardly to the nozzle or nozzles which, jointly with a nozzle or nozzles through which the products of combustion flow after acting on the turbine rotors, constitute the propulsion nozzles.

The term aft fan used herein is therefore to be understood in its normal sense as meaning a ducted fan arranged to be driven by turbine apparatus which is acted upon by the products of combustion after they have acted upon a main turbine rotor which is disposed in advance of the turbine apparatus driving the aft fan, and serves to drive an air compressor delivering air to combustion apparatus from which the products of combustion are derived.

For convenience it will be assumed herein that the products of combustion flow to a single hot propulsion nozzle and that the air from the ducted fan also flows to a single cold propulsion nozzle although in either or each case two or more nozzles may be employed. Further the terms hot propulsion nozzle" and cold propulsion nozzle will be used in normal manner to refer respectively to a nozzle (or nozzles) through which the products of combustion are ejected and a nozzle (or nozzles) through which air from the ducted fan is ejected, whether the ejection in each case is directly into the atmosphere or the nozzles are arranged (as they may be) so that the products of combustion and air pass from the individual hot and cold nozzles, into a common nozzle before final ejection into the atmosphere.

The invention is particularly applicable to constructions in which the blades of the aft fan are supported from the blades of a turbine rotor. In such arrangements as hitherto proposed a compromise had to be accepted between the tip speed requirements of the fan and those ice of the turbine rotor which drives it, and to minimise the effects of this compromise the diameters of the rotors have been made relatively large so as to embody high hub/tip ratios. This has an adverse effect upon weight and drag, and an object of the present invention is to provide an arrangement which has advantages over such known arrangements.

To this end a jet propulsion engine according to the present invention comprises a rotary air compressor arranged to deliver air to combustion chamber apparatus, a first turbine rotor arranged to be acted upon by the products of combustion from the combustion chamber apparatus and to drive the compressor, and an aft fan assembly comprising second and third turbine rotors arranged so that the products of combustion, after acting on the first turbine rotor act on the second and third turbine rotors successively, gearing connecting the second and third turbine rotors so as to determine the relationship between their speeds, and a ducted fan connected to the second or third turbine rotor.

The fan blades are conveniently supported from the blades of one of the turbine rotors, preferably the third turbine rotor (eg by a ring to which the tips of the blades of the turbine rotor in question are connected). The fan and turbine blades may be axially displaced from one another and arranged on opposite ends of the ring, which may be of frusto conical form. The construction and arrangement of the gearing connecting the second and third turbine rotors is preferably such that these turbine rotors rotate in opposite directions.

The gearing may conveniently lie within a casing situated either in front of the second and third turbine rotors and surrounded by ducting through which the products of combustion pass from the first turbine rotor to the second and third turbine rotors, or situated in rear of the third turbine rotor, and surrounded by ducting through which the products of combustion flow to the hot propulsion nozzle or nozzles through which the products of combustion are ejected.

Two constructions according to the invention respectively having the gearing referred to disposed in the two alternative situations mentioned above, are shown diagrammatically by way of example in the accompanying drawings, in which:

FIGURE 1 is a sectional side elevation in a plane containing the axis of rotation showing part of the engine embodying the three turbine rotors but not including the parts of the engine embodying the air compressor and combustion chamber apparatus since these latter parts may be constructed and arranged in conventional manner, only the upper half of the total crosssection of the assembly being shown, since the lower half would be of identical form, and

FIGURE 2 is a similar view to FIGURE 1 of another construction according to the invention.

In the construction shown in FIGURE 1 the aft fan assembly is to be assumed to be attached to the rear end of a conventional combustion turbine comprising an axial fiow compressor arranged to deliver air to combustion chamber apparatus from which the products of combustion act on a first turbine rotor--indica.ted at 1connected directly to the rotor of the compressor rotor by a shaft 1A. The aft fan unit shown comprises an annular duct 2 into which the products of combustion flow from the exit 3 of the turbine 11, the rear end of this annular duct 2'. opening through guide vanes, indicated at 4, into the inlet of turbine apparatus comprising a casing 2A containing a second turbine rotor 5, a third turbine rotor 7 from which the products of combustion flow into a jet pipe Surrounding and spaced from the casing 2A and jet pipe 8 is a shroud 9, forming with the casing 2A and jet pipe 8 an annular air duct in which lie and can rotate a ring of fan blades 10 lying between casing-support struts 10A and fixed vanes 10B and axially displaced forwardly from the third turbine rotor 7 the inner end of the blades 10 being mounted on the forward end of a fursto conical ring 7A the rear end of which is mounted on the outer ends of the blades of the rotor '7. This has the advantage that it facilitates the thermal expansion and contraction of the blades 7 which are subject to relatively high thermal stresses, relatively to the ring supporting the fan blades which may remain relatively cool and therefore impose a heavy resistance to thermal expansion of the blades 7 as would occur in constructions where the fan blades were mounted on a ring carried by the tips of the turbine blades 7 and lying in the same plane as the blades 7, unless special provision were made to minimize this effect.

The second and third turbine rotors 5 and 7 are mounted respectively on

concentric shafts

12 and 13 which extend forwardly from the rotors into the interior of a casing 14 disposed within the inner wall of the annular duct 2 within which casing the shafts l2 and 13 are connected by gearing comprising a central gear wheel 15 mounted on the shaft 12, a series of intermediate layshafts carrying gears 16 supported in the casing 14, and an outer toothed ring 17 connected to the adjacent end of the shaft 13. This gearing thus constitutes interconnecting gearing between the second and third turbine rotors 5 and 7 by which they are caused to rotate in opposite directions at speeds having a predetermined relationship to one another and with the turbine rotor 5 revolving in this example at substantially higher speed than the turbine rotor 7. The fan blades 10 are therefore driven by power derived partly from the turbine rotor direct and partly from the turbine rotor 5 through the

gearing

15, 16, 17.

In the alternative construction shown in FIGURE 2 the aft fan assembly is also to be assumed to be attached to the rear end of a conventional jet propulsion engine comprising an axial flow compressor arranged to deliver air to combustion chamber apparatus from which the products of combustion act on a first turbine rotor, indicated at 18, similar to the turbine rotor 1 shown in FIGURE 1 connected directly to the compressor rotor by a shaft 18A. The aft fan unit comprises a short annular duct 19 into which the products of combustion flow after acting on the turbine rotor 18, this annular duct containing guide vanes 20 by which the products of combustion are directed into the inlet of turbine apparatus comprising a casing shown generally at 21 containing a second turbine rotor 22 lying in advance of a third turbine rotor 23 from hich the products of combustion pass into a jet pipe 2 terminating in a hot propulsion nozzle 25.

Surrounding and spaced from the casing 21 and jet pipes 24 is a shroud 26 forming with the casing 21 and the jet pipe 24 an annular air duct, indicated generally at 27, in which lie and can rotate a ring of fan blades 28 lying in advance of fixed vanes 29 extending across the duct 27. Extending across the duct 27 and the jet pipe 24 in rear respectively of the guide vanes 29 and the turbine rotor 23 are tubular supporting members 30 connected at their inner ends to a gear casing 31 and enclosed within streamlined fairings indicated at 32, 33.

The duct 29 terminates in a cold propulsion nozzle 34 surrounding the hot propulsion nozzle 25.

Annular support 35 for the ring of fan blades 28 is carried by a frusto-conical supporting ring 36, which in turn is carried by the blades of the turbine rotor 23 as indicated.

The turbine rotor 22 is mounted upon a shaft 37 while the turbine rotor 23 is mounted upon a shaft 38, the two

shafts

37 and 38 being concentric and extending into the casing 31 which constitutes a gear casing containing gearing comprising a sunwheel 39 mounted upon the shaft 37, an internally toothed annulus 40 mounted upon the 4 shaft 38, and planet wheels 41 supported in bearings 41A from the casing 31 and each meshing with the sunwheel 39 and the toothed annulus 40.

It will thus be seen that the power transmitted to the fan blades 28 is derived partly from the turbine rotor 22 via the gearing 3 9, 41, 40, the shaft 38, the turbine rotor 23, and the ring 36, and partly directly through the ring 36 from the turbine rotor 23, and that the relationship between the speeds of the turbine rotors 22 and 23, which rotate in opposite directions, is determined by the gearing 39, 2-1, 40.

What is claimed is:

1. A jet propulsion engine comprising combustion chamber apparatus, a rotary air compressor including a rotor arranged to deliver air to the combustion chamber apparatus, a first turbine rotor arranged to be acted upon by the products of combustion from the combustion chamber apparatus and to drive the rotor of the air compressor, and an aft-fan assembly comprising second and third turbine rotors arranged so that the products of combustion, after acting on the first turbine rotor, act on the second turbine rotor and then directly on the third turbine rotor, said second and third turbine rotors being arranged to rotate in opposite directions, gearing connecting said second and third turbine rotors to determine the relationship between their speeds, and a ducted fan assembly including a set of static blades and a set of rotating blades, said rotating blades being driven from said second and third turbine rotors and connected directly to one of said turbine rotors.

2. A jet propulsion engine as claimed in claim 1 in which said ducted fan is connected directly to said third turbine rotor and is connected through said gearing to said second turbine rotor.

3. A jet propulsion engine as claimed in claim 1 in which said ducted fan comprises fan blades supported from the blades of the said third turbine rotor.

t. A jet propulsion engine as claimed in claim 1 in which said ducted fan comprises fan blades supported from the blades of the turbine rotor to which it is directly connected.

5. A jet propulsion engine as claimed in claim 1 wherein the direct connection between the said fan blades and the blades of the turbine rotor to which they are directly connected is constituted by an axial extending ring one end of which is connected to the said fan blades While its other end is connected to the blades of said turbine rotor.

6. A jet propulsion engine as claimed in claim 1 including ducting by which the products of combustion flow from the first turbine rotor to the second and third turbine rotors and a casing containing said gearing and lying within said ducting in advance of said second and third turbine rotors.

7. A jet propulsion engine as claimed in claim 1 including a propulsion nozzle, ducting by which the product of combustion flow from the third turbine rotor to said propulsion nozzle and a casing enclosing said gearing and lying within the said ducting in rear of said second and third turbine rotors.

References Cited UNITED STATES PATENTS 2,478,206 8/1949 Redding 60--226 2,505,660 4/1950 Baumann 60--226 3,282,053 11/ 1966 Messerschmitt 60226 FOREIGN PATENTS 540,171 8/1955 Belgium. 586,557 1/1943 Great Britain.

CARLTON R. CROYLE, Primary Examiner.

MARK E. NEWMAN, Examiner.

D. HART, Assistant Examiner.