Title:
Contra rotatable turbine system
Kind Code:
A1


Abstract:
A three shaft ducted fan gas turbine engine (10) has a turbine system comprising a high pressure turbine rotor (32), an intermediate pressure turbine rotor (36) and first and second low pressure turbine rotors (42,44), that straddle intermediate pressure turbine rotor (36). Rotors (42,44) are mechanically connected by a casing (46) that bridges turbine rotor (36) so as to achieve transmission of torque from the first to the second thereof. Rotor (42) also delivers gas flow to the intermediate pressure rotor (36) at a much lower temperature than is experienced by conventional arrangements.



Inventors:
Harvey, Neil W. (Derby, GB)
Rose, Martin G. (Zurich, CH)
Application Number:
10/965754
Publication Date:
10/13/2005
Filing Date:
10/18/2004
Assignee:
ROLLS-ROYCE PLC (London, GB)
Primary Class:
International Classes:
F01D5/03; F02C3/067; (IPC1-7): F01D1/02
View Patent Images:
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Primary Examiner:
CASAREGOLA, LOUIS J
Attorney, Agent or Firm:
OLIFF PLC (ALEXANDRIA, VA, US)
Claims:
1. A ducted fan gas turbine engine turbine system comprising, in flow series, a high pressure bladed turbine rotor, a fixed, bladed stator stage, a first low pressure bladed turbine rotor, an intermediate pressure bladed turbine rotor and a second low pressure bladed turbine rotor, the two low pressure bladed turbine rotors being mechanically connected for co-rotation in a direction contrary to that of the high pressure and intermediate pressure turbine rotors, said contra rotation being effected, at least in part, by appropriate shaping of the gas flow paths through said stage of stator blades.

2. A ducted fan gas turbine engine turbine system as claimed in claim 1 wherein said mechanical connection comprises a casing fixed on the outer tip extremities of the blades of each low pressure turbine rotor and bridging said intermediate pressure turbine rotor.

3. A ducted fan gas turbine engine turbine system as claimed in claim 1 wherein in situ in a ducted fan gas turbine engine, said high pressure turbine rotor is drivingly connected via a shaft to a high pressure compressor, said intermediate pressure turbine rotor is drivingly connected via another shaft to an intermediate pressure compressor immediately upstream of said high pressure compressor and said second low pressure turbine rotor is directly drivingly connected via a further shaft to a ducted fan at the upstream end of said engine.

4. A ducted fan gas turbine engine turbine system as claimed in any of claims 1 wherein said intermediate pressure turbine rotor is effectively mechanically connected to said first low pressure turbine rotor so as to impart a rotary force thereon during operation in a said engine.

5. A ducted fan gas turbine engine turbine system as claimed in claim 4 wherein said mechanical connection comprises a geared connection between the intermediate turbine rotor shaft and said first low pressure turbine rotor.

6. A ducted fan gas turbine engine turbine system as claimed in claim 5 wherein said geared connection comprises a spur gear on said intermediate turbine rotor shaft that engages a plurality of planet gears supported from fixed structure and which in turn engage a spur gear on said first low pressure turbine rotor.

7. A ducted fan gas turbine engine turbine system as claimed in claim 4 wherein said mechanical connection comprises a geared connection between the upstream end of an intermediate compressor and a ducted fan when all are assembled.

8. A ducted fan gas turbine engine turbine system as claimed in claim 5 wherein said connection of said gears is made via planet gears supported from fixed structure.

Description:

The present invention relates to turbine systems of the kind having efficacy in aircraft powered by ducted fan gas turbine engines.

Present ducted fan gas turbine engine turbine systems comprise respective high pressure, intermediate pressure and low pressure turbine stages arranged in flow series and spaced from each other by respective stages of fixed stators. The turbine stages rotate one each of two or more shafts in a common direction, respective shafts being fixed to compressors or a ducted fan.

There is documentary evidence e.g. G.B. patent 2,129,502 and U.S. Pat. No. 5,274,999, that consideration has been given to multi shaft, multi stage turbine systems, in which a fan is driven by low pressure stages that are, in all cases, downstream of the core gas turbine system, and rotate in a direction contrary to the remainder of the turbine stages. However, none of the turbine systems devised have proved sufficiently advantageous to warrant their manufacture and use.

The present invention seeks to provide an improved ducted fan gas turbine engine multi stage turbine system, low pressure stages of which rotates in a direction contrary to the remainder thereof.

According to the present invention a ducted fan gas turbine engine turbine system comprises, in flow series, a high pressure bladed rotor, a fixed, bladed stator stage, a first low pressure bladed rotor, an intermediate pressure bladed rotor and a second low pressure bladed rotor, the two low pressure rotors being mechanically connected for co-rotation in a direction contrary to that of the high pressure and intermediate pressure rotors, said contra rotation being effected by appropriate shaping of the gas flow paths through said stage of stator blades.

The invention will now be described, by way of example and with reference to the accompanying drawings, in which:

FIG. 1 is an axial cross section through a ducted fan gas turbine engine including a turbine system in accordance with one aspect of the present invention.

FIG. 2 is an axial cross section through a ducted fan gas turbine engine ncluding a turbine system in accordance with a further aspect of the present invention.

FIG. 3 is an axial cross section through a ducted fan gas turbine engine including a turbine system having an alternative fan drive system in accordance with the present invention.

Referring to FIG. 1. A ducted fan gas turbine engine, generally indicated by the numeral 10, has a short fan cowl 12 that surrounds a stage of fan blades 14. A casing 16 supports the fan cowl 12 via struts 18, and also surrounds the core gas generator 20.

The core gas generator 20 consists of the following: An intermediate pressure compressor 22, that via fixed vanes 24, receives and further pressurises ambient air after it has been slightly pressurised on passing between the root portions 26 of the rotating fan stage 14. There follows a high pressure compressor 28 that further compresses the air on its receipt from the intermediate compressor 22, and delivers some of it at a still higher pressure to combustion equipment 30 wherein it is mixed with fuel and burned. A high pressure turbine roter 32 receives the resulting hot gases and is caused to rotate in a given direction and, being directly connected to high pressure compressor 28 via a shaft 34, rotates it in the same direction. The final part of the core gas generator 20 is an intermediate pressure turbine rotor 36 which eventually receives the gases from the high pressure rotor 32 at a pressure lower than that experienced by high pressure rotor 32. The gas flow path is such that the intermediate pressure turbine rotor 36 rotates in the same direction as the high pressure turbine rotor 32. Intermediate pressure turbine rotor 36 is directly connected to intermediate pressure compressor 22 via a shaft 38 and consequently rotates intermediate compressor 22 in the same direction.

Fan stage 14 is not mechanically connected to the core gas generator in any way for the purpose of rotation thereby, but derives its rotary motion via a shaft 40 that is directly connected to the downstream one, 44, of two low pressure power turbine rotors 42 and 44. By “downstream” is meant with regard to the direction of flow of gases through the engine. Turbine rotors 42 and 44 are mechanically connected via an outer casing 46 that co-rotates with them.

Turbine rotors 42 and 44 straddle intermediate turbine rotor 36, rotor 42 being positioned upstream thereof, and consequently is the first turbine rotor to receive the flow of gas that has passed through high pressure turbine 32, and thereafter through a stage of fixed, wide chord stators 48. The stators 48 are shaped so as to define flow paths between them which will increase the exit whirl of the gases leaving the stators 48 and thereby reduce the angle of deflection of the gases relative to that imposed by conventional stators, and thus makes for a more aerodynamically efficient stator stage.

The low pressure turbine rotor 42 is arranged so that on receiving the flow from stator stage 36, it counter rotates relative to the high pressure turbine rotor 32. However, the intermediate pressure turbine rotor 36 is arranged so that, on receiving the flow from low pressure turbine rotor 42, it rotates in the same direction as the high pressure turbine rotor 32. The second low pressure turbine rotor 44, is rotated by torque transmitted by low pressure turbine stage 42, via casing 46 and, to some extent, by the gases exiting the intermediate turbine stage 36.

During the cruise regime of an aircraft powered by a known three shaft ducted fan engine, in which all the turbines and their respective shafts rotate in a common direction, the ratio of the speeds of the high pressure turbine shaft, the intermediate pressure turbine shaft and the low pressure turbine shaft is, nominally, 11:7:3. Thus, the speed of the low pressure shaft is less than half that of the intermediate shaft speed. However, in the three shaft arrangement of the present invention, whilst the ratio of the speeds of the high and intermediate pressure shafts will nominally remain the same as the prior art engine, the speed ratio between the intermediate pressure shaft and low pressure shaft will increase to, nominally, 10, by virtue of their effected counter rotation. This may necessitate the substitution of prior art quality inter shaft bearings (not shown) between intermediate pressure shaft 38 and low pressure shaft 40, by bearings (not shown) of the same quality as those used to support high pressure shaft 34 relative to surrounding fixed structure in known manner.

During operation of an associated engine, the contra rotation of intermediate pressure turbine rotor 36 and low pressure turbine rotor 42 results in the speed of rotation of the blades of intermediate pressure rotor 36 being much higher than that of the blades of low pressure rotor 42. A very high exit whirl velocity of the gases leaving the blades of low pressure rotor 42 is achieved, and still maintain good inlet conditions into intermediate pressure rotor 36. By this means, a large fraction, in the order of 40-50% of the total power generated by the low pressure system of turbine rotors 42 and 44, can be generated by turbine rotor 42.

The FIG. 1 example of the present invention depicts a further, conventional low pressure turbine rotor 50 with stators 52. The inclusion of this or any other number of conventional low pressure turbine rotors will depend on the required power regime of the ducted fan gas turbine engine, and the example should not be regarded as limitive.

A number of advantages may be derived from the turbine system of the present invention as follows:

    • a) No stator stages are utilised either side of the intermediate pressure turbine rotor 36, thus achieving a large cost saving.
    • b) The stator inlet to the low pressure turbine rotor 42 becomes a low deflection aerofoil, thereby improving its aerodynamic efficiency.
    • c) Low pressure turbine rotor 42 and intermediate pressure turbine rotor 36 both have velocity ratios that are much higher than conventional turbines, thus helping to achieve high aerodynamic efficiency.
    • d) The respective frequencies of the aerodynamic interaction between the intermediate pressure turbine rotor and the low pressure turbine rotors 42 and 44 are much higher than in conventional designs. This helps to keep noise generated by the interactions out of the audible range, and reduces the perceived noise of the turbine system.
    • e) A substantial reduction in the number of low pressure turbine blades is achieved where the values of aerofoil coefficients and axial chord proportions are similar to those used in a conventional turbine design, e.g. the reduction could equal about 30% relative to the number of blades utilised in a five stage conventional low pressure turbine.
    • f) Flowing the hot gases through stators 48 and low pressure turbine rotor 42, prior to it reaching intermediate pressure turbine rotor 36, results in the gas experiencing a considerable drop in temperature. The intermediate pressure turbine rotor 36 is subjected to considerably lower temperatures than in conventional designs. This invention does not reduce its stress. Moreover, higher stress levels can be allowed in the IP rotor. This is useful because in the current invention the flow will have expanded more, relative to conventional, by the time it leaves the IP rotor. To achieve an optimum design, the flow area of the IP rotor needs to be larger than conventional, which results in higher stress.

The first low pressure rotor 42 will experience a higher temperature than in a conventional design. However, since it is on the LP shaft which rotates at the lowest speed, the levels of stress are low and designs can be achieved using existing materials without the need for cooling.

Referring to FIG. 2. In this second embodiment of the present invention, the turbine system is generally as that described with reference to FIG. 1. However, the intermediate pressure shaft 38 is provided with a gear 54 about its periphery. Gear 54 is engaged by planet gears 56 that are supported by fixed structure 57 passing through stators 48. Low pressure turbine rotor 42 is provided with a gear ring 58 that also engages planet gears 56. The ratio of gears can be determined using known gear design technology.

The utilising of a set of gears 54, 56 and 58 enables the intermediate pressure turbine to extract from the gas flow, the power to drive the intermediate compressor 22, and the power to drive the low pressure turbine rotor 42 through the gears. There results an increase in the whirl velocity in the gas flow leaving the intermediate pressure turbine rotor, thereby improving the inlet conditions into the low pressure turbine rotor 44, which is then able to extract more power from the gas flow than is possible in the first embodiment described herein, while still achieving acceptable aerodynamic efficiency.

The FIG. 2 embodiment of the present invention will require fewer low pressure turbine rotors than the embodiment of FIG. 1. Further, if the values of blade lift coefficients and axial chord proportions are similar to those in a conventional low pressure turbine rotor design, an even greater reduction of approximately 60% in the number of low pressure turbine blades can be achieved. Moreover, the number of blades in the intermediate pressure turbine rotor 36 can be reduced by typically 10%.

The operation of the FIG. 3 embodiment in a relatively cool part of the engine enables cooling of the gear system to be more easily effected.

In this embodiment, the geardrive is located at the front of the engine, between the fan 14 and the intermediate compressor 22, and the fan disc is provided with a gear ring 158. Gear ring 158 is engaged by planet gears 156 that are supported by fixed structure 157 passing through the fixed vanes 24 at inlet to the IP compressor 22. The IP compressor 22 is provided with another gear ring 154 that also engages planet gears 156. As before, the ratio of gears can be determined using known gear design technology.





 
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