Cycle gas turbine engine

Details Cycle gas turbine engine Cycle gas turbine engine

2002-07-29

2004-03-09

Casaregola, Louis J. (Department: 3746)

Power plants

Combustion products used as motive fluid

Multiple fluid-operated motors

C060S039170

Reexamination Certificate

active

06701717

ABSTRACT:

FIELD OF THE INVENTION
The invention relates to gas turbine engines and particularly to variable cycle gas turbine engines for aero applications.
BACKGROUND OF THE INVENTION
High efficiency gas turbine engines for aero applications require a high overall pressure ratio across their compressors, in order to minimise fuel consumption. A typical overall pressure ratio for such an engine might be greater than 40:1 and this results in a need for compressors with high numbers of stages and high pressure ratios across each stage. The compressors tend to be designed for their highest efficiency at high speeds, but this may cause difficulties at lower speeds.
If the operating conditions imposed upon a compressor blade depart too far from the design intention, breakdown of airflow and/or aerodynamically induced vibration may occur. If the engine demands a pressure rise from the compressor which is higher than the compressor can sustain, “surge” occurs. This involves an instantaneous breakdown of flow through the engine and the expulsion of high pressure air from the combustion system forward through the compressor. This produces a loud “bang” and results in a loss of engine thrust.
Compressors are designed with an adequate margin to ensure that the unstable area where surge may occur is avoided. The margin between the unstable area and the working line of a particular compressor is known as the surge margin.
For compressors designed for high efficiency at high speeds, the surge margin at low speeds tends to be too low. In order to overcome the low surge margin, various adjustments may be made to the compressor at such low speeds, as follows.
One method of improving the surge margin at low speeds is to introduce airflow control into the compressor design. This may take the form of variable inlet guide vanes for the first stage and variable stator vanes for subsequent stages. As the compressor speed is reduced from its design value, these static vanes are progressively closed in order to maintain an acceptable air angle onto the following rotor blades. Alternatively or additionally air may be bled off the compressor in order to reduce its pressure ratio. This tends to increase the surge margin to an acceptable level. However, this has an adverse effect on fuel consumption.
Other known ways of providing an acceptable surge margin involve using additional compressor stages in order to reduce the pressure ratio across each stage or using lower compressor efficiencies. The former solution is expensive, whilst the latter results in poor fuel consumption.
SUMMARY OF THE INVENTION
According to one aspect of the invention there is provided a variable cycle gas turbine engine including first and second compressors, combustion apparatus and first and second turbines operable to drive the first and second compressors respectively via interconnecting shafts, wherein means are provided for varying the capacity of at least one of the turbines.
Probably the first turbine operates at a higher pressure than the second turbine, the first compressor operates at a higher pressure than the second compressor, and means are provided for varying the capacity of the second turbine.
The engine may further include a third turbine operable to drive a third compressor or fan, the third turbine and compressor or fan operating at lower pressures than the first and second turbines and compressors.
The means for varying the capacity of the turbine may include means for reducing the pressure ratio across the turbine, to reduce the pressure ratio across the compressor which it drives.
The means for varying the capacity of the turbine may include means for varying the work done by the turbine and the torque applied by the turbine to the shaft connecting the turbine to the compressor which it drives.
The variable capacity turbine may include a stator vane assembly for directing air onto a turbine rotor assembly of the turbine, the stator vane assembly comprising an annular array of substantially radially extending stator vanes circumferentially spaced apart so as to define throat areas therebetween, the stator vanes being adjustable to vary the throat areas between adjacent stator vanes. The stator vanes may be continuously adjustable so that the throat areas are continuously variable between maximum and minimum values.
The stator vanes may be adjustable to vary the angle at which air passing through the throat areas impacts the turbine rotor assembly.
Each stator vane may be pivotable about an axis which extends generally in a radial direction of the vane. Alternatively each stator vane may include a substantially fixed portion and a movable portion, the movable portion being pivotable relative to the fixed portion to vary the throat areas between adjacent stator vanes. The movable portion of each stator vane may be pivotable about an axis which extends generally in the radial direction of the vane.
The gas turbine engine may include means for bleeding air from one or more of the compressors and feeding such air into a region downstream of and in fluid communication with the second turbine, thereby to vary the capacity of the second turbine.
A stator vane downstream of the second turbine may be hollow and may include a plurality of orifices passing through a wall thereof, and means may be provided for passing air into the inside of the stator vane to allow ejection of the air through the orifices to a region downstream of the second turbine.
According to another aspect of the invention, there is further provided a method of operating a gas turbine engine as previously defined wherein when the engine runs at low power, the capacity of the variable capacity turbine is increased, thereby reducing the pressure ratio across the variable capacity turbine and reducing the pressure ratio across the compressor which it drives. As the pressure ratio across the variable capacity turbine is reduced, the pressure ratio across a turbine upstream thereof may increase, thereby increasing the pressure ratio across a compressor driven thereby.


REFERENCES:
patent: 3041822 (1962-07-01), Embree
patent: 3070131 (1962-12-01), Wheatley
patent: 3385509 (1968-05-01), Garnier
patent: 3751909 (1973-08-01), Kohler
patent: 856 036 (1952-09-01), None
patent: 0 247 984 (1987-12-01), None
patent: 490 198 (1938-08-01), None
patent: 701 324 (1953-12-01), None
Treager, Irwin E. Aircraft Gas Turbine Engine Technology. McGraw-Hill, 1970; p. 10.

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