Power plants – Combustion products used as motive fluid – Combustion products generator
Reexamination Certificate
2001-10-26
2004-05-18
Kim, Ted (Department: 3746)
Power plants
Combustion products used as motive fluid
Combustion products generator
C415S115000, C416S095000
Reexamination Certificate
active
06735956
ABSTRACT:
TECHNICAL FIELD
The invention relates to a gas turbine engine having a single stage high work high pressure turbine with blade cooling scoops to intake cooling air from an adjacent plenum for blade cooling.
BACKGROUND OF THE ART
The invention relates to cooling systems for the turbine blades of a gas turbine engine and in particular an improved cooling air supply system for an engine that has a single stage high work high pressure turbine. For such a turbine there is a high pressure drop between the vane inlet to the blade outlet. For this reason, the blade inlet and outlet gas path pressure is lower than for other types of turbines and as a result cooling air can be taken from a low pressure stage of the engine rather than the conventional high pressure stage. Conventional engines include rather complex inlet structures including tangential onboard injection, turbine cover plates with impeller surfaces and associated seals, all of which add to the mechanical complexity of the system.
Under elevated operating conditions, gas turbine engine components, such as turbine rotors and blades are conventionally cooled by a flow of compressed air discharged at a relatively cool temperature. The flow of cooling air across the turbine rotor and through the interior of the blades removes heat through heat exchange so as to prevent excessive reduction of the mechanical strength properties of the turbine blades and turbine rotor. The operating temperature, efficiency and output of the turbine engine are limited by the high temperature capabilities of various turbine elements and the materials of which they are made. The lower the temperature of the elements, the higher strength and resistance to operating stress. However, the performance of the gas turbine engine is also sensitive to the amount of air flow that is used for cooling the hot turbine components. If less air is used for cooling functions, the efficiency and performance of the engine improves. As well, aircraft engine efficiency is very sensitive to weight penalties imposed by use of complex cooling system components.
To cool turbine rotor blades conventionally a flow of high pressure cooling air is introduced at a low radius close to the engine's centre line axis. An example of a conventional cooling air system for a turbine is shown in U.S. Pat. No. 5,984,636 to Fahndrich et al. The cooling air flow is introduced with a swirl or tangential velocity component through the use of tangential on board injectors with nozzles directed at the rotating hub of the turbine rotor. The cooling air is then passed between a turbine cover plate and the surface of the turbine rotor to simultaneously cool the rotor hub and to increase the pressure through centrifugal pumping of the air as it is conducted to the blades on the periphery of the turbine hub. The requirement for tangential on board injectors, cover plates and associated running seals significantly increases the mechanical complexity of the turbine blade cooling system. However, since conventional engines have hot gas path pressures that are relatively high, the pressure of the cooling air must exceed the hot gas path pressure. Due to the high pressure of the hot gas path, conventionally it is necessary to intake high pressure cooling air and increase the pressure of the air through onboard injection and impeller action of the cover plate in order to ensure that sufficient air flow is conducted through the turbine blades for cooling purposes and exits the trailing edge of the turbine blade into the hot gas path.
It is an object of the present invention to reduce the efficiency penalty of conventional cooling air system for turbine blades by utilising low-pressure air from the low-pressure compressor stage.
It is a further object of the invention to eliminate the mechanical complexity of conventional cooling systems by eliminating tangential onboard injection, cover plates and seals
It is a further object of the invention to utilize a high work single stage high-pressure turbine with gas path pressure lower than turbines conventionally used to enable use of low-pressure sources for cooling air for the turbine blades.
Further objects of the invention will be apparent from review of the disclosure, drawings and description of the invention below.
DISCLOSURE OF THE INVENTION
The invention provides a gas turbine engine having a single stage high work high-pressure turbine with unique blade cooling scoops. The turbine blades include a cooling air inlet duct communicating with a cooling air plenum with pressure above the hot gas path pressure. A blade airfoil extends radially from the root and includes cooling air channels communicating between the cooling air inlet duct and the hot gas path of the engine. The air inlet duct includes an inlet scoop extending into the cooling air plenum with an inlet scoop aperture oriented to capture cooling air from the plenum as a consequence of the turbine rotation. The engine includes a low-pressure compressor stage in flow communication with the cooling air plenum. Advantageously, the engine includes a bearing gallery adjacent the cooling air plenum, where the bearing gallery includes a cooling air jacket in communication with the low pressure compressor stage, and the cooling air jacket communicates with the low pressure cooling air plenum. A labyrinth seal is provided between the hot gas path and the cooling air plenum.
The use of inlet air scoops in conjunction with the high work single stage high pressure turbine is feasible for the following reasons. The high work single stage high pressure turbine has a gas path pressure that is lower than conventional turbines and for this reason low pressure cooling air sources can be adopted. Of course, the cooling air pressure must be at least marginally higher than the gas path pressure in order to ensure that cooling air of sufficient quantity is conducted through the high pressure turbine blades to affect cooling. The invention, greatly simplifies turbine blade cooling systems by providing a low pressure cooling air plenum within which the high pressure turbine rotor rotates. Extending into the cooling air plenum are the blade roots of the turbine blades together with air inlet ducts with inlet scoops oriented to capture cooling air from the plenum as a consequence of the turbine rotation.
Therefore, the invention eliminates tangential onboard injectors, cover plates and associated seals that are conventionally necessary to increase the pressure of cooling air. Since the hot gas path pressure is lower for high work turbines, low pressure air can be drawn through the rotation of the inlet scoops by the rotating turbine within a cooling air plenum supplied by low pressure cooling air from the low pressure stage of the compressor.
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(Ogilvy Renault)
Kim Ted
Pratt & Whitney Canada Corp.
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