Power plants – Combustion products used as motive fluid – With dual function turbine
Reexamination Certificate
2001-04-02
2002-06-04
Koczo, Michael (Department: 3746)
Power plants
Combustion products used as motive fluid
With dual function turbine
C415S079000
Reexamination Certificate
active
06397577
ABSTRACT:
RIGHTS OF THE GOVERNMENT
The invention described herein may be manufactured and used by or for the Government of the United States for all governmental purposes without the payment of any royalty.
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engine structures, and more particularly to an improved engine structure wherein the turbine and compressor aerodynamic components are in close proximity, which reduces or eliminates the need for heavy disks and shafts.
In the operation of a gas turbine engine (Brayton cycle engine and its variants), the working fluid (usually air) is compressed to high pressure within one or more compressor stages. Fuel is then added and the fuel-air mixture is burned in a combustor. The hot high pressure combustion gas is expanded through an expansion system comprised of one or more turbine stages. Power extracted from the flow by the expansion system is used to drive the compression system and any external load (viz, propulsive fan in a turbofan engine or drive shaft in a turboshaft engine). Any pressure remaining after expansion through the turbine system may provide direct (core) jet thrust or power to other expansion systems downstream.
Conventional gas turbine engine systems typically comprise a compression system mounted on one or more rotating compressor disks connected to an expansion system mounted separately on one or more rotating turbine disks. One or more drive shafts interconnect the turbine disks and the compressor disks and transfer the required engine operating power from the turbine components to the compressor components. A linked combination of compressor stages, shaft, and turbine stages that all rotate in unison comprises a spool. An engine may have one or more spools, each free to rotate independently, but because multiple spool engines usually require the shafts to be nested, a practical limit of three spools exists, even for large engines. Most engines comprise two spools, and many only one. The design of compressor and turbine aerodynamics is therefore limited because all the components attached to a given spool must rotate at the same speed.
The compressor and turbine components may be of axial flow, radial flow or mixed flow types (combined axial and radial; the term “mixed flow” applies to a stage in which the mean flow surface radius changes from inlet to exit, such as in the first stage of a fan), but the basic topology is invariant. However, because the aerodynamic components that interact with the flow are physically separated, there is significant engine weight associated with the multiple disks and one or more drive shafts required to interconnect them.
Modern turbine engines typically operate at gas temperatures that exceed the material capabilities of the turbine components. The turbine components are cooled by ducting relatively cool air from the high pressure zone of the compressor past the combustor, and the structure required to duct the cooling air adds substantial weight and complexity to the engine.
The invention solves or substantially reduces in critical importance problems with conventional turbine engine structures by providing novel engine topology that places the turbine and compressor aerodynamic components in close axial and radial proximity to each other, which reduces or eliminates the need for heavy disks and shafts and correspondingly reduces the overall length and weight of the engine. By reducing or eliminating the need for shafts, the number of spools is limited only by the number of turbine stages, and turbine and compressor components on a given spool can be better matched aerodynamically. With the compressor and turbine adjacent to each other, ductwork structure providing cooling to the turbine is significantly shortened, may be fully integrated into the bladed ring components, and may be taken from multiple points in the compressor at the pressure needed, which reduces the adverse cycle impact of turbine cooling. The reduction of inertial masses of the rotating components in the improved engine according to the invention allows more rapid speed change, thereby improving engine responsiveness to commanded thrust or power changes. Because the engine structure according to the invention may have a hollow core, secondary engine systems, such as generators and control electronics may be mounted inside the bore of the engine, thus reducing nacelle size requirements.
The invention may be conveniently incorporated into gas turbine power and propulsion systems requiring minimum size or weight, such as in high performance fighter aircraft, helicopters, portable turbine based power systems and small land and sea vehicle power and propulsion systems. The invention may also be retrofitted into existing turbofan/low bypass turbofan engines to reduce noise and increase power and efficiency.
It is therefore a principal object of the invention to provide an improved gas turbine engine.
It is another object of the invention to provide a gas turbine engine structure of size and weight substantially smaller than conventional gas turbine engines of comparable power output.
It is another object of the invention to provide a gas turbine engine structure having a substantially reduced number of disks or shafts.
It is a further object of the invention to provide a shortened gas turbine engine structure.
It is another object of the invention to provide a gas turbine engine structure having improved cooling and reduced cooling requirements to the turbine stage of the engine.
It is a further object of the invention to provide a gas turbine engine structure having improved responsiveness to power and thrust change commands.
These and other objects of the invention will become apparent as a detailed description of representative embodiments proceeds.
SUMMARY OF THE INVENTION
In accordance with the foregoing principles and objects of the invention, a gas turbine engine structure is described wherein the compressor region and the turbine region of the engine are disposed substantially concentrically of each other between fixed inner and outer casings, with the combustor disposed at one common end of the compressor and turbine, and wherein the rotor components of the engine include one or more rings or bands with the turbine blades and compressor blades mounted respectively on the inner and outer surface thereof, and wherein the compressor and turbine stator components are mounted respectively on the inner surface of the outer casing and the outer surface of the inner casing, or, alternatively, wherein the turbine and compressor stator components are mounted respectively on the inner surface of the outer casing and the outer surface of the inner casing, which structure geometries eliminate much of the weight associated with the disks and interconnecting shafts that characterize conventional engines.
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patent: 577017 (1946-05-01), None
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Kundert Thomas L.
Scearce Bobby D.
The United States of America as represented by the Secretary of
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