Power plants – Combustion products used as motive fluid – With exhaust treatment
Reexamination Certificate
1993-06-23
2003-12-16
Nelson, Peter A. (Department: 3641)
Power plants
Combustion products used as motive fluid
With exhaust treatment
C060S039183, C342S004000
Reexamination Certificate
active
06662546
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to aircraft gas turbine engines with low observable fan design features which reduce the radar signature of the engine and its aircraft and more particularly to fixed angle high swirl inlet guide vanes coated with radar absorbing material.
2. Description of Related Art
Gas turbine engines generally comprise a compressor for compressing air flowing through the engine, a combustor in which fuel is mixed with the compressed air and ignited to form a high energy gas stream, and a turbine for driving the compressor. One type of gas turbine engine for an aircraft is the turbojet in which thrust is provided by the high velocity gas stream exiting the turbine.
A second type of aircraft gas turbine engine is the turbofan in which a fan is mounted forward of the compressor and is driven by a second turbine or power turbine mounted downstream of the first turbine. The fan produces a flow of pressurized air which is split into two portions. The first portion enters an outer bypass duct for bypassing the core engine and the second portion enters the compressor of the core engine. One advantage of the turbofan engine over the turbojet is that it is more fuel efficient for a given thrust output of the engine.
Another feature which may be utilized to increase the thrust output of a gas turbine engine is an augmentor. In an augmented gas turbine engine, an exhaust duct is provided downstream of the turbine(s). Additional fuel is injected into the exhaust duct and is ignited to increase the energy of the gas stream. The gas stream is ejected through an exhaust nozzle to increase the thrust output of the engine. Augmenters for high performance military aircraft are particularly useful and desirable. A typical augmentor disposed in the exhaust duct of the engine requires protection from the extremely high temperatures associated with the gas flow within the augmentor. A cooling liner is typically positioned within the exhaust duct so as to form a cooling duct therebetween and a portion of the bypass flow is diverted into this cooling duct for cooling of the same and to optionally convey cooling air to an exhaust nozzle to cool the nozzle.
A characteristic of turbofan engines, especially low fan pressure ratio (pressure increase across the fan section of the engine) turbofan engines, is relatively low specific fuel consumption at subsonic speeds. A characteristic of turbojet and relatively high fan pressure ratio turbofan engines is relatively high specific thrust characteristics at supersonic speeds. In order to satisfy the need for aircraft which must efficiently operate over a wide range of subsonic and supersonic speeds, so-called variable cycle engines have been developed. Such variable cycle engines are characterized by the capability to change the bypass ratio of the engine during operation. For example. U.S. Pat. Nos. 4,010,608 and 4,961,312 by Simmons and 4,175,384 by Wagenknecht et al disclose variable cycle engines. These variable cycle engines include an outer bypass duct and a variable area bypass injector for modulating the flow through the bypass duct thereby varying the engine cycle.
A fundamental problem in the design and operation of augmented variable cycle engine is the matching of bypass airflow requirements to the airflow requirements of the combustion process in and cooling of the augmentor combustor and the cooling of the augmenter liner. Control of the bypass air has typically been accomplished by variable bypass systems that use bypass doors such as those disclosed in the patents referenced above. The Wagenknecht patent (4,175,384) further discloses a variable area bypass injector (VABI) wherein the VABI comprises generally two variable geometry flow devices, a double-bypass selector valve and a static pressure valve. The VABI in Wagenknecht is included in a double bypass, variable cycle engine for selectively mixing co-annular bypass streams. Further disclosed is a multiple stage fan section having a variable inlet guide vane and variable stator vanes upstream of a mid-stage fan VABI as well as between all fan rotor blades is disclosed for modulating the airflow through the fan section of the engine and the fan VABI. U.S. patent application No. 07/484,083, entitled “Variable Cycle Turbofan Ramjet Engine” by Johnson et al, filed Feb. 22. 1990, and presently allowed discloses the use of an variable stator vane aft of the first by-pass duct.
Modern day military aircraft and particularly combat aircraft have a great need to employ techniques for reducing the radar signature of the aircraft and its gas turbine engine. Such radar signature reducing techniques may also be referred to as low observable or LO techniques. One such LO technique is to use inlet guide vanes having fixed forward portions and variable angle aft portions such that the inlet guide vanes are angled with respect to the line of sight through the inlet so as to prevent direct observation of the rotating fan blades. These conventional RAM coated inlet guide vanes are designed to essentially swirl the inlet flow to hide at least a large part of the fan blades then deswirl the flow before it enters the blades. Such a technique is disclosed in U.S. patent application No. 06/536,792 entitled “Low Radar Cross Section Engine Inlet Axial Line of Sight Blocked” by Giffin et al, filed Sep. 26, 1983, and presently allowed; and essentially blocks all possible linear lines of sight into the rotating equipment and provides the inlet guide vanes with a coating of radar absorbing material or RAM, of which there are several known materials. However, the use of variable stator vanes in the inlet including inlet guide vanes that are fully or partially variable increases the radar signature of the inlet because they introduce more radar observable edges.
Centrifugal forces acting on the fan blades is an obstacle to using RAM coatings on fan blades for additional LO capabilities and performance. One effect of inlet preswirl is to reduce rotor inlet Mach Number which enhances the tolerance to the thickness of coatings on the fan blades. Therefore fixed angle inlet guide vanes are better for radar signature reducing techniques and there is a great need for an aircraft gas turbine engine inlet and fan design incorporating RAM coated fixed angle inlet guide vanes to more effectively reduce the radar signature of aircraft and their gas turbine engines.
SUMMARY OF THE INVENTION
The present invention provides an aircraft gas turbine engine with low observable fan design features which reduce the radar signature of the engine and its aircraft. An aircraft gas turbine engine fan section has in direct serial flow relationship an inlet having fixed angle high swirl angle inlet guide vanes, a first stage of first fan rotor blades, a stage of first variable angle stator vanes, and a preferably mid-stage fan bleed. The inlet guide vanes are angled to block the linear line of sight of the rotating fan blades through the inlet and include a RAM treatment which preferably is a coating of a radar absorbing material on the surface of the inlet guide vanes.
The exemplary embodiment has a multi-stage fan section with the mid-stage fan bleed disposed between the first and second stages of fan rotor blades and the engine is a variable cycle gas turbine engine. The mid-stage fan bleed is used to provide cooling air, preferably to cool a nozzle which in one embodiment has a variable throat. The nozzle's convergent flaps and seals and its divergent flaps and seals may be internally cooled by this cooling air. Aft of the throat area A
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this cooling air may be used to film cool the divergent flaps and/or seals through film cooling slots or holes, as long as the local pressure at that point is sufficiently below the pressure of the cooling air which is substantially that of the fan bleed air so as to allow it to flow into the nozzle flow stream.
ADVANTAGES
The present invention provides an improved means for applying LO techniques to aircraft gas turbine engines that use va
General Electric Company
Herkamp Nathan D.
Nelson Peter A.
Rosen Steven J.
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