Power plants – Combustion products used as motive fluid – Combustion products generator
Reexamination Certificate
1999-09-17
2003-03-11
Freay, Charles G. (Department: 3746)
Power plants
Combustion products used as motive fluid
Combustion products generator
C060S039826, C060S748000
Reexamination Certificate
active
06530223
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to combustors in gas turbine engines and, in particular, to a gas turbine engine combustor having a pilot dome oriented in substantially perpendicular relation to a main dome.
It will be appreciated that emissions are a primary concern in the operation of gas turbine engines, particularly with respect to the impact on the ozone layer by nitrous oxides (NOx), carbon monoxide (CO), and hydrocarbons. In the case of supersonic commercial transport aircraft flying at high altitudes, current subsonic aircraft technology is not applicable given the detrimental effects on the stratospheric ozone. Accordingly, new fuel injection and mixing techniques have been and continue to be developed in order to provide ultra-low NOx at all engine operating conditions.
One combustion system, known as a dry low emission (DLE) combustor, premixes fuel and air in a manner so that the fuel-air ratios are below stoichiometric levels (also known as “lean”). The DLE combustor is described in greater detail in U.S. Pat. Nos. 5,675,971 and 5,680,766, for example, and falls generally within a class of gas turbine engine combustors known as lean, premixed, prevaporized (LPP). While the DLE combustor is able to produce ultra-low NOx across a broad range of conditions for stationary land-based operations, it is a heavy and relatively complex system. Thus, such DLE design was found to be unacceptable for use in aircraft engines due to cost and weight considerations.
Further, a key component found to provide extremely low levels of NOx at moderate to high power conditions for such aircraft engine was the use of a series of simple mixing tubes as the main fuel injection source. It was found, however, that flame stability and emissions characteristics of a combustor incorporating only such mixing tubes was less capable at low power. Thus, it was determined that an independent pilot fuel injector system would be beneficial for such combustor to improve low power flame stability and meet landing-takeoff (LTO) and idle cycle emissions requirements.
The use of combustion staging has been in practice within the gas turbine engine art for many years to expand the operational range of combustion systems, as well as to provide a broad range of gas turbine power output and applicability. This has typically been accomplished by staging the fuel in a plurality of fuel air mixing devices or modulating the mixing devices independently. In addition, air staging has been performed by having separate and/or isolated annular or cannular combustion zones that can be controlled independently to provide low emissions and a broad range of operation. To date, however, such staging by pilot and main combustion zones has been within substantially the same annular plane.
In light of the foregoing, it would be desirable for a gas turbine engine combustor to be developed which provides ultra-low emissions during all operating conditions. It would also be desirable for such combustor to be simple in construction so as to minimize weight and cost, as well as fit within size parameters available for existing gas turbine engine combustors.
BRIEF SUMMARY OF THE INVENTION
In an exemplary embodiment of the invention, a combustor for a gas turbine engine having a longitudinal axis therethrough is disclosed as including an outer liner having a forward end and an aft end, an inner liner having a forward end and an aft end, a first dome formed upstream of the outer liner forward end so as to define a first combustion zone radially oriented to the longitudinal axis, and a dome plate having an outer portion connected to an upstream end of the first dome and an inner portion connected to the inner liner forward end, wherein a second combustion zone is defined by the dome plate, the outer liner, and the inner liner substantially perpendicular to the first combustion zone.
Further, a plurality of circumferentially spaced fuel air mixers are positioned with respect to a corresponding segment of the first dome so as to provide a swirled fuel air mixture into the first combustion zone. Likewise, a plurality of fuel air mixers are positioned upstream of the dome plate for providing an unswirled fuel air mixture into the second combustion zone. In this way, a vortex flow created in the first combustion zone moves radially inward to mix with the axial flow injected into the second combustion zone. Preferably, the axial flow injected through the dome plate is aligned with an aft component of the vortex flow.
REFERENCES:
patent: 5444982 (1995-08-01), Heberling et al.
patent: 5490380 (1996-02-01), Marshall
patent: 5540056 (1996-07-01), Heberling et al.
patent: 5596873 (1997-01-01), Joshi et al.
patent: 5619855 (1997-04-01), Burrus
patent: 5682747 (1997-11-01), Brown et al.
patent: 5749219 (1998-05-01), Dubell
patent: 5791148 (1998-08-01), Burrus
“HSCT Computer Model Takes Shape at NASA,” pp. 68-76, by James Ott, Hampton, VA, in Aviation Week & Space Technology, Oct. 13, 1997.
Dodds Willard J.
Halila Ely E.
Heberling Paul V.
Taylor Jack R.
Andes William Scott
Freay Charles G.
Hess Andrew C.
Rodriguez William
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