Rotary kinetic fluid motors or pumps – With passage in blade – vane – shaft or rotary distributor...
Reexamination Certificate
2000-07-25
2001-08-21
Lopez, F. Daniel (Department: 3745)
Rotary kinetic fluid motors or pumps
With passage in blade, vane, shaft or rotary distributor...
C416S095000, C415S178000, C415S116000
Reexamination Certificate
active
06276896
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates, generally, to gas turbine engines. More particularly, the invention relates to the impeller of an axi-centrifugal compressor for gas turbine engines for aircraft. The invention has particular utility for improving the efficiency of a gas turbine engine by allowing a higher compressor discharge temperature.
2. Background Information
Though it does not depict any existing engine,
FIG. 1
illustrates the current state-of-the-art for axi-centrifugal gas turbine aircraft engines, and is included to provide a frame of reference for the subsequent discussion of prior art and for the present invention. The direction is to the left in the figure, and aft is to the right. Axi-centrifugal gas turbine engines are very compact and efficient. The main airflow goes through a series of axial compressor stages
10
then through the impeller
12
which has a plurality of blades
14
which redirect the flow radially with centrifugal force into diffuser pipes
16
which increase the pressure and reduce the velocity of the airflow as it is redirected toward combustors
18
. In the combustors, the air is mixed with fuel, ignited, and the resulting gas passed through blades of high-pressure turbines
20
. A small portion of the main airflow, called cooling air bleed, is removed from the main airflow in front of impeller
12
and is directed afterward along the hub of the impeller to the high-pressure turbines
20
where it used to cool the blades
21
of the second stage high-pressure turbine before reentering the main airflow stream.
Referring also to
FIG. 2
, the impeller
12
is almost always made of titanium rather than steel due to titanium's higher strength to density ratio, which makes it ideal for rotating machinery components. Furthermore, titanium is much less expensive to purchase and machine than high strength steels. However, at sustained temperatures above 1000° F., the strength of titanium diminishes rapidly with increasing temperature. With current titanium impellers, such as is illustrated in
FIG. 2
, the maximum compressor discharge temperature, usually identified by the symbol T
3
, is limited to 1100° F.
The impeller temperatures are non-uniform. The peak temperature occurs near the rim
24
on the back face
22
at point
26
where radiant heat from the turbines
20
is reflected forward. Also, there is leakage of the hot main airflow around rim
24
onto the back face
22
of impeller
12
, further exacerbating the heating of back face
22
. The temperature at point
26
is approximately 150° F. higher than anywhere else on impeller
12
at the high-power engine conditions. If, at those conditions, the temperature of the impeller at point
26
could be reduced 150° F., that would allow the compressor discharge temperature T
3
to be increased by 150° F. to 1250° F., thereby significantly increasing the overall engine efficiency.
The temperature at point
26
cannot be reduced simply by blowing cooling air at the back face
22
near rim
24
because the main airflow crosses a gap between the rim
24
and the diffuser pipes
16
. The flow parameters across this gap are critical. Any cooling air directed at the back face
22
near rim
24
of impeller
12
would impinge on the main airflow and disrupt that critical flow sufficiently to destroy the effectiveness of the airflow into and through the diffuser pipes
16
, resulting in a drastic reduction of engine efficiency.
U.S. Pat. No. 4,793,772 to Zaehring and U.S. Pat. Nos. 4,920,741 and 4,961,309, both to Liebl, disclose circulating cooling air in a chamber formed outside of the stub shaft to cool the last compressor section of an axial compressor. U.S. Pat. No. 4,808,073 to Zaehring et al. discloses vane-like ribs on the inside of the rear stub shaft which direct cooling air from the center shaft outwardly along the stub shaft and against the outer portion of the last rotor disk. These devices and methods work because the stub shaft connects to the last stage compressor rotor near the rim of the rotor. The shaft that connects to an impeller of an axi-centrifugal compressor connects near its hub rather than its rim, therefore, these devices and methods are not applicable to an impeller for an axi-centrifugal compressor.
The present invention provides an improved impeller for an axi-centrifugal gas turbine and a method of cooling it which reduces the temperature near the outer rim 150° F. over conventional impellers without disrupting the critical airflow between the impeller and the diffuser pipes.
BRIEF SUMMARY OF THE INVENTION
The present invention provides an apparatus and method for cooling an impeller for an axi-centrifugal compressor of a gas turbine engine. A curvic coupling joint is used between the aft end of the impeller and the turbine shaft. A generally annular disk-shaped shield structure is attached to the aft face of the impeller and extends from the impeller rim to the turbine shaft just aft of the curvic coupling. The shield structure is offset from the aft face of the impeller thereby forming a cavity therebetween. The shield structure has a plurality of circumferentially spaced apertures through which cooling air passes to first circulate through the cavity by means of radial vanes to cool the impeller, then exit through the curvic coupling joint. The cooling air is then directed afterward along the turbine shaft to be used to cool high-pressure turbine blades.
The shield structure has a shield portion generally parallel to the flat portion of the aft face of the impeller, and a plenum portion located radially inward from the shield portion. The shield portion and the plenum portion may be made as one unit or two separate components, preferably with a radial overlap between the components. The apertures are located near the juncture of the shield portion and plenum portion.
A plurality of radial vanes extend forward from the shield portion to the flat portion of the aft face of the impeller. The radial vanes are arranged in pairs straddling each aperture with each pair of vanes being joined together at their inner ends by a joining portion to form a U-shape. The joining portion partially surrounding the aperture. The radial vanes have outer ends that terminate radially inward from the rim of the shield structure to allow cooling air to flow around them.
The plenum portion has an inner rim with an aft surface which mates with an axial piloting ring disposed circumferentially on the turbine shaft and located just aft of the curvic coupling. The plenum ring has radial vanes extending forward and inward to aid inward airflow.
Cooling air is extracted from the main airflow at the diffuser exit and routed selectively through a heat exchanger or a bypass of the heat exchanger. At least a portion of the cooling air is injected through the apertures, into the cavity between the shield plate and the aft face of the impeller. It circulates radially outward along the vanes on the shield plate then around the ends of the vanes and radially inward to the cavity between the plenum ring and the aft face of the impeller, then inward through the curvic coupling and on back to the turbines.
The features, benefits and objects of this invention will become clear to those skilled in the art by reference to the following description, claims and drawings.
REFERENCES:
patent: 3490852 (1970-01-01), Carlstrom et al.
patent: 3572996 (1971-03-01), Borden
patent: 3635586 (1972-01-01), Kent et al.
patent: 3644058 (1972-02-01), Barnabei et al.
patent: 3647313 (1972-03-01), Koff
patent: 3748060 (1973-07-01), Hugoson
patent: 3814539 (1974-06-01), Klompas
patent: 3989410 (1976-11-01), Ferrari
patent: 4277222 (1981-07-01), Barbeau
patent: 4674955 (1987-06-01), Howe et al.
patent: 4759688 (1988-07-01), Wright et al.
patent: 4793772 (1988-12-01), Zaehring et al.
patent: 4820116 (1989-04-01), Hovan et al.
patent: 4854821 (1989-08-01), Kernon et al.
patent: 4920741 (1990-05-01), Liebl
patent: 4923370 (1990-05-01), Larson et al.
patent: 4961309 (1990-10-01
Burge Joseph C.
Poire Norman Paul
Dowrey & Associates
Lopez F. Daniel
Woo Richard
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