Apparatus and method for cooling Axi-Centrifugal impeller

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 space 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.

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.degree. 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 T3, is
 limited to 1100.degree. 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.degree. 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.degree. F., that would
 allow the compressor discharge temperature T3 to be increased by
 150.degree. F. to 1250.degree. 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.degree. 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.

DETAILED DESCRIPTION
 Referring to FIG. 3, an example of the preferred embodiment of the present
 invention is illustrated and generally indicated by the reference numeral
 30. The improved impeller 30 has an impeller body 32 with blades 34 on the
 front, a rim 36 and a back face with a generally flat portion 38 and a
 curved portion 39 much like the prior art impeller 12 illustrated in FIG.
 2. A shield plate 40, preferably made of titanium, is attached to and
 offset from the back face flat portion 38 and has a plurality of integral
 radial vanes 42 that contact the back face thereby forming a plurality of
 cavities through which cooling air is circulated, as will be described
 below, to cool the back face of the impeller body 32 thereby reducing the
 temperature near the rim 36 by approximately 150.degree. F., and also
 reducing the impeller thermal stress gradients from the front face to the
 back face. Shield plate 40 preferably has a thermal barrier coating, such
 as ceramic, on its aft face to reduce heat transfer across shield plate
 40.
 Referring also to FIG. 4, shield plate 40 generally has an annular disk
 shape with an outer rim and an inner rim. Near the outer rim of shield
 plate 40 is a circumferential rail 44 extending forward which has an
 integral radial step 46 extending outward around circumferential rail 44.
 Near the rim 36 of the impeller body 32 is mating ring 48 which extends
 aftward from back face flat portion 38 and is preferably disposed radially
 outward from and adjacent to circumferential rail 44. Mating ring 48 has a
 recess 50 which receives radial step 46 of circumferential rail 44.
 Because of the high rim speeds of these components, the circumferential
 rail 44, radial step 46 and mating ring 48 are continuous full hoop
 structures with uniform axial cross sections. The high rim speeds at this
 radial location result in centrifugal forces that would cause any lug or
 other discontinuous portion of these components to exceed the tensile
 strength of the material locally.
 To install shield plate 40 on impeller body 32, the impeller body 32 is
 heated and shield plate 40 is cooled sufficiently to allow the mating ring
 48 to pass over radial step 46. As the impeller body 32 cools and shield
 plate 40 warms, radial step 46 engages recess 50 making a mechanical
 joint. When the components reach room temperature, they are locked
 together. This mechanical joint axially retains the outer end of the
 shield plate against the impeller and seals against cooling air leakage
 through the joint. During operation, the centrifugal loads tend to make
 this joint close tighter. The relative radial positions of the
 circumferential rail and the mating ring could be reversed, but that
 arrangement may tend to allow the joint to open under centrifugal loads.
 Shield plate 40 has circumferential ring 52 at the outer rim which is
 located radially outward from circumferential rail 44 and mating ring 48,
 and extends generally radially outward from shield plate 40.
 Circumferential ring 52 provides a heat shield for the local area around
 mating ring 48 and circumferential rail 44.
 Referring again to FIG. 3, the impeller body 32 is connected to the turbine
 shaft 54 by a curvic coupling joint 56. The turbine shaft 54 has a
 piloting ring 58 just aft of the curvic coupling 56. A plenum ring 60,
 preferably made of titanium, is mounted on the turbine shaft 54 with its
 aft end 62 mated to the piloting ring 58. Piloting ring 58 preferably
 restrains end 62 both radially and axially. The plenum ring 60 extends
 radially outward and forward to a radial piloting ring 64 on the inner rim
 of shield plate 40, thereby creating a cavity 66 between the plenum ring
 60 and the impeller body 32. The plenum ring 60 preferably follows the
 contour of the back face curved portion 39, but it need not do so.
 Alternatively, it may span straight across between axial piloting ring 58
 and radial piloting ring 64, as indicated by the phantom lines 60A, but
 that makes cavity 66 significantly larger. Shield plate 40 and plenum ring
 60 are preferably made as two separate pieces as shown for cost reasons,
 but they may be made as a single structure, thereby eliminating the joint
 between the shield plate 40 and plenum ring 60.
 A plurality of radial vanes 61 extend forward from plenum ring 60 similar
 to vanes 42 on shield ring 40. Vanes 61 are needed to pump the cooling air
 inward through cavity 66. The cooling air from the radial vanes 42 of
 shield plate 40 is passed through cavity 66 and through the teeth of the
 curvic joint 56 where it is directed aftward to cool the second stage
 high-pressure turbine blades.
 Referring also to FIG. 3A, where the plenum ring 60 meets the shield plate
 40, the plenum ring 60 has a mating portion 68 extending radially outward
 adjacent the aft face of shield plate 40 near its inner rim. A piloting
 surface 70 on the inner forward portion of mating portion 68 engages and
 mates to the radial piloting ring 64 on shield plate 40. The shield plate
 40 has a plurality of circumferentially spaced apertures 72 at a uniform
 radius which align with apertures 74 in plenum ring 60. Clocking between
 shield plate 20 and plenum ring 60 is maintained by any known means, such
 as dowel pins, or alignment features integral with the components, to
 maintain alignment of apertures 72 and 74. Apertures 72 and 74 allow for
 passage of cooling air from tangential on-board injection (TOBI) nozzles
 76 forward through plenum ring 60 and shield plate 40 into spaces between
 radial vanes 42. At the outer rim of the plenum ring 60 a shield ring 78
 extends aftward from the outer edge of mating portion 68. The shield ring
 78 shields the apertures 74 and the exit port of the TOBI nozzles 76 from
 the compressor discharge flow stream which leaks past the rim 36 of
 impeller body 32.
 Pressure drop in the cooling air system is minimized by the use of the TOBI
 nozzles which are directed partially in the direction of impeller rotation
 so that air exiting the TOBI nozzles has a velocity component in the
 direction of impeller rotation. There are preferably a plurality of TOBI
 nozzles 76 circumferentially spaced around the engine. Each TOBI nozzle 76
 is positioned so that its exit port is located radially inward from and
 adjacent to shield ring 78, and is also aligned with apertures 74 and 72.
 A brush seal 86 (shown in FIG. 7) extends from a stationary structure to
 shield plate 40 above shield ring 78. The brush seal 86 limits the cooling
 air from flowing radially outward instead of through the apertures 74, and
 also shields the cooling air from inflowing leakage of the main airflow
 around rim 36.
 Referring to FIGS. 5 and 6, the relationship between radial vanes 42 and
 apertures 72 is shown. Radial vanes 42 are arranged in pairs straddling
 each aperture. Each pair of vanes 42 has a joining portion 80 connecting
 their inner ends to form a U-shape. The joining portion 80 is adjacent the
 aperture 72 and follows the shape of it to partially surround it. In the
 embodiment shown in FIG. 5, apertures 72 are round and joining portion 80
 is semicircular. Radial vanes 42 do not extend all the way to
 circumferential rail 44. Their terminal ends 82 leave a gap between the
 ends 82 and circumferential rail 44 which allows cooling air to flow
 around ends 82. When the shield plate 40 is installed on impeller body 32,
 the forward edge 84 of each radial vane 42 butts against the back face 38
 of impeller body 32 as shown in FIG. 3. Cooling air entering apertures 72
 from TOBI nozzles 76 flows radially outward from apertures 72 along the
 inside of each pair of radial vanes 42, then around ends 82 and radially
 inward between the pairs of vanes 42 as shown by arrows A. The resulting
 cooling is a forced-vortex, closed-circuit system, and therefore
 independent of the inflow leakage on the impeller back face from the main
 gas stream.
 Referring to FIG. 7, cooling of the impeller is accomplished by removing
 approximately 1.7 percent of the compressor design point air flow
 (designated 1.7 W.sub.a), routing it through a heat exchanger 90, then
 through the TOBI nozzles 76 where it circulates between the back face of
 the impeller body 32 and the shield plate 40 as illustrated in FIG. 5,
 then passes between the plenum ring 60 and impeller body 32, through the
 teeth of the curvic joint, and then aft as in the prior art FIG. 1 to the
 high-pressure turbines 20 where it is used to cool the second stage
 high-pressure turbine blades 21 before being recombined with the main flow
 through the turbine blades. Approximately 0.2% W.sub.a leaks past the
 brush seal 86 above the TOBI nozzles, and another 0.05% W.sub.a will leak
 outward through the joint between the impeller body 32 and the shield
 plate 40 at the impeller rim 36. This cools this local area and leaves
 approximately 1.5 W.sub.a available to cool the back face of the impeller,
 and then the turbines.
 The 1.7 W.sub.a cooling air is removed from the main gas stream at the
 diffuser pipes 16 through semi total pressure scoops to provide the
 maximum available pressure for the impeller cooling system. Pipes from the
 scoops are manifolded together, then the air is routed through the heat
 exchanger. This high-pressure air can accommodate the pressure drop
 associated with the cooling of the air in the heat exchanger. There is
 still sufficient pressure to circulate the cooling air along the impeller
 back face and through the curvic joint at the rear the impeller and then
 aft to cool the second stage high-pressure turbine blades 21. The cooling
 system for the impeller is thereby integrated with that for the
 high-pressure turbines, rather than being a separate add-on system. The
 flow volume, temperature and pressure of the cooling air exiting from the
 impeller back face circuit at the curvic joint 56 are sized to be suitable
 for cooling the second stage high pressure turbine blades, similar to the
 conditions at the cooling circuit entrance in front of the impeller for
 the prior art configuration shown in FIG. 1.
 This cooling air is piped externally to heat exchanger 90, which may be a
 conventional cross-flow type wherein the cooling air flows through pipes
 which are cooled by fuel from the main aircraft fuel tank flowing across
 them. However, an endothermic type heat exchanger is preferred because it
 would significantly reduce the heat exchanger weight and bulk. Fuel flow,
 indicated by the arrows F, is from the aircraft fuel tank, through the
 heat exchanger then to nozzles at the combustor 18. Cooling air flow,
 indicated by the arrows A, is from the scoops at diffuser pipes 16,
 through the heat exchanger 90, and then to TOBI nozzles 76.
 When T.sub.3 is more than 200.degree. F. below maximum, such as at
 low-power conditions, valve V1 is closed and valve V2 is open so bleed air
 bypasses the heat exchanger and its associated pressure drop. This reduces
 the active air system pressure differential at low-power conditions. When
 T.sub.3 is within 200.degree. F. of maximum, such as at higher power
 conditions, valve V1 is open and valve V2 is closed so that bleed air
 flows through the heat exchanger. The functions of V1 and V2 may be
 performed by a single device or by two or more devices using devices and
 switching methods known in the art. Valve V1 and a valve V3, which is
 located in the heat exchanger air exit pipe, are closed if a temperature
 indicator mounted inside the heat exchanger 90 indicates a temperature
 above normal operating limits. With all air to the heat exchanger shut
 off, there can be no combustion inside the heat exchanger. The functions
 of valve V1 and V3 may also be performed by a single device.
 The above described improved impeller for axi-centrifugal gas turbine
 engines and method of cooling it provides a closed circuit system that
 integrates with existing systems for cooling the second stage
 high-pressure turbine blades and has no impact on the engine primary flow
 streams. All of the impeller cooling system flows are forced-vortex to
 minimize flow variability and system pressure drops. The system provides
 for a temperature reduction of approximately 150.degree. F. at the rim of
 the impeller, which reduces the impeller thermal stress gradients from the
 front face to the back face and allows the compressor discharge
 temperature T3 to be increased approximately 150.degree. F., thereby
 increasing the overall engine efficiency.
 The descriptions above and the accompanying drawings should be interpreted
 in the illustrative and not the limited sense. While the invention has
 been disclosed in connection with the preferred embodiment or embodiments
 thereof, it should be understood that there may be other embodiments which
 fall within the scope of the invention as defined by the following claims.