Patent Application
20100098530
This application claims priority to German Patent Application DE 10 2008 052 101.9 filed Oct. 20, 2008, the entirety of which is incorporated by reference herein.
This invention relates to a compressor for a gas turbine, in particular an aircraft gas turbine, with a rotor hub carrying rotor blades, with a stator equipped with stator vanes, with a shroud associated to the stator vanes, and with an arrangement providing sealing between the shroud and rotor hub to prevent leakage.
The compressor, in particular an axial-flow compressor, includes a rotor shaft with one or several compressor stages having a rotor and a stator. The compressor may be provided with a so-called inlet guide vane assembly upstream of the first stage. The rotor equipped with rotor blades is connected to the compressor or the rotor shaft, respectively, and rotates in the casing of the compressor. The stator of each stage is equipped with stator vanes and is not connected to the compressor shaft. It is therefore stationary. Rotating and non-rotating blade rows alternate in the axial direction of the rotor shaft in a compressor.
Suitable attachment must be provided for the stators. For this, two options exist: 1) The stator has a hub gap, with the rotor shaft extending underneath the stator. In this arrangement, which is found, for example, on high-pressure compressors, the stator is connected to the compressor casing only. 2) On the radially inward side of the stator vanes the stator is provided with a so-called shroud, with the shroud being a ring connecting to the stator at the hub. This arrangement involves that, between rotor and stator hub upstream and downstream of the stator, an axial gap is formed which leads to flow leakage. This leakage must be minimized by a sealing arrangement in the form of seals relative to the rotor shaft. This arrangement is found on the sealing system for a gas turbine according to U.S. Pat. No. 6,932,349 B2. The smaller the gap between seal and shroud, the smaller the leakage flow between the forward and the rearward axial gap.
The arrangement with a hub shroud is mechanically complex, heavy and expensive. Also, the replacement of worn seals incurs considerable assembly effort and cost investment.
A strong transverse duct flow, which interacts with the leakage flow through the shroud and is disturbed, occurs on the stationary hub shroud from the pressure side to the suction side of the stator. The transverse duct flow tends to flow to the vane suction side. As a consequence, flow separation on the stator vane and, thus, severe losses occur. The leakage flow through the axial gap, which is situated before and behind the shroud, is driven by a strong pressure gradient between the leading edge and the trailing edge of the stator and, as it interacts with the main flow, can consequently get large and produce severe losses. Furthermore, heating of the flow is encountered within the shroud cavity. In order to reduce these adverse aerodynamic effects, leakage must be minimized. This can be accomplished by either increasing the number of seals or reducing the gap between seal and shroud. In either case, the geometry is complicated in terms of design and manufacture. As a consequence, increased costs are incurred. Since control of the sealing gap is only difficult to achieve, the operational risk for the compressor is increased as well. Summarizing, then, leakage should be completely prevented as it adversely affects the efficiency of the compressor and, consequently, the fuel consumption of the engine.
It is a particular object of the present invention to avoid these disadvantages of the state of the art.
The present invention, in a broad aspect, provides that the sealing arrangement between shroud and rotor hub is formed by a discharge arrangement for leakage air. Provision is thereby made for an almost complete suppression or elimination of leakage air and, concurrently, a simplification of design and manufacture.
Since special seals are dispensed with in this arrangement, no assembly effort is required for the replacement thereof in the case of wear. This provides for reduced costs and risks as well as improved efficiency and lower fuel consumption.
The present invention is implemented in that the state-of-the-art seals are dispensed with and the leakage flow within the shroud is discharged to the outside through an exhaust cavity within the shroud and the stator vane.
Discharge is accomplished in that the static pressure applied to the exhaust cavity is lower than the pressure prevailing in the shroud cavity. The static pressure must here be that much lower than the pressure prevailing in the shroud cavity, that the desired amount of reduced mass flow is removed.
This can be accomplished, for example, in that the exhaust cavity is connected via suitable lines to either an upstream compressor duct or to a secondary air system of the engine, with the removed air being used as turbine cooling air or for the pressurization of the aircraft cabin or other engine or aircraft systems, for example.
The amount of the removed, absolute leakage mass flow is controllable in that a defined throat is provided within the exhaust cavity or within the discharge arrangement.
Complete suppression of the leakage flow between the rearward and the forward axial gap requires that the shroud geometry, more precisely the flow-wetted area of the shroud, be appropriately attuned. The required area can be determined by way of the pressures before and behind the shroud and the exhaust mass flow.
Generally, the leakage mass flow through a shroud cavity with seals amounts to approx. 0.5-1 percent of the total flow. Consequently, the exhaust cavity can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged.
By way of the exhaust cavity, both the boundary layer before the stator and the boundary layer behind the stator are ingested into the shroud cavity and finally discharged to the outside. This leads to a significant reduction of the stator losses and, thus, to an increase in compressor efficiency. With use being also made of the discharged air, further potential for lowering fuel consumption is provided.
The exhaust cavity itself can have any cross-sectional shape and be provided, for example, as discrete round, elliptical or polygonal openings in any number and position on the shroud. Within the shroud, the cavity can either remain discrete or expand into a circumferential cavity. Vane passage can be provided by one or several tubes or by hollowing the vane, for example.
The shroud without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals. This provides for weight and cost savings. Furthermore, the leakage flow, which entails both losses and heating of the fluid and the components, is nonexistent. Exhaustion of the air leads to ingestion of the boundary layer of the compressor side wall, which further contributes to a reduction of losses. Utilization of the exhausted air for secondary systems provides for a further increase in efficiency. This means increased compressor efficiency and reduced fuel consumption. In summary, an increase in stage efficiency of approx. 0.2 percent is obtainable.
The present invention is more fully described below, in light of the accompanying figures showing an embodiment of a compressor for a gas turbine, here an aircraft gas turbine:
a to c show different air passage sections for the leakage air in a stator vane.
In this arrangement, a leakage flow occurs between the rotor hub 3 and the shroud 8 of the stator 5 in the forward and the rearward axial gap 12, 13, which is reduced by the two seals 16, 17 relative to the rotor shaft 2. The smaller the radial gap 18 between the seals 16, 17 and the shroud 8, the lower the flow leakage between the forward and the rearward axial gap 12, 13.
As illustrated in
Provided between the shroud 8 and the rotor hub 3 as sealing arrangement 20 is a discharge arrangement 23 for the leakage air which is removed from the shroud cavity 11 via an air discharge opening 24 and an air discharge duct 25 through the shroud 8 and the stator vane 7. The air discharge duct 25 here forms an exhaust cavity 30 for leakage air.
This arrangement provides for almost complete suppression of leakage air and, concurrently, simplification of design and manufacture. Since special seals 16, 17—as in the state of the art according to FIG. 1—are here dispensed with, no assembly effort for a replacement thereof is required in the case of wear. This provides for a reduction of costs and risks as well as improved efficiency and, thus, reduced fuel consumption of the aircraft gas turbine using the compressor according to the present invention.
The embodiment according to the present invention is implemented in that the state-of-the-art seals 16, 17 in accordance with
Discharge is accomplished as per
As shown in
The amount of the removed, absolute leakage mass flow is controllable in that a defined throat 29 (
Complete suppression of the leakage flow 22 between the rearward and the forward axial gap 13, 12 requires that the shroud geometry, more precisely the flow-wetted area 40 of the shroud 8, be appropriately attuned. The required area 40 (
In accordance with the state of the art, the leakage mass flow through the shroud cavity 11 with seals 16, 17 generally amounts to approx. 0.5-1 percent of the total flow. Consequently, the exhaust cavity 30 can be designed such that an amount of this size is exhausted, or, if used for example for pressurizing the aircraft cabin or for turbine cooling, it should be dimensioned such that the required mass flow, which normally amounts to approx. 2-4 percent, is discharged.
By way of the exhaust cavity 30, both the boundary layer 31 before the stator 7 and the boundary layer 32 behind the stator 7 are ingested into the shroud cavity 30 and finally discharged to the outside (
A variant of the air discharge is shown in
The exhaust cavity 30 itself can have any cross-sectional shape. See
The shroud 8 without seals according to the present invention is significantly easier to design and manufacture than an arrangement with seals 16, 17 in accordance with the state of the art as per
| Number | Date | Country | Kind |
|---|---|---|---|
| 10 2008 052 101.9 | Oct 2008 | DE | national |
