Patent Description:
Gas turbine engines provide power by compressing air using a compressor, adding fuel to this compressed air, combusting this mixture such that it expands through the blades of a turbine and exhausting the produced gases. The turbine consists of a disc, rotating about the central shaft of the engine, and a plurality of blades extending radially out of the disc towards the engine casing of the engine. Expansion of the combustion gases through the turbine causes its blades to rotate at high speed and the turbine, in turn, drives the compressor.

The distance between the tips of the blades and the inner surface of the compressor casing is known as the tip clearance. It is desirable for the tips of the blades to rotate as close to the casing without rubbing as possible because as the tip clearance increases, a portion of the compressed gas flow will pass through the tip clearance decreasing the efficiency of the compressor. This is known as over-tip leakage. The efficiency of the compressor, which partially depends upon tip clearance, directly affects the specific fuel consumption (SFC) of the engine. Accordingly, as tip clearance increases, SFC also rises.

As the disc and the blades rotate, centrifugal and thermal loads cause the disc and blades to extend in the radial direction. The casing also expands as it is heated but there is typically a mismatch in radial expansion between the disc/blades and the casing. Specifically, the blades will normally expand radially more quickly than the housing, reducing the tip clearance and potentially leading to "rubbing" as the tips of blade come into contact with the interior of the casing. Over time in use, the casing heats up and expands away from the blade tip, increasing the tip clearance. This may result in a tip clearance at stabilized cruise conditions that is larger than desired resulting in poor efficiency.

Conventionally, tip clearances are set when the engine is cold to allow for radial extension of the disc and blades due to centrifugal and thermal loads, to prevent rubbing. This means that there is initially a large tip clearance, such that the engine is relatively inefficient. When the engine is running, the blades will eventually extend radially to close this clearance, making the engine run more efficiently. Over a longer period of time, however, the temperature of the casing will rise and the casing will expand radially, which will again increase the tip clearance.

The running tip clearance of the high-pressure compressor (HPC) of an aircraft engine has a significant bearing on the efficiency of the HPC module. This, in turn, impacts other module attributes such as turbine durability as well as the engine fuel burn metric. Consequently much effort has been expended in ensuring that the running tip clearance is at the smallest mechanically feasible value.

<CIT> discloses features of the preamble of claim <NUM>.

In accordance with the present invention, there is provided a system as claimed in claim <NUM>, a gas turbine engine compressor as claimed in claim <NUM> and a process as claimed in claim <NUM>. Various embodiments of the invention are set out in the dependent claims.

Other details of the compressor case/tip clearance control system are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements.

<FIG> is a simplified cross-sectional view of a gas turbine engine <NUM> in accordance with embodiments of the present invention. Turbine engine <NUM> includes fan <NUM> positioned in bypass duct <NUM>. Turbine engine <NUM> also includes compressor section <NUM>, combustor (or combustors) <NUM>, and turbine section <NUM> arranged in a flow series with upstream inlet <NUM> and downstream exhaust <NUM>. During the operation of turbine engine <NUM>, incoming airflow F<NUM> enters inlet <NUM> and divides into core flow FC and bypass flow FB, downstream of fan <NUM>. Core flow FC continues along the core flowpath through compressor section <NUM>, combustor <NUM>, and turbine section <NUM>, and bypass flow FB proceeds along the bypass flowpath through bypass duct <NUM>.

Compressor <NUM> includes stages of compressor vanes <NUM> and blades <NUM> arranged in low pressure compressor (LPC) section <NUM> and high pressure compressor (HPC) section <NUM>. Turbine section <NUM> includes stages of turbine vanes <NUM> and turbine blades <NUM> arranged in high pressure turbine (HPT) section <NUM> and low pressure turbine (LPT) section <NUM>. HPT section <NUM> is coupled to HPC section <NUM> via HPT shaft <NUM>, forming the high pressure spool. LPT section <NUM> is coupled to LPC section <NUM> and fan <NUM> via LPT shaft <NUM>, forming the low pressure spool. HPT shaft <NUM> and LPT shaft <NUM> are typically coaxially mounted, with the high and low pressure spools independently rotating about turbine axis (centerline) CL.

Combustion gas exits combustor <NUM> and enters HPT section <NUM> of turbine <NUM>, encountering turbine vanes <NUM> and turbines blades <NUM>. Turbine vanes <NUM> turn and accelerate the flow of combustion gas, and turbine blades <NUM> generate lift for conversion to rotational energy via HPT shaft <NUM>, driving HPC section <NUM> of compressor <NUM>. Partially expanded combustion gas flows from HPT section <NUM> to LPT section <NUM>, driving LPC section <NUM> and fan <NUM> via LPT shaft <NUM>. Exhaust flow exits LPT section <NUM> and turbine engine <NUM> via exhaust nozzle <NUM>. In this manner, the thermodynamic efficiency of turbine engine <NUM> is tied to the overall pressure ratio (OPR), as defined between the delivery pressure at inlet <NUM> and the compressed air pressure entering combustor <NUM> from compressor section <NUM>. As discussed above, a higher OPR offers increased efficiency and improved performance. It will be appreciated that various other types of turbine engines can be used in accordance with the embodiments of the present disclosure.

Referring to <FIG>, an exemplary portion of a gas turbine compressor <NUM> section is shown. The compressor <NUM> disclosed can achieve a technical effect through a thermal contraction (or inhibit thermal expansion) of the case <NUM> through the use of air in conjunction with a case architecture featuring surface features <NUM>. In the invention, the inner case <NUM> proximate the compressor <NUM> comprises at least one surface feature <NUM>.

The surface features <NUM> are configured to create enhanced heat transfer from the inner case <NUM> to cooling air <NUM> flowing over the surface <NUM> of the inner case <NUM>. Enhanced heat transfer can be considered heat transfer rates that are greater than the heat transfer rate in the absence of the surface feature <NUM>. The surface features <NUM> can produce vortices <NUM> in the cooling air that produce convective heat transfer. The surface features <NUM> can include trip strip, ridges, nubs, raised and dimpled contours, vortex generators, raised flanges, and the like, that are configured to interrupt laminar flow boundaries along the surface <NUM> and cause mixing of the cooling air <NUM>.

Various portions of the inner case <NUM> architecture can be employed for the use of the surface features <NUM>. In an exemplary embodiment, portions nearest the rotating blades <NUM> can be utilized. The portions of the inner case <NUM> that are configured to maintain the tip clearance <NUM> between the blade tip <NUM> and inner case <NUM> can be configured with the surface features <NUM>. Either axial portions and/or radial portions of the inner case <NUM> can be employed as well. The surface features <NUM> can be employed in surfaces of the inner case <NUM> or components <NUM> that contact the cooling air <NUM>.

In an exemplary embodiment, the surface feature <NUM> can be formed as circumferential rings <NUM> that produce a symmetric response to the case <NUM>. The surface features <NUM> can be formed into a connector <NUM> that produces a symmetric response to the case <NUM>. The surface features <NUM> can be formed as circumferential ribs along an air seal support <NUM> that produces a symmetric response to the case <NUM>. In an exemplary embodiment the inner case <NUM> can be configured to experience the cooling effects of the surface features <NUM> to produce a predetermined blade tip to case clearance <NUM> change for from about <NUM>% to about <NUM>%. In an exemplary embodiment the predetermined blade tip to case clearance <NUM> change can be from about <NUM>% to <NUM>% blade cord.

The compressor <NUM> includes passageways <NUM> that are used to direct the air <NUM> through the compressor <NUM>, specifically into the location of the inner case <NUM> that includes the surface features <NUM>. Compressor <NUM> cooling air <NUM>, such as, upstream cooling bleed air, can be utilized to change the temperature of the inner case <NUM> to adjust the case <NUM> dimensions in order to maintain the proper predetermined tip clearance <NUM> between the case <NUM> and blade tip <NUM>.

A collection manifold <NUM> is fluidly coupled to the cooling air <NUM> to collect the air <NUM> and direct the cooling air <NUM> to a distribution manifold <NUM>. The distribution manifold <NUM> is configured to fluidly couple the cooling air <NUM> with the portion of the case <NUM> that includes the surface features <NUM>. The cooling air <NUM> flows over the surface feature <NUM> and exchanges thermal energy to cool the inner case <NUM>. Radial or axial portions of the inner case <NUM> that are required to control the tip clearance can receive the cooling air <NUM>.

A valve <NUM> is fluidly coupled between the collection manifold <NUM> and the distribution manifold <NUM>. The valve <NUM> is positioned to control the flow of cooling air <NUM>. The valve <NUM> can be adjusted to direct the cooling air <NUM> toward the distribution manifold <NUM> or to a bypass manifold <NUM>. The valve <NUM> can be used to control the temperature of the case <NUM> and control the tip clearance <NUM> dimensions between the blade tip <NUM> and case <NUM> responsive to the heat transfer enhancement of the surface features <NUM>. In an exemplary embodiment, the temperature differential employed to change the tip clearance <NUM> can be from about <NUM> degrees Kelvin (<NUM> degrees Fahrenheit) to about <NUM> degrees Kelvin (<NUM> degrees Fahrenheit).

A controller <NUM> is coupled with the valve <NUM> and configured to control the valve position. The controller <NUM> can be a microprocessor and the like, configured to receive input and send output to control the valve <NUM>. The valve <NUM> can be controlled to maintain/reduce the allocation of the cooling air <NUM> to the distribution manifold <NUM> and/or the bypass manifold <NUM>, since the total flow rate of bleed air <NUM> remains fixed and the valve position determines whether or not it is utilized to effect the thermal contraction of the case <NUM>. The valve <NUM> is not intended to control the mass flow rate of bleed air <NUM>. That mass flow rate is fixed by the requirement of turbine cooling, which is the intended final destination of bleed air <NUM>.

The controller <NUM> is utilized to control the cooling air <NUM> flow direction to change the tip clearance <NUM> by changing the temperature of the inner case <NUM>. The controller <NUM> can operate based on a predetermined schedule derived from engine operational data. For example, flight profile, predetermined schedules, and engine conditions can be utilized to modify the cooling air <NUM> flow and resultant temperature of the case <NUM> changing the thermal expansion and changing the dimensions of the case <NUM> relative to the tip <NUM>. In another embodiment, the controller <NUM> can be operated based on instrumentation and controls <NUM> coupled to the controller <NUM> and based on real time information (temperature, dimensions, operational mode) from the gas turbine engine <NUM>. The instrumentation and controls <NUM> include sensors (temperature, pressure, flow rate, altitude), programs, signals, communications links, engine operational data and the like. The instrumentation and controls <NUM> sense parameters and characteristics of the gas turbine engine <NUM> and provide data through signals. In an exemplary embodiment, the cooling air <NUM> can be directed over the surface features <NUM> during engine cruise conditions and redirected during engine transient conditions.

The distribution manifold <NUM> can be configured to be portions of the cooling flow pathway structures in the compressor <NUM>. The distribution manifold <NUM> can comprise the internal heat shielding baffles situated along the path the cooling air <NUM> travels. The distribution manifold <NUM> can be shaped by key placement of supports/baffles <NUM> and casing surfaces <NUM> to direct cooling air <NUM> flow. The distribution manifold can be modified in order to further control the flow of the cooling air <NUM> proximate the surface features <NUM>. The distribution manifold <NUM> can be configured, with a portion of a flow area <NUM> being convergent or divergent, as a nozzle, so as to increase flow velocity, or as a diffuser, so as to decrease flow velocity of the cooling air <NUM>. The distribution manifold <NUM> can be configured to provide a predetermined cooling air <NUM> velocity across the surface features <NUM> as well as maintain air pressure and mass flow rates of the cooling air <NUM>. A technical advantage of configuring the distribution manifold <NUM> includes the capacity to better control the cooling air <NUM> velocity and thus the cooling effects on the inner case <NUM> in coordination with the surface features <NUM>.

In an exemplary embodiment, an insert <NUM> can be coupled to the distribution manifold <NUM> and be configured as convergent or divergent, nozzle or diffuser. The insert <NUM> can change flow parameters proximate the surface features <NUM>. The insert <NUM> can be placed proximate the surface features <NUM>, upstream or downstream depending on the flow and cooling characteristics needed. The technical advantage the insert <NUM> can include the capacity to modify a particular inner case <NUM> design for a predetermined compressor <NUM> without the need to have to redesign the inner case <NUM>. The insert <NUM> can have certain flow characteristics that are effective for a particular inner case <NUM>. The insert <NUM> allows for flow characteristic modifications post construction of the original case <NUM>.

The combination of the surface features <NUM>, cooling air <NUM> with valve <NUM> and controller <NUM> and additional distribution manifold <NUM> creates an advantageous compressor case/blade tip clearance control system, or simply tip clearance control system <NUM> configured to enhance the cooling effect across the component <NUM> being cooled and thus control the tip clearance <NUM> between the inner case <NUM> and blade tip <NUM>. In an exemplary embodiment, the portion of the case <NUM> being cooled can be an air seal support <NUM>.

A technical advantage of the compressor case/blade tip clearance control system incorporated with the case is for better control the tip clearance between the case and the blade tips of the high pressure compressor.

A technical advantage of the compressor case/blade tip clearance control system incorporated within the case includes improving engine cycle performance and maintaining the bleed flow rate, thereby enhancing high pressure compressor life.

Another technical advantage of the compressor case/blade tip clearance control system incorporated within the case includes the capacity to control the flow of air supplied to the case and actively control the tip clearance responsive to gas turbine engine conditions.

Another advantage of the tip clearance control system is that it provides for architectures that enable control of the velocity of case cooling air over circumferential surface features that create vortices in the cooling air proximate the case in order to increase the convective heat transfer coefficient of the component being cooled.

Claim 1:
A compressor case to blade tip clearance system comprising:
a rotor having blades with tips (<NUM>);
a case including an inner case (<NUM>) comprising at least one surface feature (<NUM>);
a cooling air passageway;
a distribution manifold (<NUM>) disposed in the cooling air passageway, wherein the surface feature (<NUM>) is fluidly coupled to the distribution manifold (<NUM>), said at least one surface feature (<NUM>) configured to interact with cooling air (<NUM>) in the cooling air passageway;
a tip clearance (<NUM>) located between the tips (<NUM>) and the inner case (<NUM>), wherein said tip clearance (<NUM>) is maintained responsive to a flow of the cooling air (<NUM>) over said at least one surface feature (<NUM>);
a collection manifold (<NUM>) fluidly coupled to the distribution manifold (<NUM>);
a valve (<NUM>) fluidly coupled between said collection manifold (<NUM>) and said distribution manifold (<NUM>), said valve (<NUM>) configured to control said flow of cooling air (<NUM>) over said at least one surface feature (<NUM>); and
a controller (<NUM>) coupled to said valve (<NUM>), said controller (<NUM>) configured to actuate said valve (<NUM>) to control the air flow rate to change the tip clearance (<NUM>) by changing the temperature of the case (<NUM>),
characterised in that the system further comprises:
instrumentation and controls (<NUM>) coupled to the controller (<NUM>), said instrumentation and controls (<NUM>) configured to activate said controller (<NUM>) responsive to gas turbine engine real time operational condition information, wherein said instrumentation comprises sensors selected from the group consisting of temperature sensor, pressure sensor, flow rate sensor and altitude sensor.