Compressor

A compressor section for a gas turbine engine, the compressor section including a circumferential row of stator vanes, a circumferential row of rotor blades, and a casing arrangement. A row of bleed ports are provided in the casing arrangement. Each of the bleed ports extends radially and is curved such that an outlet of the port is rearward of and is circumferentially offset in a direction of rotation of the rotor blades from an inlet of the port.

TECHNICAL FIELD

The present disclosure concerns a compressor section and/or a gas turbine engine.

BACKGROUND

A gas turbine engine typically comprises a series of rotatable components, both in the compressor and turbine of the engine, which are housed within a casing. The rotatable components each comprise an array of blades, each having an aerofoil cross section. The blades are attached to a central hub or drum. The blades of the rotatable components accelerate the air through the engine and/or extract energy from the air. Each of the rotatable components are coupled with a static component which comprises an array of vanes that are also of aerofoil cross section. The static components are connected to the radially inner and/or outer casing components.

Typically air is bled from the compressor to provide cooling and sealing air to the turbine. Air can be bled from the compressor using one or more bleed ports. The bleed ports may be annular ports that extend circumferentially around the casing of the compressor. Alternatively, a series of discrete bleed ports may be provided circumferentially around the casing.

The use of bleed ports is important for cooling and sealing, but the provision of such holes can be a cause of aerodynamic losses.

SUMMARY

The present disclosure seeks to minimise the total pressure losses associated with the provision of bleed holes.

According to a first aspect there is provided a compressor section for a gas turbine engine. The compressor section comprises a circumferential row of stator vanes, a circumferential row of rotor blades, and a casing arrangement. The casing arrangement may be configured such that the rotors pass the casing arrangement during rotation in one axial region and the stator vanes are connected to the casing arrangement in another axial region. A row of bleed ports are provided in the casing arrangement, wherein each of the bleed ports extends radially and is curved such that an outlet of the port is rearward of and is circumferentially offset in a direction of rotation of the rotor blades from an inlet of the port.

In the present application, a forward and rearward direction is with respect to the axial flow of air through a gas turbine engine.

Each of the bleed ports may be configured to have a larger cross section at the outlet than at the inlet. For example, the bleed ports may be considered as divergent bleed ports. An inner surface of the port may be continuous between the inlet and the outlet, e.g. to diffuse the air flow through the port.

The outlet may be larger than the inlet in a circumferential and/or axial direction.

The bleed ports may have a circular or oval cross section.

The bleed ports may be provided axially adjacent to and forward of the row of rotor blades. Alternatively, the bleed ports may be provided axially adjacent to and rearward of the row of rotor blades.

The bleed ports may be provided axially rearward of the stator vanes.

The casing arrangement may circumscribe the row of stator vanes and the row of rotor blades.

The bleed ports may extend radially outwardly through the casing arrangement.

The casing arrangement may include a rotor casing proximal to the rotors and a stator casing to which the stator vanes are connected. The bleed ports may extend through both the stator casing and the rotor casing.

The bleed ports in the rotor casing may be co-axial to the bleed ports in the stator casing.

The rotor casing may include a rotor casing portion having a flange. The stator casing may include at least two stator casing portions each having a flange. The rotor casing portion flange may be positioned between and connected to the flanges of the two stator casing portions.

The bleed port may be provided adjacent to the flange of one of the stator casing portions.

An exit diffuser may be provided at the outlet of each of the bleed ports.

The casing arrangement may include a stator casing, and the exit diffuser may be connected to the stator casing.

The exit diffuser may be defined by a plurality of members that are wider at a position adjacent the outlet of the bleed port and narrower at the outlet of the diffuser. For example, the members that define the exit diffusers may be substantially triangular in radial cross section.

The ratio of the width of an outlet to the width of an inlet of the exit diffuser may be greater than the ratio of the width of the outlet to the width of the inlet of the bleed ports.

The compressor section may include a plurality of circumferential rows of stator vanes and a plurality of circumferential rows of rotor blades, each row of stator vanes being axially adjacent a row of rotor blades. Bleed holes may be provided at a single axial location, e.g. in a region of only one of the rows of stator vanes and rotor blades. Alternatively, bleed holes may be provided at multiple axial locations, e.g. in a region of two or more rotor blades and stator vanes.

According to a second aspect there is provided a compressor section for a gas turbine engine. The compressor section comprising a circumferential row of stator vanes, a circumferential row of rotor blades, and a rotor casing circumscribing the rotor blades and a stator casing circumscribing the stator vanes and to which the stator vanes are connected. The rotor casing includes a rotor casing portion having a flange, and the stator casing includes at least two stator casing portions each having a flange, and wherein the rotor casing portion flange is positioned between and connected to the flanges of the two stator casing portions.

The compressor section may comprise a plurality of bleed ports, and the bleed ports may be provided adjacent to the flange of one of the stator casing portions.

The bleed ports may be provided upstream of the flanges of the rotor and stator casing portions. The bleed ports may be adjacent to and upstream of the rotor blades. The flanges may be substantially axially aligned with the rotor blades.

The compressor section of the second aspect may have one or more features of the compressor section of the first aspect.

According to a third aspect there is provided a gas turbine engine comprising the compressor section of the first aspect and/or second aspect.

The high pressure compressor section may be the compressor section of the first and/or second aspect.

According to a fourth aspect there is provided a method of manufacturing a compressor section for a gas turbine engine. The method comprises providing a casing arrangement and connecting a circumferential row of stator vanes to the casing arrangement. A circumferential row of rotor blades is arranged such that they pass the casing arrangement during rotation. The method further comprises providing a row of bleed ports in the casing arrangement, wherein each of the bleed ports extends radially and is curved such that an outlet of the port is rearward of and is circumferentially offset in a direction of rotation of the rotor blades from an inlet of the port.

The bleed ports may be defined using machining.

The compressor section may be the compressor section of the first and/or second aspect.

DETAILED DESCRIPTION

With reference toFIG. 1, a gas turbine engine is generally indicated at10, having a principal and rotational axis11. The engine10comprises, in axial flow series, an air intake12, a propulsive fan13, an intermediate pressure compressor14, a high-pressure compressor15, combustion equipment16, a high-pressure turbine17, an intermediate pressure turbine18, a low-pressure turbine19and an exhaust nozzle20. A nacelle21generally surrounds the engine10and defines both the intake12and the exhaust nozzle20.

The compressed air exhausted from the high-pressure compressor15is directed into the combustion equipment16where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines17,18,19before being exhausted through the nozzle20to provide additional propulsive thrust. The high17, intermediate18and low19pressure turbines drive respectively the high pressure compressor15, intermediate pressure compressor14and fan13, each by suitable interconnecting shaft.

Referring now toFIGS. 2 and 3, the high pressure compressor includes a plurality of rows of stator vanes24and a plurality of rows of rotor blades26. The rows of blades and vanes are arranged circumferentially and each row of rotor blades has a row of stator vanes associated with it. Each of the stator vanes and rotor blades has a portion that has an aerofoil profile. The rotor blades are connected to a hub (not shown inFIG. 2 or 3), e.g. mounted or welded thereto.

A casing arrangement28circumscribes the stator vanes24and the rotor blades26. The stator vanes24are connected to the casing arrangement. A plurality of bleed holes30are provided in the casing arrangement and extend radially through the casing arrangement. In the present example, the bleed holes are provided just upstream of the seventh row of rotor blades. However, in alternative embodiments the bleed holes may be provided upstream of any one or more rotor stages of the compressor.

In the present example, the casing arrangement includes a stator casing (including stator casing portions32and34) and a rotor casing36. Each of the rotor casing and stator casing are made from a plurality of axial sections that are connected together. The rotor casing36provides a surface past which the rotor blades pass during operation. The stator vanes24are connected to the stator casing.

The stator casing portions32,34are connected to the rotor casing36using a flange arrangement38. Each of the stator casing portions and the rotor casing includes a flange40,42,44that extends radially outwardly from the respective casing. In the present example, the flanges are in an axial position corresponding to one of the row of rotor blades26(e.g. are axially aligned with the rotor blade). In the region of the flange, the rotor casing is positioned radially inward of the stator casing. The flange42of the rotor casing is received between the flanges40,44of the stator casing portions and the flanges are fixed in position using a fastener, e.g. one or more bolts.

As shown inFIG. 3, the bleed ports30are positioned forward (e.g. upstream) of the flange arrangement38and are axially adjacent the flange40of one of the stator casing portions32. The bleed ports30may be provided axially adjacent to and forward of the row of rotor blades26. Alternatively, as shown inFIG. 5, the bleed ports30may be provided axially adjacent to and rearward of the row of rotor blades26.

Each of the bleed ports30are defined by a port46formed in the rotor casing and a port48formed in the stator casing. The ports formed in the rotor casing are aligned with (e.g. coaxial with) the ports formed in the stator casing, so as to form a continuous bleed port. The ports46and48may be considered to be through holes in the respective rotor casing and stator casing.

Each bleed port30is profiled so as to curve in a radial direction from an inlet50to an outlet52of the bleed port. To achieve this curve both or one of the ports46,48formed in the rotor casing and stator casing may be curved. As can be seen fromFIG. 2, the ports are curved such that an inlet of the port is axially forward of an outlet of the port (e.g. the ports are curved in an axial direction C of air flow through the compressor). As can be seen fromFIG. 3, the ports are also curved such that the outlet of the port is circumferentially offset from the inlet of the port in a direction of rotation B of the rotors (e.g. the ports are curved in a direction of rotation of the rotor blades). The cross sectional area of the bleed port30at any point may be circular or oval.

The bleed ports30are also arranged to diffuse the flow through the bleed port. The inlet50to the bleed ports is smaller than the outlet52to the bleed ports. The bleed ports are arranged so that an inner surface of the port is continuous between the inlet and the outlet.

Referring toFIG. 4, in an alternative example of a compressor section, the bleed ports include an exit diffuser54. The diffuser is defined by members56connected to an outer surface of the stator casing (portion32). The members56are arranged so that an outlet58of the diffuser is larger than the outlet52of the port. The members56converge substantially to a point so that the outlets of the diffusers are adjacent each other with a minimal width wall separating them. In this example, the members are substantially triangular in radial cross section.

During use, air flow through the compressor section is bled through the bleed holes30and is directed to the turbine for sealing, cooling and/or other suitable purposes. The arrows (some labelled A) inFIGS. 2 to 4indicate the flow of air through the bleed holes.

The use of the described bleed ports increases static pressure of the bleed from the compressor. The profile of the bleed ports remove swirl from the bleed flow and diffuses the flow to increase the static pressure, which can contribute to improved aerodynamic efficiency of a gas turbine engine.

The use of a “triple flange”, that is the connection of the rotor casing flange to the flanges of the stator casing portions means that the profile of the bleed ducts can be more easily produced (e.g. machined) in the rotor and stator casings.