Axial-flow compressor, more particularly one for an aircraft gas-turbine engine

Abstract

On an axial-flow compressor, more particularly one for an aircraft gas-turbine engine, including at least one rotor disposed in a casing and having compressor blades extending from a rotor hub as well as one stator, a slot-type recess (14) is formed in the compressor blades (7), with the recess originating at the trailing edge (13) of the compressor blades (7) and immediately adjacing the rotor hub (6). This enables secondary flows occurring in the transition area between rotor hub and compressor blades and, consequently, rotor losses to be reduced, thereby providing for increased compressor efficiency and reduced fuel consumption.

Description

This invention relates to an axial-flow compressor, more particularly one for an aircraft gas-turbine engine, including at least one rotor disposed in a casing and having compressor blades extending from a rotor hub as well as one stator.

The major components of an axial-flow compressor for increasing the pressure of the air supplied are the rotor, which is rotatably driven in a casing, and the stator, which is arranged on the casing inner wall and includes individual stator vanes. The rotor has compressor blades spacedly disposed on the periphery of a drive shaft or hub. Multi-stage compressors employed on gas-turbine engines include two or more rotor disks which are joined to a drum and connected to a drive shaft and on whose outer circumference the compressor blades are either separably mounted and, therefore, easily replaceable or, in the case of a blisk-type rotor, integrally formed on. While the compressor blades of a blisk extend immediately from the outer circumference of the rotor disk, the compressor blades mounted axially or circumferentially to the rotor disk each have a blade root (platform, shroud) integrally connected with them and being held on the rotor disk by a locating provision formed on the blade root underside.

On gas-turbine engines, distinction is made between the fan, the low-pressure, the intermediate-pressure and the high-pressure compressor, with either being required to satisfy high requirements on aerodynamic performance. In the transition area between the compressor blades and the rotor hub, i.e. the compressor blade and the periphery of the rotor disk of integrally formed-on blades or, respectively, the compressor blade and the blade root of blades replaceably mounted on the rotor disk, secondary flow phenomena, for example edge separation, channel swirls or transverse channel flows, occur which, in particular on rotors with high hub load, such as the fan or the low-pressure compressor of a gas-turbine engine, give rise to flow separation at the blade surface and the rotor hub (blade root, rotor disk) and locally entail high pressure losses. This leads to a reduction of compressor efficiency and, consequently, an increase in fuel consumption.

A broad aspect of the present invention is to provide an axial-flow compressor with enhanced aerodynamic performance, more particularly one for aircraft-gas turbines.

It is a particular object of the present invention to provide solution to the above problematics by an axial-flow compressor designed in accordance with the features of patent claim 1. Advantageous developments of the present invention become apparent from the sub-claims.

On the basis of an axial-flow compressor including at least one rotor disposed in a casing and having compressor blades extending from a rotor hub as well as one stator, and with the compressor blades of said axial-flow compressor having a blade tip, a leading edge and a trailing edge and a chord length measured at the rotor hub as well as a height between rotor hub and blade tip measured at the center of the blade, the present invention, in its essence, provides a slot-type recess originating at the trailing edge of the compressor blades and immediately adjacing the rotor hub which enables the secondary flows usually occurring in the transition area between rotor hub and compressor blades to be significantly reduced and, consequently, the losses due to secondary flow phenomena decreased, thereby providing for increased compressor efficiency.

The rotor hub can be formed by the blade root of separately manufactured compressor blades mounted on the periphery of a rotor disk or by the rotor hub itself in the case of compressor blades integrally connected with the rotor disk.

In a further embodiment of the present invention, the maximum height of the slot-type recess is two percent of the blade height and can be constant or varying in longitudinal direction, for example gradually decrease from the trailing edge.

In a further embodiment of the present invention, the minimum length of the recess is 10 percent of the chord length and the maximum length is 50 percent of the chord length.

The present invention is more fully described in light of the accompanying drawing showing a preferred embodiment. In the drawing,

FIG. 1 is a longitudinal section of a low-pressure compressor for an aircraft gas-turbine engine, and

FIG. 2 is a schematic representation of a compressor blade attached to a rotor hub.

The axial-flow compressor shown in FIG. 1, here a low-pressure compressor of an aircraft gas-turbine engine, includes five rotors 2 joined to a rotor drum 1. The rotor drum 1 is disposed in a casing 3 and connected to a drive shaft 4. The rotors 2 each have a rotor disk 2 which, so to speak, is a rotor hub 6 connected to a drive shaft 4 and from whose periphery the compressor blades 7 extend.

On the low-pressure compressor according to FIG. 1, the compressor blades 7 are provided as replaceable, individual blades separably mounted on the rotor disk 5 which have a blade root 8 (inner shroud, inner platform) forming the outer periphery of the rotor hub 6 or the rotor disk 5. Arranged between the compressor blades 7 of the adjacent rotors 2 is a stator 9 including stator vanes 10 fixed to the casing 3. Also, on rotors 2 provided as blinks, the compressor blades 7 can be integrally formed on the rotor disks 5 (rotor hub 6), with the compressor blades 7 here extending immediately from the periphery of the rotor disk 5 or the rotor hub 6, respectively.

FIG. 2 schematically shows a compressor blade 7 extending immediately from a hub 6, i.e. the periphery of the rotor disk 5, and having a blade tip 11, a leading edge 12, a trailing edge 13 and a chord length C at the rotor hub 6 as well as a height H between rotor hub 6 and blade tip 11 at half chord length C. Originating at the trailing edge 13 of the compressor blade 7 and formed in the portion thereof immediately adjacent to the rotor hub 6 (rotor disk 5, blade root 8) is a slot-type recess 14 whose height S does not exceed two percent of the blade height H and whose length L is at least 10 percent of the chord length C and at most 50 percent of the chord length C. Height S of the recess 14 along the chord length C can be constant or variable. In the representation in FIG. 2, height S of the recess 14, which here has a length L of 10 percent of the chord length C, gradually decreases towards the center of the blade.

By virtue of the slot-type recess 14 formed in the compressor blade 7 in the portion adjacent to the rotor hub 6 (blade root or rotor disk, respectively), the formation of channel swirls, transverse flows and edge separations in the transition area between rotor hub 6—here the blade root 8 or the rotor disk 5—is significantly reduced and, thus, the secondary flow characteristics and the inflow of the subsequent stator 9 improved, so that—in particular with rotors with high hub load, such as the fan of a gas-turbine engine—the rotor losses are reduced, enabling compressor efficiency to be increased and, ultimately, fuel consumption reduced.

LIST OF REFERENCE NUMERALS

• 1 Rotor drum
• 2 Rotor
• 3 Casing
• 4 Drive shaft
• 5 Rotor disk (rotor hub)
• 6 Rotor hub
• 7 Compressor blade
• 8 Blade root (rotor hub)
• 9 Stator
• 10 Stator blade
• 11 Blade tip
• 12 Leading edge
• 13 Trailing edge
• 14 Slot-type recess
• C Chord length
• H Height of blade
• S Height of slot-type recess

Claims

1. Axial-flow compressor, more particularly one for an aircraft gas-turbine engine, including at least one rotor (2) disposed in a casing (3) and having compressor blades (7) extending from a rotor hub (6) as well as one stator (9), with the compressor blades (7) having a blade tip (11), a leading edge and a trailing edge (12, 13) and a chord length (C) measured at the rotor hub (6) as well as a height (H) between rotor hub (6) and blade tip (11) measured at the blade center, characterized in that a slot-type recess (14) for reducing the secondary flows in the transition area between rotor hub and compressor blades is formed in the compressor blades (7), with the recess (14) originating at the trailing edge (13) of the compressor blades (7) and immediately adjacing the rotor hub (6).

2. Axial-flow compressor in accordance with claim 1, characterized in that compressor blades (7) provided with a blade root (8) are replaceably mounted on the periphery of a rotor disk (5), with the blade root forming the rotor hub (6).

3. Axial-flow compressor in accordance with claim 1, characterized in that the compressor blades (7) are integrally connected with the rotor disk (5), with the latter immediately forming the rotor hub (6).

4. Axial-flow compressor in accordance with claim 1, characterized in that the maximum height (S) of the slot-type recess (14) is two percent of the blade height (H).

5. Axial-flow compressor in accordance with claim 4, characterized in that the height (S) of the recess (14) varies in longitudinal direction.

6. Axial-flow compressor in accordance with claim 5, characterized in that the height (S) of the recess (14) gradually decreases departing from the trailing edge (13).

7. Axial-flow compressor in accordance with claim 4, characterized in that the height (S) of the recess is constant.

8. Axial-flow compressor in accordance with claim 1, characterized in that the minimum length (L) of the recess (14) is 10 percent of the chord length and the maximum length (L) is 50 percent of the chord length (C).