GB2253443A

GB2253443A - Gas turbine nozzle guide vane arrangement

Info

Publication number: GB2253443A
Application number: GB9104609A
Authority: GB
Inventors: Michael Josef Zdybel; George Fredrick Dyer
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1991-03-05
Filing date: 1991-03-05
Publication date: 1992-09-09
Also published as: GB9104609D0

Abstract

Separation of the gas flow from an outer casing 10 of a nozzle guide vane assembly is prevented by gas injection. Preferably gas injection is carried out through a first hole 15A situated at the junction of the vane suction surface, the outer casing 10 and the throat line 13 between two adjacent vanes, a second hole 15B situated downstream of the first at the junction of the vane suction surface and the outer casing and a third hole 15C situated on the throat line 13. By preventing gas flow separation, a low pressure region downstream of the nozzle guide vanes is avoided thereby preventing turbine rotor blade vibration otherwise caused by passage of rotor blade tips through such a region. <IMAGE>

Description

NOZZLE GUIDE VANE ARRANGEMENT This invention relates to nozzle guide vane arrangements for use in gas turbines.

A typical gas turbine engine is shown in figure 1. This comprises an air intake 1, compressor stage 2, fuel burning region 3, turbine stage 4 and outlet jet pipe 5 in flow series.

In figure 2 the turbine stage 4 is shown in more detail in cross section. The turbine stage 4 includes a turbine rotor 5 comprising a plurality of turbine blades 6 attached to the rim of a turbine disc 7. The disc 7 is mounted on a shaft 8 for rotation about an axis 9.

The compressor stage 2 is also attached to the shaft 8, allowing the turbine 5 to drive the compressor stage 2.

An outer casing 10 defines the outer limit of the gas path through the turbine stage 4, while an inner casing 11 defines the inner limit of the gas path. A plurality of inlet guide vanes 12 extend between the outer and inner casings 10 and 11 across the gas path. The vanes 12 defined a plurality of gas flow channels, one between each pair of adjacent vanes 12. The inner casing 11 is supported on the guide vanes 12 which are in turn supported by the outer casing 10. It should be noted that although the casing 10 is referred to herein as the outer casing because it defines the outer boundary of the gas flow through the turbine stage 4 the engine may include further structures such as supports or bypass ducts outside of it.

The inlet guide vanes 12 are aerofoils in cross-section and direct the gas flow from the fuel burning region in the best direction to maximise the efficiency of the momentum transfer from the gas flow to the turbine 5.

A problem in gas turbines is preventing damage due to vibration. Vibration is a particular problem in the turbine blades where blade design can be dictated by the need to ensure that the blades have no natural mode of vibration at frequencies at which they will be driven to vibrate in operation.

One way in which the turbine blades are driven into vibration is by the effect of the nozzle guide vanes on the gas flow through the turbine. Each nozzle guide vane has a low pressure region in the gas flow downstream of the vanes and adjacent the outer casing 10 associated with it. As the turbine rotates the turbine blade tips move in and out of these low pressure regions. When the turbine blade tips move into and out of these low pressure regions the gas loads on them change abruptly and these changing loads drive each turbine blade into vibration at a frequency equal to the rotational frequency of the turbine multiplied by the number of nozzle guide vanes.

This invention was intended to eliminate or reduce this source of excitation of the turbine blades.

It has been realised that the low pressure region adjacent the outer casing is produced by a separation of the gas flow through the nozzle guide vanes from the outer casing caused by interaction between the boundary layer of gas adjacent the outer casing and the suction surface leg of the horseshoe and/or passage vortices generated between each pair of adjacent vanes. This interaction causes the gas flow to separate from the surface of the outer casing in the suction surface diffusion region downstream of the throat. The throat being the point at which the gas passage between two adjacent vanes has its minimum cross-sectional area.

In a first embodiment this invention provides a nozzle guide vane arrangement for use in a gas turbine comprising an outer casing, a plurality of vanes each having a suction surface and defining between them gas flow channels with throats between them and gas injection means arranged to inject gas into regions where a vane suction surface meets the outer casing adjacent the throat of a gas flow channel.

In a second embodiment this invention provides a method of preventing gas flow separation in a nozzle guide vane arrangement for use in a gas turbine comprising an outer casing and a plurality of vanes each having a suction surface and defining between them gas flow channels with throats by injecting gas into regions where a vane suction surface meets the outer casing adjacent the throat of a gas flow channel.

This gas injection prevents the gas flow through the channel separating from the outer casing surface in the diffusion region downstream of the throat and so reduces the magnitude of the low pressure regions produced.

As a result the excitation of the blades as described above is reduced, allowing more freedom in the design of the blades.

The invention operates as follows. Upstream of each vane leading edge the boundary layer of gas flowing along the outer casing surface stagnates and splits to flow around the suction and pressure surfaces of the vane. As the boundary layer splits it separates from the outer casing surface and generates vortices. The suction surface leg of the horseshoe vortex, also known as the corner vortex, forms from the flow separation around the vane suction surface while the pressure surface legs of the horseshoe and passage vortices form from the flow separation around the vane pressure surface.

As a result, between each adjacent pair of vanes the suction surface leg of the horseshoe vortex from one vane interacts with the pressure surface legs of the horseshoe and passage vortices of the adjacent vane.

Between the vanes, downstream of the flow separation a new laminar boundary layer of gas flow along the outer casing surface is set up, but in the suction surface diffusion region downstream of the throat the horseshoe and passage vortices cause this new boundary layer to separate from the outer casing surface. This second separation causes a significant pressure loss in the gas stream flowing between the vanes. By injecting gas into regions where a vane suction surface meets the outer casing adjacent the throat of the gas flow channel defined between the vanes the second separation can be prevented and the pressure loss reduced. As a result the magnitude of the low pressure regions in the gas stream downstream of the vanes is reduced.

A nozzle guide vane arrangement employing the invention will now be described by way of example only with reference to the accompanying diagrammatic figures in which; Figure 3 shows a view along the axis of a gas turbine of a set of nozzle guide vanes employing the invention; Figure 4 shows a cross section along the arc X-X of figure 3; Figure 5 shows a pair of adjacent vanes from the assembly of figure 3 in perspective view; and Figure 6 shows a part of figure 5 in more detail, identical parts having the same reference numerals throughout.

Referring to figures 3 to 6 a set of nozzle guide vanes 12 used in a gas turbine stage of an aircraft jet engine are shown. The vanes 12 are evenly spaced around the circumference of the annular gas flow path through the turbine. This gas flow path is defined between the outer and inner casings 10 and 11 respectively and the gas flow is in the direction of the arrows 16.

Each pair of adjacent vanes 12 has a gas flow passage between them having its narrowest cross section at a throat line 13.

Each vane 12 has a first fillet 14 where it contacts the outer casing 10 and a second fillet (not shown) where it contacts the inner casing 11.

In the region where each vane 12 contacts the outer casing 10 adjacent to its throat line three air injection holes 15A, 15B and 15C are formed on the suction surface side of the vane 12. Injection hole 15A is situated on the first fillet 14 where it crosses the throat line 13, injection hole 13B is situated on the first fillet 14 just downstream of the throat line 13 and the injection hole 15C is situated in the outer casing 10 on the throat line 13 and adjacent to the fillet 14.

All three holes 15A, 15B and 15C pass through the outer casing 10 and are supplied with compressed air from the compressor stage 2 by ducting (not shown) running outside of the outer casing 10 and are inclined relative to the surface of the casing 10. When the turbine is operating compressed air is injected into the gas flow past the vanes 12 from the holes 15A, 15B and 15C, the inclination of the holes 15A, 15B and 15C directs the air in the same direction as the gas flow between the vanes 12.

By injecting air through the holes 15A, 15B and 15C separation of the gas flow from the surface of the outer casing 10 in the diffusion region downstream of the throat line 13 can be prevented. As a result the magnitude of the low pressure areas generated downstream of vanes 12 is reduced.

The injection of air at or near to the point where the throat line meets the vane suction surface is necessary to prevent this gas flow separation, however the optimum size, positioning, number and air flow rate of these holes will vary from one gas turbine design to another.

Although three injection holes are shown other numbers of holes arranged at or near to this point may be best in some gas turbine designs.

Although only three circular air injection holes are shown additional air injection holes or differently shaped air injection holes could be provided.

Although the example shown is in an aircraft jet engine the invention may be applied to other types of gas turbine.

Claims

1A nozzle guide vane arrangement for use in a gas turbine comprising an outer casing, a plurality of vanes each having a suction surface and defining between them gas flow channels with throats between them and gas injection means arranged to inject gas into regions where a vane suction surface meets the outer casing adjacent the throat of a gas flow channel.

2 An arrangement as claimed in claim 1 in which the gas injection means comprise a first gas injection hole at the point where the vane suction surface meets the outer casing on the throat of the gas flow channel, a second gas injection hole situated downstream of the first at a point where the vane suction surface meets the outer casing and a third gas injection hole situated on the throat.

3 An arrangement as claimed in claim 2 in which there is a fillet where the vane suction surface meets the outer casing and the first and second gas injection holes are situated on the fillet.

4 An arrangement as claimed in any preceding claim in which the gas injection means are arranged to inject air.

5 A method of preventing gas flow separation in a nozzle guide vane arrangement for use in a gas turbine comprising an outer casing and a plurality of vanes each having a suction surface and defining between them gas flow channels with throats by injecting gas into regions where a vane suction surface meets the outer casing adjacent the throat of a gas flow channel.

6 A nozzle guide vane arrangement substantially as shown in or described with reference to figures 3 to 6 of the accompanying drawings.