Gas turbine bleed air arrangement with an inlet having a non-uniform profile

Abstract

A gas turbine engine includes at least one bleed air passage extending radially and in fluid communication with at least one plenum extending generally circumferentially. The passage has an upstream inlet in fluid communication with the bleed air passage and an outlet for releasing air from the plenum. An upstream leading edge of the inlet or a downstream trailing edge of the inlet has a non-uniform profile.

Description

TECHNICAL FIELD

[0001] The present invention relates to a gas turbine engine of the type used in an aircraft. The invention is particularly, but not exclusively, applicable to commercial aircraft. BACKGROUND

[0002] A conventional gas turbine engine comprises an air intake, a series of compressors, a combustor, a turbine and an exhaust. The operation of a conventional gas turbine engine will be understood by those skilled in the art and is therefore not described in detail.

[0003] The invention described herein relates to an arrangement which is able to reduce / dampen a flow phenomenon which can cause an aero-acoustic-resonance of an engine bleed air system or other cavity, such as a plenum in one example.

[0004] Resonance can be dangerous for rotating turbomachinery as the resulting fluctuating pressure waves can cause engine / component failure. Resonance can also generate noise, but this is a secondary issue.

[0005] In respect of this secondary issue, aircraft noise is an important aspect of aircraft design and aircraft manufacturers are subject to increasing demands to reduce noise emissions in order to comply with increasing noise limits and environmental constraints at airports. If resonance can be reduced or prevented, then the associated noise caused by resonance can be prevented.

[0006] In order to prevent compressor stall at different engine operating conditions, a bleed air passage can be provided, typically within a duct between the low / intermediate pressure and high pressure compressors. The bleed air passage is arranged to release air from the main (core) flow path in order to divert air from the compressor and prevent compressor stall at certain operating conditions. This allows the gas turbine engine to continue operation over a wide range of operating conditions.

[0007] The air released by the bleed air passage can be communicated into one or more radially extending cavities, each cavity being referred to as a bleed plenum. Each plenum collects air before it is released into a bypass duct which surrounds the engine core. Each plenum provides a volume to receive air which has been diverted from the engine core.

[0008] However, when little or no air is released through the bleed air passage, the high speed air passing through the main flow path that interacts with the inlet or entrance to the bleed air passage can create unstable flow phenomena that can induce undesirable air-acoustic effects in the form of fluctuating pressure waves due to resonance in the bleed air passage and / or bleed air plenum. These pressure waves can damage the compressor and other structures. To prevent the occurrence of these damaging pressure waves, as described herein, the inventors have established that the geometry of the plenum can advantageously be designed such that air-acoustic resonance of the plenum does not occur within the engine operating range.

[0009] The inventors have devised a modified bleed air arrangement that addresses these problems in a convenient manner. SUMMARY

[0010] Aspects of the invention described herein are set out in the appended claims.

[0011] Viewed from a first aspect of the invention described herein, there is provided a gas turbine engine or engine component with a duct comprising at least one generally radially extending bleed air passage in fluid communication with an outlet for releasing air from the duct, wherein the bleed air passage has an inlet that is in fluid communication with the duct of the engine and that is defined between an upstream leading edge and a downstream trailing edge measured in the direction of air flow through the duct, and wherein the upstream leading edge of the inlet or the downstream trailing edge of the inlet has a non-uniform profile.

[0012] By modifying the manner in which air flows through the bleed air opening in the duct of the engine, the inventors have established that noise and potentially damaging vibrations can be avoided. The noise and vibrations can be caused by high speed air entering the bleed air passage and / or passing through the inlet to the bleed air passage. Contrary to intuition, the inventors have established that by replacing the smooth and uniform surface that typically forms the intersection of the engine duct and the bleed air passage (i.e. the perimeter or circumference of the inlet to the bleed air passage) with a non-smooth or non-uniform surface, acoustic resonance can be avoided.

[0013] In particular, by providing a non-uniform or non-smooth surface to the upstream edge (leading edge) and / or the downstream edge (trailing edge), the vortex that would otherwise be created can be mixed to avoid resonance from occurring.

[0014] The term "non-uniform" is intended to mean a surface that is not smooth and non-uniform in shape, such as a traditional straight or curved line that forms the conventional or existing inlet to the bleed air passage.

[0015] The non-uniform surface causes the vortex to break away or shed from the edge of the inlet at different locations with respect to the direction of air flow.

[0016] Problems occur in conventional bleed air arrangements in gas turbine engines due to the interaction between the breakdown of the shear layer in the bleed air passage inlet and acoustic standing waves (or acoustic response) in the bleed air system or cavity. Under certain conditions, the acoustic resonance frequency of the bleed air system coincides with the natural frequency of self-sustained shear layer instability in the shear layer at the bleed air passage inlet. When this occurs, the resulting unsteady pressure oscillations can have an adverse effect on surrounding components, such as the low pressure compressor (LPC) rotor stages and bleed air piping and valves. The vibrations within these components can damage the engine and its operation.

[0017] The non-uniform profile can be selected from a variety of profiles or shapes, each arranged to release or "trip" vortices from the edge. For example, a sawtooth (zigzag or triangular) shape can be formed on the edge of the bleed air passage inlet to create a surface that trips vortices at different locations, more particularly at different distances from the trailing edge of the bleed air inlet in the direction of air flow.

[0018] In other words, the non-uniform profile has some portions of its length (measured in the vertical direction relative to the air flow) more downstream (measured along the duct) than other portions. Advantageously, these portions can alternate, thereby providing an alternating profile that alternates between portions further downstream (in the flow direction along the duct) and some portions further upstream (in the flow direction along the duct) of the edge. Thus, a sinusoidal or zigzag or even a fractal or random edge can be achieved.

[0019] The purpose of the sawtooth is to prevent coherent accumulation or roll-up of Rossiter modes in the shear layer between the main duct flow and the slower fluid within the bleed air system or cavity. If the shear layer instability and coherent accumulation are disturbed, the forcing mechanism of the acoustic resonance is removed. The solution can consist of a sawtooth of the slot leading edge, or of a turbulence generator at the inlet leading edge, where the sawtooth disrupts the coherence of the Rossiter modes, and the turbulence generator affects the properties of the shear layer such that the Rossiter frequency is shifted from the cavity resonance frequency.

[0020] The trailing edge or leading edge can be modified to include a non-uniform profile. For example, in one arrangement, the trailing edge of the inlet can include a generally uniform profile and the leading edge of the inlet can include a non-uniform profile. In another arrangement, the leading edge of the inlet can include a generally uniform profile and the trailing edge of the inlet can include a non-uniform profile. Thereby, the vortex shedding can be optimized to reduce the risk of coherence in the shear layer instability. Additionally or alternatively, only a portion of the edge can have a non-uniform profile. For example, some portions of the trailing edge or leading edge can have a non-uniform profile as described herein while other portions can be smooth or uniform. Thus, the shedding performance can be optimized for each engine design and operating parameter.

[0021] It will be appreciated that the term non-uniform does not encompass microscopic or very small non-uniformities that would not affect the air flow. Rather, non-uniform refers to a modified surface profile that is sufficient to affect vortex shedding.

[0022] The shape or profile of the non-uniform surface can also be selected in dependence on the operating parameters of the engine. For example, the non-uniform profile of the leading or trailing edge can be in the form of a contoured surface having portions that extend a greater distance along the direction of air flow through the engine than other portions.

[0023] Thus, vortex shedding is induced at different locations along the direction of air flow. By releasing or shedding vortices at different locations, the vortices are induced to interact in a less coherent manner than from a generally, for example, straight or uniform surface or edge. Subsequent interaction as the vortices enter the air flow induces the vortices to lose coherence and can avoid or eliminate any natural frequencies.

[0024] In another arrangement, the non-uniform profile of the leading or trailing edge can be in the form of a meandering or sinusoidal edge or even a fractal or random profile.

[0025] In examples in which the bleed passage communicates with a plenum, the resonance is compounded by the plenum, which is in fluid communication with the bleed passage and receives bleed air from the duct via the bleed passage inlet. The plenum can for example be in the form of a circumferentially extending chamber around a central portion of the engine. The plenum can additionally be a single volume or can be subdivided into a plurality of individual plenums. By including one or more outwardly extending protrusions extending into the plenum, acoustic resonance in the plenum can be attenuated.

[0026] Such outwardly extending protrusions within the plenum can be in the form of convex surfaces extending into the corresponding plenum.

[0027] The non-uniform profile or shape of the leading or trailing edge of the bleed passage inlet can be any suitable shape as described above.

[0028] The size of the trailing edge serrations can be related to the size of the rossby mode vortices. The leading edge serrations or vortex generators can be very small to create turbulence, or comparable to the size of the rossby mode vortices (about 5% to 50% of the distance between the leading and trailing edges).

[0029] Indeed, as described herein, the non-uniform portions or parts of the leading and / or trailing edges are purposefully or intentionally arranged to induce vortices within the engine. This is counter-intuitive in itself.

[0030] Advantageously, the inventors have established that the non-uniformity or undulation of the leading edge (LE) and the trailing edge (TE) can have an amplitude of 2-20% of the height of the duct measured vertically across the passage immediately adjacent to the leading and trailing edges (this is represented by reference hc in Figure 6C below). Figure 3A

[0031] Furthermore, the inventors have established that the non-uniformity or undulation of the leading edge (LE) and the trailing edge (TE) can have an amplitude of 2-50% of the length of the opening to the bleed passage (in the flow direction along the duct). This is represented by reference L in Figure 6D below. Figure 3B

[0032] Still further, the inventors have established that the wavelength of the non-uniformity or undulation (i.e. the number of non-uniformities per unit length measured along the leading / trailing edge) can also advantageously be: (i) 2-20% of the height of the duct measured vertically across the passage immediately adjacent to the leading and trailing edges (this is represented by reference hc in Figure 6C below); or (ii) 2-50% of the length of the opening to the bleed passage (in the flow direction along the duct) (this is represented by reference L in Figure 6D below). Figure 3A Figure 3A Figure 3B

[0033] In another arrangement, a portion of the leading or trailing edge of the inlet can be moveable with respect to an adjacent portion or portions of the inlet. Thus, the profile of the non-uniform portion of the leading or trailing edge can be conveniently changed or modified by the relative movement. This can be achieved by a suitable actuator arrangement within the wall of the duct.

[0034] Additionally or alternatively, a portion of the trailing edge of the inlet can be provided with sound-absorbing material. Thus, flow disturbances impinging on the trailing edge can be reflected as sound waves which can be absorbed to improve the acoustic performance of the engine. In one example, a portion of the trailing edge of the inlet can be provided with a porous material, such as a foam, or an acoustic liner, such as a Helmholtz resonator. The foam can be a semi-permeable porous metallic material.

[0035] The skilled person will appreciate that the components described herein are located around the core of a gas turbine engine, the engine having a rotating main shaft. Thus, each bleed or cavity inlet can have a generally arcuate shape and be coaxial with the main shaft of the engine.

[0036] ​​​​​Viewed from a further aspect, there is provided an intermediate compressor structure for a gas turbine engine, the intermediate compressor structure comprising at least one bleed air passage allowing air to exit the compressor structure, the at least one bleed air passage having an inlet defined between an upstream leading edge and a downstream trailing edge measured in the direction of air flow through the engine, and wherein the upstream leading edge of the inlet or the downstream trailing edge of the inlet has a non-uniform profile.

[0037] Such a structure can be adapted to be positioned within a gas engine construction between a low pressure compressor and a high pressure compressor.

[0038] Viewed from a further aspect, there is provided a method of modifying a gas turbine engine, the gas turbine engine comprising at least one radially extending bleed air passage having an upstream inlet and an outlet for releasing air from the passage, wherein the inlet is defined between an upstream leading edge and a downstream trailing edge measured in the direction of air flow through the engine, and wherein the method comprises the step of altering the upstream leading edge of the inlet or the downstream trailing edge of the inlet so as to have a non-uniform profile.

[0039] Thus, a conventional gas turbine engine can be modified to incorporate the modified bleed air passage described herein, whilst retaining other engine components.

[0040] Viewed from a further aspect, there is provided a method of manufacturing an aero engine, the aero engine comprising at least one radially extending bleed air passage having an upstream inlet and an outlet for releasing air from the passage, wherein the inlet is defined between an upstream leading edge and a downstream trailing edge measured in the direction of air flow through the engine, and wherein the method comprises the step of altering the upstream leading edge of the inlet or the downstream trailing edge of the inlet so as to have a non-uniform profile.

[0041] According to such a method, the upstream leading edge of the inlet and / or the downstream trailing edge of the inlet can be altered to comprise a shape that undulates or is generally serrated in the direction of air flow through the duct of the engine. Furthermore, the trailing edge of the inlet can be provided with an acoustically absorbent surface as discussed above. BRIEF DESCRIPTION OF DRAWINGS

[0042] Aspects of the present application will now be described, by way of example only, with reference to the accompanying drawings in which:

[0043] Figure 1 A cross-sectional view of a gas turbine engine incorporating a bleed air system is shown;

[0044] Figure 2A A cross-sectional view through a bleed air system of an aero engine is shown;

[0045] Figure 2B and Figure 2CThe application is shown applied to plenum and non-plenum bleed arrangements;

[0046] Figure 3A and Figure 3B Cross-sectional views through plenum and non-plenum arrangements are shown respectively;

[0047] Figure 4 and Figure 5 A modified bleed front arrangement is shown;

[0048] Figure 6 A modified bleed trailing edge arrangement is shown; and

[0049] Figure 7A , Figure 7B and Figure 7C Examples of wavy profile trailing edges are shown.

[0050] Any reference in this specification to a document or a piece of art located on a server, website, or business database is not, and should not be taken as, an acknowledgement or any form of suggestion that the present invention is not entitled to antedate such prior art. Any reference in this specification to a document or a piece of art that is available on the internet or on an intranet or that is publicly available does not, and should not be taken as an acknowledgement or any form of suggestion that the present invention is not entitled to antedate such prior art. Any reference in this specification to a document or a piece of art that is publicly available, or that is posted on a server or a website does not, and should not be taken as an acknowledgement or any form of suggestion that the present invention is not entitled to antedate such prior art. The citation of any document or piece of art in this specification is not, and should not be taken as an acknowledgement or any form of suggestion that this document or piece of art is part of the common general knowledge of the public prior to the date of the application as claimed in this specification. The words "comprising", "containing", and similar words, when used in this specification are not to be interpreted as meaning "consisting of" or "consisting exclusively of". In other words, they are intended to mean "including but not limited to". The present invention is further described with reference to the following examples. It will be appreciated that the claimed invention is not intended to be limited in any way by these examples. It will also be appreciated that the present invention encompasses not only the individual embodiments described herein but also combinations of the embodiments described herein.

[0051] The various embodiments described herein are presented only by way of example. The embodiments are presented as representative examples only, and are not exhaustive and / or exclusive. It will be appreciated that the advantages, embodiments, examples, functions, features, structures, and / or other aspects described herein are not to be considered limitations on the scope of the claimed invention or on equivalents of the claims, and that other embodiments and modifications can be utilized and can be made without departing from the spirit and scope of the claimed invention. The various embodiments of the present invention can suitably comprise, consist of, or consist essentially of, the essential, or critical, elements described herein, in addition to other elements, components, features, parts, steps, means, etc. In addition, this disclosure can include other inventions now existing or which can be developed in the future that are not presently claimed but which can be claimed in the future.

[0052] While the application is susceptible to various modifications and alternative forms, specific embodiments are shown by way of example in the drawings and are herein described in detail. It should be understood however that the drawings and detailed description thereto are not intended to limit the application to the particular form disclosed but on the contrary, the intention is to cover all modifications, equivalents and alternatives falling within the spirit and scope of the claimed application.

[0053] It will be appreciated that features of aspects of the application described herein can be used, as appropriate, in any suitable combination. DETAILED DESCRIPTION

[0054] Figure 1 A cross-sectional view of a gas turbine engine 1 is shown, incorporating a bleed air system generally shown at 13.

[0055] The bleed air system can optionally include a plenum as described herein, or can alternatively be in the form of a passage to allow air communication without a plenum. The invention described herein is applicable to bleed air arrangements with and without a plenum.

[0056] The skilled person will be aware of the main components of a gas turbine engine and their operation. Broadly speaking, the engine 1 comprises an air intake 2 which allows air to flow into the engine to a fan 3 at the upstream end of the engine. All the components are housed within a nacelle 4.

[0057] The engine comprises a bypass passage downstream of the fan and a central engine core containing a compressor, a combustor and a turbine. The core of the engine is formed by a first low pressure compressor 5 and a second high pressure compressor 6. There can also be an intermediate compressor. This multi-stage compressor arrangement takes air from ambient pressure and temperature to high temperature and pressure. The compressed air is then communicated to a combustion chamber 7 where fuel is injected and combustion takes place.

[0058] The combustion gases exit from the rear of the combustion chamber 7 and first impinge on the high pressure turbine 9 before passing through the core nozzle 11 to exit the rear of the engine. Thrust from the engine is generated by two gas flows: first from the fan nozzle 8 (which receives thrust from the fan) and second from the exhaust of the core nozzle 11.

[0059] A transition duct 14 is arranged to receive air from the low or intermediate pressure compressor 5 and communicate these radially inwards to supply the high pressure compressor 6.

[0060] As shown, all the compressors are coaxial with the central axis of the turbine. The low / intermediate pressure compressor 5 has a larger outer radius (measured from the central axis of the compressor) than the outer radius of the high pressure compressor 6 for efficiency reasons (examples discussed above).

[0061] This requires that the duct or passage communicating air between the two compressors is generally S-shaped to communicate the compressed air towards the central axis of the turbine and into the high pressure turbine 6.

[0062] As described herein, it is desirable to be able to bleed or bleed off some air from the engine within the last few stages of the low / intermediate compressor or the transition duct. Thus, the bleed can also be positioned between the last LPC rotor and its OGV, although in such a design the OGV is typically at the inlet to the transition duct and can be considered to be in / part of the transition duct. This can be used to control the amount of air passed to the high pressure compressor and for example to prevent the low pressure compressor from stalling.

[0063] As Figure 1 shown in Figure 1 5, an outlet 15 is provided which provides an openable passage to allow air to be selectively passed from the transition duct 14 to an annular chamber, commonly referred to as a plenum.

[0064] The plenum can be arranged downstream of the low pressure compressor. In particular, the plenum can be arranged radially outward of the core and the bleed passage is typically downstream of the LPC.

[0065] The plenum is an annular chamber which extends around all or part of the engine and is arranged to receive air bled from the main flow path. In effect, the plenum acts as a collection chamber or reservoir for air bled from the main flow path.

[0066] Figure 2A An enlarged sectional view of the bleed system and its position relative to the core flow path or main flow path and the transition duct is illustrated. In a conventional engine, air enters the plenum through a bleed passage which is connected to the main flow path at position A. The air then exits the plenum through a port or valve B which is in close proximity to a bypass passage C of the engine. There are also other arrangements with a valve at the end of the bleed bleed pipe D before the air enters the plenum.

[0067] Figure 2B and Figure 2C illustrate arrangements with and without a plenum respectively. The plenum (P) in Figure 2B forms an annular chamber into which air can be communicated and then bled. Figure 2C illustrates an arrangement with an air passage but without a plenum.

[0068] Figure 3A and Figure 3B illustrate sectional views through the duct, bleed passage and plenum Figure 3A illustrates a plenum in communication with the bleed passage, Figure 3B illustrates a bleed passage in communication with the duct alone). It will be appreciated that the plenum can be in the form of an annulus which extends around the engine and is coaxial with an axis which extends along the core of the engine. The plenum can be a continuous chamber or can be a number of discrete chambers each performing the same function of collecting bled air. Similarly, the bleed passage inlet can be a generally continuous annular shape.

[0069] As Figure 3A and Figure 3B illustrated schematically in FIG. 1, air passing through the main flow path 14 can induce flow phenomena at the inlet of the passage 15, such as shear layer instability (Rossiter mode), or a convected flow phenomenon from the upstream compressor to the inlet of the passage 15, which subsequently interacts with the bleed cavity 13. Disadvantageously, a coupling effect can occur at point P, where there is a coupling between the cavity acoustics of the plenum and the Rossiter mode of air passing through the inlet to the passage.

[0070] The invention described herein relates to the air-acoustic effects of this arrangement and how the air flow in the main flow path is generating undesirable sound waves A w and vibration effects within the bleed system and engine, such as the low pressure compressor LPC.

[0071] The bleed cavity 13 defines a volume having a characteristic resonance frequency Fr, which depends on various factors of the bleed design and volume. Air in the main gas path passing through the inlet of the bleed passage 15 can induce excitation of the air within the plenum at, for example, a frequency F e .

[0072] In the case where the excitation frequency is close to or equal to the resonance frequency (i.e. when F r = F e , an acoustic resonance A r may occur within the plenum, generating highly undesirable air-acoustic effects and potentially damaging vibrations within the engine.

[0073] The frequency at which this resonance can occur will depend on the operating conditions of the engine and the acoustic characteristics of the plenum.

[0074] As discussed above, the interaction between the acoustic properties of the bleed system and the acoustic effects of the air flow over the inlet onto the bleed passage creates significant problems in engine noise, pressure loss and vibration.

[0075] In particular, with reference to the air flow over the inlet, the shear layer over the inlet to the passage is naturally unstable. This instability will grow with the development of the shear layer over the opening, causing oscillations of the shear layer. The oscillating shear layer will generate unsteady forces upon impinging the trailing edge of the opening.

[0076] This unsteady force in turn will generate acoustic waves propagating away from the trailing edge. Under the right conditions, the device acoustic waves will reinforce the initial shear layer instability and amplify the shear layer oscillations. The amplified oscillations in turn will generate even stronger forces and stronger acoustic response on the open trailing edge, and so on. Depending on the open geometry (e.g. length of the opening) and flow conditions (e.g. velocity and temperature) a feedback loop with a specific frequency will be established, known as the Rossiter mode.

[0077] The cavity created by the opening in the gas path has a specific acoustic resonance frequency. Under certain conditions, the natural frequency of the shear layer oscillations will be close enough to the acoustic resonance frequency of the cavity to create a feedback loop between the acoustic response in the cavity and the Rossiter mode. Then, the amplitude of the radiated acoustic waves can become very strong and lead to high transient loads on the engine components, such as compressor blades.

[0078] The invention described herein reduces or eliminates the strong acoustic forces by breaking or weakening the feedback loop in the Rossiter mode.

[0079] By modifying the geometry so that the length of the opening varies along the edge and thereby reduces the coherence of the shear layer oscillations (Rossiter mode), the feedback loop can be broken at the leading or trailing edge of the cavity.

[0080] Another way to weaken the feedback is to modify the trailing edge so that the acoustic pulse from the shear layer hitting the trailing edge is weakened. This can be done by treating the trailing edge with an acoustic liner (Helmholtz resonator) or a porous material (e.g. foam).

[0081] Yet another way to weaken the formation of the Rossiter mode is to modify the leading edge with small geometric features such as vortex generators or flow trip devices to change the initial conditions of the shear layer in the opening to bleed air. These in Figure 4 are illustrated by triangles 18.

[0082] Figure 4 A bleed air passage arrangement is illustrated. Here, the leading edge LE has a straight profile 16, i.e. a generally uniform shape or profile. Triangles 18 indicate trip devices at the leading edge. The illustrated vortices 17 are created as air breaks away from the edge of the inlet. The vortices travel TE. As shown, the vortices 17 are generally similar, all arrive at TE at the same time and travel across the width of the inlet at the same speed. This uniformity of the vortices allows the Rossiter mode to be enhanced, and then resonance can occur. As discussed above, a feedback loop can quickly be generated, producing undesirable pressure waves, vibrations and noise. This is further exacerbated by the interaction of the Rossiter mode with the plenum acoustic properties (as described above). The effect of the trip devices 18 is to disrupt the formation of coherent vortices involved in the feedback loop that leads to the Rossiter mode.

[0083] Figure 5 A modified air intake according to the invention described herein is shown. Here, the leading edge LE has been provided with a serrated or undulating surface 18, which has portions that extend further (axially or radially) than other portions. Specifically, the amplitude of the interval between the crests and troughs of the wave or sine wave is x. The amplitude x is the distance between the maximum and minimum values ​​of the wave.

[0084] like Figure 5 As shown, vortex 19A begins or leaves the leading edge at a point earlier than vortex 19B. Therefore, the two vortices 19a and 19b travel across the inlet width and arrive at the trailing edge TE at different times. Thus, adjacent vortices do not constructively interfere with each other, avoiding resonance.

[0085] Figure 6 An alternative arrangement is shown, in which the leading edge LE is uniform, i.e., unmodified, while the trailing edge TE is modified. In this example, the trailing edge has a serrated or undulating profile. Here, although vortices 20A, 20B, and 20C depart from the leading edge at the same point, they meet the trailing edge at different points, specifically at different distances along the airflow direction across the inlet. Due to the undulating profile, the vortices... Figure 6 Impacts measured from left to right at different locations Figure 6 On the undulating trailing edge of the eddy current. The different impacts of the eddy current on the modified trailing edge are incoherent, thus preventing resonance.

[0086] Figure 7A , Figure 7B and Figure 7C Example trailing edges of undulating profiles are shown. In each example, the trailing edge comprises a non-uniform trailing edge in the form of multiple adjacent concave indentations (in these examples).

[0087] As shown in the figure, the inlet of the bleed air channel can be like... Figure 7C The curve shown has a corresponding non-uniform trailing edge profile. The non-uniform concavity can also optionally taper along the inner surface of the air intake channel.

Claims

1. A gas turbine engine assembly comprising a duct, the duct comprising at least one bleed passage extending generally radially, the bleed passage being in fluid communication with an outlet for releasing air from the duct, wherein the bleed passage having an inlet, the inlet being in fluid communication with the duct of the engine, the inlet being defined between an upstream leading edge and a downstream trailing edge measured in the direction of air flow through the duct, and wherein the upstream leading edge of the inlet has a non-uniform profile extending into the inlet and the bleed passage, or the downstream trailing edge of the inlet has a non-uniform profile extending into the inlet and the bleed passage, wherein the non-uniform profile of the leading edge or the trailing edge is in the form of a contoured surface, portions of the contoured surface extending to a greater distance along the direction of air flow through the engine than other portions.

2. The engine assembly of claim 1, wherein, the trailing edge of the inlet comprises a generally uniform profile, and the leading edge of the inlet comprises the non-uniform profile.

3. The engine assembly of claim 2, wherein, the leading edge of the inlet comprises a vortex generating surface profile.

4. The engine assembly of claim 1, wherein, the leading edge of the inlet comprises a generally uniform profile, and the trailing edge of the inlet comprises a non-uniform profile.

5. The engine assembly of any one of claims 1-4, wherein, the non-uniform profile of the leading edge or the trailing edge is in the form of a meandering, sinusoidal, triangular or random edge.

6. The engine assembly of any one of claims 1-4, wherein, a portion of the leading edge of the inlet is movable relative to one or more adjacent portions of the inlet to vary the profile of the leading edge.

7. The engine assembly of any one of claims 1-4, wherein, a portion of the trailing edge of the inlet is movable relative to one or more adjacent portions of the inlet to vary the profile of the trailing edge.

8. The engine assembly of any one of claims 1-4, wherein, the amplitude of the non-uniformity of the leading edge and / or the trailing edge is 2-20% of the height of the duct measured perpendicularly across the passage immediately adjacent to the leading or trailing edge of the inlet.

9. The engine assembly of any one of claims 1-4, wherein, the amplitude of the non-uniformity of the leading edge and / or the trailing edge is 2-50% of the length of the inlet opening to the bleed passage in the direction of flow along the duct.

10. The engine assembly of claim 9, wherein, the wavelength of the non-uniformity is the number of non-uniformities per unit length measured along the leading edge / trailing edge, which is also: (i) 2-20% of the height of the duct measured perpendicularly across the passage immediately adjacent to the leading or trailing edge; and (ii) 2-50% of the length of the inlet opening to the bleed passage in the direction of flow along the duct.

11. The engine assembly of any one of claims 1-4, wherein, a portion of the trailing edge of the inlet is provided with sound absorbing material.

12. The engine assembly of claim 11, wherein, a portion of the trailing edge of the inlet is provided with porous material.

13. The engine assembly of claim 11, wherein, a portion of the trailing edge of the inlet is provided with foam.

14. The engine assembly of any one of claims 1-4, wherein, a portion of the trailing edge of the inlet is provided with acoustic lining.

15. The engine assembly of any one of claims 1-4, wherein, a portion of the trailing edge of the inlet is provided with a Helmholtz resonator.

16. An intermediate compressor structure for a gas turbine engine, the intermediate compressor structure comprising at least one bleed passage allowing air to exit the compressor structure, the at least one bleed passage having an inlet, the inlet being defined between an upstream leading edge and a downstream trailing edge measured in the direction of air flow through the engine, and wherein, the upstream leading edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, or the downstream trailing edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, wherein the non-uniform profile of the upstream leading edge of the inlet or the downstream trailing edge of the inlet is in the form of a contoured surface, portions of the contoured surface extending to greater distances along the direction of air flow through the engine than other portions.

17. A method of modifying a gas turbine engine, the gas turbine engine including at least one bleed air passage extending radially, the at least one bleed air passage having an upstream inlet and an outlet for releasing air from the passage, wherein the inlet being defined between an upstream leading edge and a downstream trailing edge measured along a direction of air flow through the engine, and wherein the method includes the step of: altering the upstream leading edge of the inlet so as to have a non-uniform profile extending into the inlet and the bleed air passage, or altering the downstream trailing edge of the inlet so as to have a non-uniform profile extending into the inlet and the bleed air passage, wherein the non-uniform profile of the upstream leading edge of the inlet or the downstream trailing edge of the inlet is in the form of a contoured surface, portions of the contoured surface extending to greater distances along the direction of air flow through the engine than other portions.

18. A method of manufacturing an aeroengine, the aeroengine including at least one bleed air passage extending radially, the at least one bleed air passage having an upstream inlet and an outlet for releasing air from the passage, wherein the inlet being defined between an upstream leading edge and a downstream trailing edge measured along a direction of air flow through the engine, and wherein the method includes the step of: altering the upstream leading edge of the inlet so as to have a non-uniform profile extending into the inlet and the bleed air passage, or altering the downstream trailing edge of the inlet so as to have a non-uniform profile extending into the inlet and the bleed air passage, wherein the non-uniform profile of the upstream leading edge of the inlet or the downstream trailing edge of the inlet is in the form of a contoured surface, portions of the contoured surface extending to greater distances along the direction of air flow through the engine than other portions.

19. The method of claim 18, wherein, the upstream leading edge of the inlet and / or the downstream trailing edge of the inlet is altered to include a contoured shape in the direction of air flow through the engine.

20. The method of claim 18 or 19, wherein, the trailing edge of the inlet is provided with an acoustically absorbent surface. the upstream leading edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, or the downstream trailing edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, wherein the non-uniform profile of the upstream leading edge of the inlet or the downstream trailing edge of the inlet is in the form of a contoured surface, portions of the contoured surface extending to greater distances along the direction of air flow through the engine than other portions. the upstream leading edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, or the downstream trailing edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, wherein the non-uniform profile of the upstream leading edge of the inlet or the downstream trailing edge of the inlet is in the form of a contoured surface, portions of the contoured surface extending to greater distances along the direction of air flow through the engine than other portions. the upstream leading edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, or the downstream trailing edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, wherein the non-uniform profile of the upstream leading edge of the inlet or the downstream trailing edge of the inlet is in the form of a contoured surface, portions of the contoured surface extending to greater distances along the direction of air flow through the engine than other portions. the upstream leading edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, or the downstream trailing edge of the inlet has a non-uniform profile extending into the inlet and the bleed air passage, wherein the non-uniform profile of the upstream leading edge of the inlet or the downstream trailing edge of the inlet is in the form of a contoured surface, portions of the contoured surface extending to greater distances along the direction of air flow through the engine than other portions.

Images