GB2110767A - A shrouded rotor for a gas turbine engine - Google Patents

Abstract

The rotor, e.g. for a turbine, has a shroud whose outer surface is formed to reduce the inefficiency caused by leakage flow of gas over the outside of the shroud. The forward region 27 of this surface is formed to cooperate with static structure 28 to define a narrow clearance which directs the leakage flow into a plenum chamber 32. The leakage flow remains attached to the intermediate section 31 of the surface which is flat or gently curved to encourage this effect. At its rearward portion the shroud has enclosed ducts 36 which direct the leakage flow in the same direction as the main gas flow leaving the blades, and which preferably do not involve a change in static pressure, so that recirculation of the leakage flow is discouraged and all the leakage flow passes through the ducts 36. Blade cooling air may also be discharged through the ducts 36. <IMAGE>

Description

SPECIFICATION A shrouded rotor for a gas turbine engine This invention relates to a shrouded rotor for a gas turbine engine, and is particularly relevant to the turbine rotors.

In such rotors a well-known source of inefficiency lies in the leakage of gas or air over the radially outer surface of the shroud. This leakage causes inefficiency in two main ways, thus the basic fact that some gas by-passes the blading of the stage produces a direct penalty while the fact that this flow re-enters the main gas flow at a non-conforming angle and velocity results in losses and inefficiencies in subsequent stator and rotor stages.

In the past, many ways have been tried to overcome or at least reduce this problem.

Thus the obvious way to reduce the leakage is to improve the seal between the shroud and the associated static structure, but it is not easy to improve the seal beyond a certain level because allowance must be made for relative expansions. An alternative is to make use of the leakage air in assiting the driving of the rotor, which may have the added benefit of redirecting the air to match more closely the main gas flow. This approach is exemplified in our patent 1423833, and provides a significant advantage over plain shrouds.

There is, however still scope for improvements, and the present invention provides an aerodynamic structure allowing effective capture and use of the leakage flow.

According to the present invention a shrouded rotor for a gas turbine has a shroud outer surface comprising a forward portion which cooperates with static structure to form a close annular clearance, an intermediate portion presenting a smooth surface to which leakage gas may attach by the Coanda effect, a plurality of inlets into which said leakage gas flows directly from said intermediate portion, and enclosed ducts leading from said inlets across the rearward portion of the shroud, the ducts being curved and shaped to direct the leakage gas in the same direction as the main gas flow leaving the rotor blades.

Preferably the ducts are shaped to allow the static pressure at their inlets to equal that at their outlets, so that leakage round the outer surface of the ducts is discouraged. This leakage may also be deterred by arranging a close annular clearance between the outer surface of the shroud portion carrying the ducts and the associated static structure.

It may also be desirable to eject some spent cooling air from the blade into the enclosed ducts.

The invention will now be particularly described, merely by way of example, with reference to the accompanying drawings in which: Figure 1 is a partly sectioned view of a gas turbine engine having a shrouded high pressure turbine rotor in accordance with the inventor; Figure 2 is an enlarged view of the shroud portion of the shrouded high pressure turbine rotor of Fig. 1 and Figure 3 is a developed section through part of the shroud of Fig. 2 on the line 3-3.

In Fig. 1 there is shown a gas turbine engine comprising a casing 10 within which are mounted in flow series low pressure and high pressure compressors 11 and 1 2 respectively a combustion system 13, and high pressure and low pressure turbines 14 and 1 5 respectively.Operation of the engine overall is conventional in that air is drawn into the intake 1 6 formed by the leading edge of the casing 10, and is compressed in the low and high pressure compressors 11 and 1 2. The compressed air is mixed with fuel in the combustion system 1 3 and burnt, and the hot gases thus produced flow through and energise the turbine 14 and 1 5. Each turbine is connected to its respective compressor by a drive shaft; thus high pressure turbine 1 4 drives high pressure compressor 1 2 through a high pressure shaft 1 7 while low pressure turbine drives low pressure compressor through a low pressure shaft 1 8.

The gas exhausting from the low pressure turbine 1 5 exhausts from the engine in the final nozzle 1 9 formed by the downstream extremity of the casing 10, producing propulsive thrust. As described the engine is of course a conventional two shaft design; it will be understood that the same layout could form the gas generator portion of a fan engine of the two or three shaft variety.

Both of the turbines 1 4 and 1 5 consist of the same basic components. They thus have nozzle guide vanes which direct the hot gases onto the rotor blades of a turbine rotor so as to drive this rotor. In the case of the high pressure turbine the nozzle guide vanes are shown in Fig. 1 at 20 and the rotor blades at 21. Fig. 2 iliustrates the radially outer part of the vanes 20 and blades 21 in an enlarged view.

In Fig. 2 part of the aerofoil of the vanes 20 can be seen at 22 while the platform 23 provides the outer boundary of the gas flow through the stator stage. The tip portion of the rotor blade aerofoil is visible at 24, and the rotor tip shroud can be seen at 25. These shrouds are in the form of segments, each carried by one aerofoil, which abut to form a complete annulas. Their function is to form an outer boundary for the airflow through the rotor stage, and to prevent air leaking over the blade tips. Clearly they perform this duty effectively since they are integral with the blade tips, but there is still the possibility of leakage over the outside of the shrouds and it is with this leakage that the invention is concerned.

The inner surfaces 26 of the shrouds 25 are of smooth aerodynamic profile so that they present as little obstruction as possible to the flow of gas across them. The radially outer surfaces, however, are formed in a more complex fashion and cooperate with fixed annular structure outside the ring of shrouds to reduce the effect of leakage flow.

Thus the forward region 27 of the shroud is made as a plain cylindrical surface which cooperates with the cylindrical inner surface 28 of static structure to form a small clearance which provides a partial seal and limits the flow therethrough. Of course, because clearance has to be left to allow for differential expansions etc the seal between the surfaces at 27 and 28 cannot be perfect but will allow a degree of leakage. This leakage enters the passage at a static pressure P equal to that in the main gas flow. In the present instance the inner surface 28 of the static structure is formed by the inner ends of the radially extending cells of an annular honeycomb structure which can be seen at 29.The use of such a honeycomb is advantageous in that it provides a compliant layer which will deform if there should be accidental contact between the shroud surface 27 and the honeycomb.

The cylindrical surface 28 only extends over the forward portion of the shroud, and ends at 30 to leave the following, intermediate portion of shroud surface 31 open to a plenum chamber 32. The surface 31 is gently concavely curved and smooth so that the leakage air proceeding from the gap between surfaces 27 and 28 will attach itself to the surface by virtue of the Coanda effect. The plenum chamber 32 is formed in the juncture of the end-surface 30 of the forward part of the static structure and the inner, cylindrical surface 33 of a further layer of honeycomb 34 forming a static structure. The purpose of the plenum chamber is to allow the pressure of the leakage air to stabilise at a value, P2, whose significance will become evident in due course.P2 is of course less than P1, the pressure at entry to the passage between the surfaces 27 and 28.

In its downstream region the shroud is formed with a thickened portion 35 within which are formed enclosed ducts 36, each extending from an inlet 37 forming a continuation of the surface 31, to outlets at 38. As can be seen from Fig. 2 the inlets 35 and ducts 36 are shaped to encourage the leakage flow to flow through them, while as can be seen from Fig. 3 they curve to direct the leakage flow in substantially the same direction as outlet as that of the main gas flow through the blade aerofoils. This main flow is shown in large arrows while the leakage flow is depicted by small arrows.

The shaping of the ducts 36 is also such as to produce no loss of static pressure in the flow therethrough. This means that the pressure at the outlet 38 is the same as that at the inlets 37, i.e. P1 must also therefore be the static pressure in the main gas flow just downstream of the edge of the shroud 25.

Since the pressure at the outlet 38 equals that in the chamber 32, there should not be any tendency for the leakage gas or any other gas to flow outside the thickened portion 35 between its outer surface 39 and the inner surface 33 of the honeycomb 34. However, to reduce the chance of leakage occuring in either direction during off-design conditions it is preferably to maintain a minimal clearance between the surfaces 32 to 39 similar to that between the surfaces 27 and 28.

It will be seen that the structure described allows the total leakage air to be entrained into ducts and disposed of with the same direction and static pressure as the main gas flow. This reduces the losses due to the leakage.

One further point illustrated in Fig. 2 is that the duct 36 may be used to provide a convenient disposal route for spent cooling air from the blade aerofoil. In the crude scheme illustrated in Fig. 2 a cooling air passage 40 extends longitudinally of the aerofoil and its cooling air flow exits through a hole 41 into one c-the passages 36. Clearly this could be done with any of the cooling air passages in the trailing region of the aerofoil, or with some complication in connecting passages could be applied to leading edge tubes. This then has the advantage that the energy in the cooling air may be partially recovered and this air does not have substantially deleterious secondary effect

Claims (6)

1. A shrouded rotor for a gas turbine having a shroud outer surface comprising a forward portion which cooperates with static structure to form a close annular clearance, an intermediate portion providing a smooth surface to which leakage gas may attach to the Coanda effect, a plurality of inlets into which said leakage gas flows directly from said intermediate portion, and enclosed ducts leading from said inlets across the rearward portion of the shroud, the ducts being curved and shaped to direct the leakage flow in the same direction as the main gas flow leaving the rotor blades.

2. A shrouded rotor as claimed in claim 1 and in which the enclosed ducts are shaped to ensure the static pressure at their outlets equals that at their inlets.

3. A shrouded rotor as claimed in claim 2 and in which said ducts are formed in a thickened part of said shroud whose outer surface cooperates with static structure to provide a partial seal.

4. A shrouded rotor as claimed in any one of the preceding claims and comprising a plenum chamber formed between said inter mediate portions of the shroud outer surface and static structure.

5. A shrouded rotor as claimed in any one of the preceding claims and comprising passages allowing spent cooling air to flow from the interior of the aerofoil of the blades of the rotor into said ducts.

6. A shrouded rotor substantially as hereinbefore particularly described with reference to the accompanying drawings.