US20130177414A1

US20130177414A1 - Method and system for controlling the clearance at the blades tips of a turbine rotor

Info

Publication number: US20130177414A1
Application number: US13/701,700
Authority: US
Inventors: Damien Bonneau; Marc Croixmarie; Denis Denece Franck Roger; Bruno Robert Gaully
Original assignee: SNECMA SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 2010-06-03
Filing date: 2011-06-01
Publication date: 2013-07-11
Also published as: RU2012157775A; WO2011151602A1; RU2566510C2; CA2801193A1; EP2576994A1; FR2960905B1; BR112012030635A2; CN103003529A; CN103003529B; FR2960905A1

Abstract

A system and method for controlling clearance between tips of moving blades of an aeroplane gas-turbine engine and a turbine shroud of an outer casing surrounding the blades. The method includes controlling, according to an operating speed of the engine, a valve positioned in an air duct opening at a stage of the compressor of the engine and leading into a control housing positioned around the outer surface of the turbine shroud and supplied with air coming only from the compressor stage. The valve is opened to cool the turbine shroud during a high-speed phase corresponding to takeoff and climb phases of an aeroplane propelled by the engine and during a nominal-speed phase following the high-speed phase and corresponding to a cruise phase of the aeroplane.

Description

BACKGROUND OF THE INVENTION

The present invention relates to the general field of turbomachine turbines for aeroplane gas-turbine engines. It relates more particularly to control of the clearance between, on the one hand, the tips of moving blades of a turbine rotor and, on the other hand, a turbine shroud of an outer casing surrounding the blades.
To increase the performance of a turbine, it is a known practice to minimize to the extent possible the clearance existing between the tip of the turbine blades and the shroud that surrounds them. This blade tip clearance is dependent on the dimensional variations between the rotating portions (disk and blades constituting the turbine rotor) and the fixed portions (outer casing, including the turbine shroud that is part of it). These dimensional variations are both of thermal origin (connected with the variations in temperature of the blades, the disk and the casing) and of mechanical origin (connected in particular with the centrifugal force exerted on the turbine rotor).
To minimize this clearance, it is a known practice to resort to active control systems. These systems generally operate by directing onto the outer surface of the turbine shroud cool air taken from a compressor and/or from the fan of the turbine engine. The cool air delivered to the outer surface of the turbine shroud has the effect of cooling the latter and thus limiting its thermal expansion. Such active control is controlled for example by the full authority control system (or FADEC) of the turbine engine and is a function of the different operating levels thereof.
Document EP 1,860,281 describes an example of an active control system wherein air taken from the fan of the turbine engine cools the turbine shroud during the cruising flight phases. Such a system, however, exhibits numerous disadvantages such as its considerable use of space inside the nacelle of the turbine engine, the strong dependency of its effectiveness on aerothermal conditions existing within the nacelle and performance losses connected with the tapping of air flow from the fan which does not participate in providing thrust.
Another active control system consists of tapping air at two different stages of the compressor of the turbine engine and modulating the delivery of each of these tapped streams to control the temperature of the mixture to be directed onto the outer surface of the turbine shroud. Though effective, such a system exhibits the disadvantage of resorting to a complex and bulky valve to modulate the cooling air stream. In particular, in the case of an application to a small sized turbine engine, the use of such a valve is not optimal in terms of mass and cost.

OBJECT AND SUMMARY OF THE INVENTION

The present invention therefore has as its principal object to correct such disadvantages by proposing an active control solution that is minimalist in terms of mass and cost.
This goal is achieved by a method for controlling the clearance between the tips of moving blades of a turbine rotor of an aeroplane gas-turbine engine and a turbine shroud of an outer casing surrounding the blades, the method comprising controlling, according to the operating speed of the engine, a valve positioned in an air duct opening at a compressor stage of the engine and leading into a control housing positioned around the outer surface of the turbine shroud, said control housing being supplied with air taken solely from said compressor stage.
In accordance with the invention, the valve is opened to cool the turbine shroud of the outer casing during a high-speed operating phase corresponding to the takeoff and climb phases of an aeroplane propelled by the engine and during a nominal-speed phase following the high-speed phase and corresponding to the cruise phase of said aeroplane.
Correlatively, the invention provides a system for controlling clearance between blade tips of a turbine rotor of an aeroplane gas-turbine engine and a turbine shroud of an outer casing surrounding the blades, the system comprising an air duct intended to open at a compressor stage of the engine and to lead to a control housing positioned around the outer surface of the turbine shroud and intended to be supplied with air tapped solely from said compressor stage, a valve positioned in the air duct, and a circuit capable of controlling the valve so as to open it during a high-speed operating phase corresponding to the takeoff and climb phases of an aeroplane propelled by the engine and during a nominal-speed phase following the high-speed phase and corresponding to the cruise phase of said aeroplane.
By high-speed phase, it is meant a speed phase greater than the nominal-speed phase of the turbine engine. In an aeroplane gas-turbine engine, the nominal-speed phase is the flight cruise phase which is chosen during the greater portion of the flight, and the high-speed phase is a phase higher than the flight cruise phase, used in particular during the takeoff and climb phases of the aeroplane.
The invention is remarkable in particular in that it uses a single air tap at the compressor that guarantees a sufficient pressure differential to ensure a delivery of cool air to the turbine shroud (the control housing exhibits only a single and unique air supply source). In addition, this air tapped at the compressor is delivered only into the control housing and does not supply any other components of the engine. Also, when the valve is closed, no air is actually tapped from the compressor, which limits head loss within the same. In this manner it is possible to minimize the air ducts and air taps in the engine and to employ the simplest possible valve (in terms of structure and of control). The result is a low cost control system with low mass.
Preferably, the valve is closed during a flight idle-speed phase following the nominal-speed phase and corresponding to the approach phase of the aeroplane prior to landing.
Preferably also, the valve is closed during a ground idle-speed phase preceding the nominal-speed phase and corresponding to the taxiing phase of the aeroplane prior to takeoff.
The idle-speed phase is a level lower than the nominal-speed phase of the turbomachine. In an aeroplane gas-turbine engine, the idle-speed phase is therefore a phase lower than the flight cruise phase.
Advantageously, the air delivery leading to the outer surface of the turbine shroud is gradually reduced during transition between the high-speed phase and the nominal-speed phase. In the case of a variable position valve, such a progressive reduction of the air delivery can be obtained by gradually closing the valve. In the case of an on-off valve, the gradual reduction of the air delivery can be obtained by alternation of open and closed phases of the valve.
The invention also provides an aeroplane gas-turbine engine having a clearance control system as previously defined.

BRIEF DESCRIPTION OF THE DRAWINGS

Other features and advantages of the present invention will appear from the description given below, with reference to the appended drawings which illustrate an embodiment of it that is in no way limiting in nature. In the figures:

FIG. 1 is a schematic longitudinal section view of a gas turbine aero-engine equipped with a control system according to the invention;

FIG. 2 is a magnified view of the engine of FIG. 1 showing in particular its high-pressure turbine;

FIG. 3 shows the curves illustrating one variation of the operating level and the corresponding variations in the radial dimension of the rotor and of the stator in a gas turbine aero-engine; and

FIGS. 4A through 4C show curves representative of examples of control of an on-off valve used in an embodiment of the control system according to the invention.

DETAILED DESCRIPTION OF ONE EMBODIMENT

FIG. 1 shows schematically a turbojet 10 of the bypass, twin spool type to which the invention particularly applies. Of course, the invention is not limited to this particular type of gas turbine aero-engine.
In well-known fashion, the turbojet 10 with longitudinal axis X-X includes in particular a fan 12 which delivers a stream of air into a primary flow passage 14 and into a secondary flow passage 16 coaxial with the primary flow passage. From upstream to downstream in the direction of flow of the gas stream passing through it, the primary flow passage 14 includes a low-pressure compressor 18, a high-pressure compressor 20, a combustion chamber 22, a high-pressure turbine 24 and a low-pressure turbine 26.
As shown more precisely by FIG. 2, the high-pressure turbine 24 of the turbojet includes a rotor consisting of a disk 28 on which is mounted a plurality of moving blades 30 positioned in the passage of the primary flow 14. The rotor is surrounded by a turbine casing 32 including a turbine shroud 34 carried by an outside turbine casing 36 through mounting braces 37.
The turbine shroud 34 may be formed of a plurality of adjacent sectors or segments. On the inner side, it is provided with a layer 34 a of abradable material and surrounds the blades 30 of the rotor, leaving a clearance 38 with their tips 30 a.
In accordance with the invention, a system is provided that allows control of the clearance 38 by reducing, in a controlled manner, the inner diameter of the outer turbine casing 36.
To this end, a control housing 40 is positioned around the outer turbine casing 36. This housing receives cold air by means of an air duct 42 opening at its upstream end into the passage of the primary flow at one of the stages of the high-pressure compressor 20 (for example by means of a scoop that is known per se and therefore not shown in the figures). In particular, the control housing is supplied with air by only this single tap at the compressor (there are no other air sources supplying the housing).
The cool air circulating in the air duct 42 is entirely discharged onto the outer turbine casing 36 (for example by means of multiple perforation of the walls of the control housing 40) causing it to cool and hence reducing its inner diameter. In particular, the air tapped at the high pressure compressor stage does not supply any other components than the control housing.
As shown in FIG. 1, a valve 44 is positioned in the air duct 42. This valve is controlled by the full authority control system (or FADEC) 46 of the turbojet depending on the operating levels of the turbojet.
By controlling the valve 44 as a function of the different phases of flight of the aeroplane, it is thus possible to vary over the course of a mission the inner diameter of the outer turbine casing 36—and hence the inner diameter of the turbine shroud 34—and consequently to control the clearance existing between the turbine shroud and the tip of the blades 30 of the high-pressure turbine rotor.
FIG. 3 shows the variation of this clearance 38 over the course of a typical mission of the aeroplane as obtained by the control system and method according to the invention.
On this figure different curves are shown, namely: a curve 100 illustrating the rotation speed of the high-pressure spool of the turbojet, a curve 200 illustrating the outer diameter of the high-pressure turbine rotor (disk 28 and blades 30), a curve 300 illustrating the inner diameter of the stator of the high-pressure turbine (outer turbine casing 36 and turbine shroud 23) as controlled by the control system according to the invention, and a curve 300 a (dotted) illustrating the inner diameter of the stator as it would be in the absence of control.
These different curves are shown according to the different phases of operation of the turbojet representative of a typical mission, namely: a ground idle-speed phase GI (corresponding to the taxiing phase of the aeroplane prior to takeoff), followed by a high-speed phase TO+CL (corresponding to the takeoff and climb phases of the aeroplane), followed by a nominal-speed phase CR (corresponding to the cruise phase of the aeroplane), followed by a flight idle-speed phase FI (corresponding to the approach of the aeroplane prior to landing), followed by a reverse thrust phase REV (corresponding to braking of the aeroplane on the ground), followed by another ground idle-speed phase GI.
As shown by the curve 100, it will be noted that the high-speed phase TO+CL takes place at a speed higher than the nominal-speed of the turbojet (CR phase). The idle-speed phases (ground and flight) take place at speeds lower than the nominal-speed of the turbojet, the flight idle-speed phase FI having a speed that is also lower than that of the ground idle-speed phase GI. It will also be noted that the nominal-speed phase CR is adopted during the greater portion of the mission.
Control of the valve 44 according to the invention is as follows:

- During the ground idle-speed phase GI, the valve is closed and the inner diameter of the stator remains substantially unchanged. During the transition phase between the GI phase and the TO+CL phase, the valve is still closed and the stator is free to expand under the influence of the hot air in the passage of the primary flow. During this same transition phase, it will be noted that the rotor begins to expand mechanically under the influence of centrifugal force.
- During the high-speed TO+CL phase, the valve 44 is open, which cools the stator and consequently reduces its inner diameter. The clearance is small and considerably reduced compared to what it would be in the absence of control. The result is a strong gain in performance in this phase. It will be noted more precisely that the opening of the valve takes place after the pinch point is passed, that is once the transition point is reached between the mechanical expansion phase of the rotor and the thermal expansion of the rotor.
- During the nominal-speed phase CR, the valve 44 is held open to cool the stator and to thus obtain a small clearance, which is beneficial for the performance of the engine.

It will be noted that at the end of the TO+CL phase, during transition to the nominal-speed phase CR, the delivery of air directed toward the stator is progressively reduced. It will also be noted that during the CR phase, this same air delivery can be greater or smaller depending on the flight altitude. Different ways of obtaining a reduction in the delivery of air will be detailed later in connection with FIG. 4.

- During the flight idle-speed phase FI, the valve 44 is again closed so that the stator is free to expand under the influence of the hot air flowing in the passage of the primary flow. The clearance opens during this approach phase of the aeroplane prior to landing in order to provide for a contingency requiring the aeroplane to take off again (and hence a resumption of high speed).
- Finally, during the thrust reversal REV and ground idle-speed GI phases, the valve 44 is kept closed.

Different valve structures can be used to implement such a clearance control. The valve 44 can be of the controlled delivery type (under FADEC control), which facilitates the control of the air delivery directed toward the stator particularly at the end of the TO+CL phase and in the CR phase.
However, for cost and reliability reasons, it is advantageous to resort to an on-off type valve. To obtain modulation of the air delivery directed toward the stator with this type of valve, it is possible to alternate opening and closing phases of the valve.
FIGS. 4A through 4C show different deliveries that can be obtained with this type of control of the on-off valve. In these figures are shown square wave signals illustrating, in the ordinate, the position of the valve (0=valve open and 1=valve closed) and in the abscissa the time, t. The curves Ca through Cc illustrate the average air delivery supplied by the valve depending on the different open times thereof: the more the valve is open longer (at each opening cycle), the higher is the average air delivery supplied by the valve (and conversely).
In this manner it is understood that by operating on the one hand on the opening frequency and on the other hand on the cyclic opening/closing ratio of the valve, it is possible to obtain a variation of the average delivery of the air directed toward the stator.
Different on-off type valve architectures are well known to those skilled in the art and will therefore not be described here. Preferably, an electrically controlled valve will be selected which will remain in the closed position in the absence of a supply of electrical power (thus guaranteeing that the valve will remain closed in the event of a control failure).

Claims

1-10. (canceled)

11. A method for controlling clearance between tips of moving blades of a turbine rotor of an aeroplane gas-turbine engine and a turbine shroud of an outer casing surrounding the blades, the method comprising:

controlling, according to an operating speed of the engine, a valve positioned in an air duct opening into a compressor stage of the engine and leading into a control housing positioned around an outer surface of the turbine shroud, the control housing being supplied with air coming only from the compressor stage; and

opening the valve for cooling the turbine shroud of the outer casing during a high-speed operating phase corresponding to takeoff and climb phases of an aeroplane propelled by the engine and during a nominal-speed phase following the high-speed phase and corresponding to a cruise phase of the aeroplane.

12. The method according to claim 11, wherein the valve is closed during a flight idle-speed phase following the nominal-speed phase and corresponding to an approach phase of an aeroplane prior to landing.

13. The method according to claim 11, wherein the valve is closed during a ground idle-speed phase preceding the nominal-speed phase and corresponding to a taxi phase of the aeroplane prior to takeoff.

14. The method according to claim 11, wherein delivery of air leading toward the outer surface of the turbine shroud is gradually reduced during a transition between the high-speed phase and the nominal-speed phase.

15. The method according to claim 14, wherein the valve is an adjustable position valve, the gradual reduction of the air delivery leading toward the outer surface of the turbine during the transition being obtained by gradually closing the valve.

16. The method according to claim 14, wherein the valve is an on-off valve, the gradual reduction in the delivery of air leading toward the outer surface of the turbine shroud during the transition being obtained by alternating opening and closing phases of the valve.

17. A system for controlling clearance between tips of moving blades of a turbine rotor of an aeroplane gas-turbine engine and a turbine shroud of an outer casing surrounding the blades, the system comprising:

an air duct configured to open at a compressor stage of the engine and to lead into a control housing surrounding an outer surface of the turbine shroud and configured to be supplied with air coming only from the compressor stage;

a valve positioned in the air duct; and

a circuit configured to control the valve for opening the valve during a high-speed phase corresponding to takeoff and climb phases of an aeroplane propelled by the engine and during a nominal-speed phase following the high-speed phase and corresponding to a cruise phase of the aeroplane.

18. The system according to claim 17, wherein the valve is an adjustable position valve.

19. The system according to claim 17, wherein the valve is an on-off valve.

20. An aeroplane gas-turbine engine comprising a clearance control system according to claim 17.