U.S. patent application number 13/701700 was filed with the patent office on 2013-07-11 for method and system for controlling the clearance at the blades tips of a turbine rotor.
This patent application is currently assigned to SNECMA. The applicant listed for this patent is Damien Bonneau, Marc Croixmarie, Franck Roger, Denis Denece, Bruno Robert Gaully. Invention is credited to Damien Bonneau, Marc Croixmarie, Franck Roger, Denis Denece, Bruno Robert Gaully.
Application Number | 20130177414 13/701700 |
Document ID | / |
Family ID | 43471088 |
Filed Date | 2013-07-11 |
United States Patent
Application |
20130177414 |
Kind Code |
A1 |
Bonneau; Damien ; et
al. |
July 11, 2013 |
METHOD AND SYSTEM FOR CONTROLLING THE CLEARANCE AT THE BLADES TIPS
OF A TURBINE ROTOR
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.
Inventors: |
Bonneau; Damien; (Melun,
FR) ; Croixmarie; Marc; (Auvernaux, FR) ;
Denece; Franck Roger, Denis; (Saint Michel Sur Orge, FR)
; Gaully; Bruno Robert; (Marolles en Hurepoix,
FR) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Bonneau; Damien
Croixmarie; Marc
Denece; Franck Roger, Denis
Gaully; Bruno Robert |
Melun
Auvernaux
Saint Michel Sur Orge
Marolles en Hurepoix |
|
FR
FR
FR
FR |
|
|
Assignee: |
SNECMA
Paris
FR
|
Family ID: |
43471088 |
Appl. No.: |
13/701700 |
Filed: |
June 1, 2011 |
PCT Filed: |
June 1, 2011 |
PCT NO: |
PCT/FR2011/051261 |
371 Date: |
March 11, 2013 |
Current U.S.
Class: |
416/1 ;
416/23 |
Current CPC
Class: |
F01D 5/02 20130101; F01D
11/24 20130101 |
Class at
Publication: |
416/1 ;
416/23 |
International
Class: |
F01D 5/02 20060101
F01D005/02 |
Foreign Application Data
Date |
Code |
Application Number |
Jun 3, 2010 |
FR |
1054366 |
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.
Description
BACKGROUND OF THE INVENTION
[0001] 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.
[0002] 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).
[0003] 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.
[0004] 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.
[0005] 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
[0006] 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.
[0007] 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.
[0008] 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.
[0009] 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.
[0010] 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.
[0011] 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.
[0012] 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.
[0013] 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.
[0014] 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.
[0015] 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.
[0016] The invention also provides an aeroplane gas-turbine engine
having a clearance control system as previously defined.
BRIEF DESCRIPTION OF THE DRAWINGS
[0017] 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:
[0018] FIG. 1 is a schematic longitudinal section view of a gas
turbine aero-engine equipped with a control system according to the
invention;
[0019] FIG. 2 is a magnified view of the engine of FIG. 1 showing
in particular its high-pressure turbine;
[0020] 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
[0021] 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
[0022] 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.
[0023] 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.
[0024] 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.
[0025] 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 34a of abradable material and surrounds the blades 30
of the rotor, leaving a clearance 38 with their tips 30a.
[0026] 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.
[0027] 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).
[0028] 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.
[0029] 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.
[0030] 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.
[0031] 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.
[0032] 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 300a (dotted) illustrating the inner diameter of the stator
as it would be in the absence of control.
[0033] 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.
[0034] 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.
[0035] Control of the valve 44 according to the invention is as
follows: [0036] 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. [0037] 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. [0038] 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.
[0039] 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. [0040] 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). [0041] Finally, during the
thrust reversal REV and ground idle-speed GI phases, the valve 44
is kept closed.
[0042] 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.
[0043] 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.
[0044] 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).
[0045] 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.
[0046] 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).
* * * * *