[0001] This invention relates to a turbine and is particularly concerned with minimising
the effects of cooling air leakage in a turbine which is air cooled.
[0002] It is common practice to provide at least some of the aerofoil blades in the turbine
of a gas turbine engine with some form of internal cooling. Typically, that cooling
is provided by cool air which has been tapped from the air compression section of
the engine. It is important that the cooling air is directed to the interiors of the
blades which require cooling, without leaking into regions where it could have an
adverse effect upon the overall operating efficiency of the turbine.
[0003] One region in which air leakage problems can occur is between turbine discs carrying
turbine blades and structures adjacent those discs. Typically, cooling air from the
compression section of the gas turbine engine flows along the radially inner regions
of the engine before being deflected in radially outward directions between the disc
and structure adjacent thereto. The air is then directed into cooling passages provided
within turbine blades carried by one of the discs.
[0004] Conventionally, in order to inhibit the leakage of cooling air into the hot gas stream
which operationally flows over the turbine blades, an annular gas seal is positioned
between the disc and the structure adjacent thereto. Typically, the seal is of the
labyrinth type comprising annular, axially extending parts provided on both the disc
and the adjacent structure which cooperate to define a barrier in the form of a tortuous
path for air attempting to flow in a radially outward direction. While such seals
are partially effective in providing a barrier to air flowing in radially outward
directions, there remains a certain degree of undesirable leakage of cooling air into
the hot gas stream.
[0005] Some of that leaked cooling air passes through gaps defined between the sub-platform
regions of adjacent turbine blades. This flow may be inhibited by the provision of
lock plates which are primarily positioned on the disc to inhibit blade movement.
Such an arrangement is described in GB1209419, US-A-4 953 888 and US-A-2 920 865.
[0006] It is an object of the present invention to provide a turbine in which the deleterious
effects of such cooling air leakage into the hot gas stream, particularly such leakage
past the lock plates, have upon the overall efficiency of the turbine are reduced.
[0007] According to the present invention, a turbine comprises at least one rotatable disc
carrying an annular array of aerofoil blades, each of said blades having an aerofoil
portion operationally located in an annular gas passage extending through said turbine
for the flow of gas through said turbine, means being provided to direct cooling air
into passages provided internally of said aerofoil blades to provide cooling thereof,
said cooling air operationally flowing, at least partially, in radially outward directions
over at least part of the upstream external surface of said disc prior to a part thereof
being diverted to provide cooling of said aerofoil blades, a plurality of lock plates
being provided on the downstream side of said disc to provide locking of said blades
on said disc, means being provided radially inwardly of said aerofoil portions to
direct at least some of the remaining cooling air towards said lock plates, each of
said lock plates having an aperture therein, deflection means being provided on each
of said lock plates and associated with said aperture in said lockplate to deflect
said cooling air directed towards said lock plates into a region downstream of said
disc in a direction having a circumferential component generally opposite to that
in which said disc operationally rotates.
[0008] Each of said aerofoil blades is preferably provided with a shank radially inwardly
of its aerofoil portion said means to direct at least some of said remaining cooling
air towards said lock plates comprises a plurality of passages defined by the shanks
of said aerofoil blades.
[0009] Each of said deflector means is preferably in the form of cowling attached to its
associated lock plate.
[0010] The present invention will now be described, by way of example, with reference to
the accompanying drawings in which:
Figure 1 is a partially broken away perspective view of part of turbine in accordance
with the present invention.
Figure 2 is a view on arrow A of Figure 1.
Figure 3 is a perspective view of a portion of the embodiment shown in Figure 2.
[0011] Referring to Figure 1, a turbine 10 for a gas turbine engine (not shown) is shown
in a partial, broken away view. It is of generally conventional configuration comprising
an annular array of stator vanes 11 which are located upstream of an annular array
of aerofoil rotor blades 12. The turbine 10 is provided with several more axially
alternate annular arrays of stator vanes and aerofoil blades, but these have been
omitted in the interests of clarity.
[0012] The stator vanes 11 each comprise an aerofoil portion 13 which is situated in an
annular gas passage 14 which extends through the turbine 10. The radially inner and
outer extents of the gas passage 14 in the region of the vane aerofoil portions 13
are respectively defined by inner and outer platforms 15 and 16 which are integral
with the aerofoil portion 13. The inner platforms 15 of circumferentially adjacent
vanes 11 abut to define a generally continuous gas passage-defining surface as do
the outer platforms 16.
[0013] Each stator vane 11 is respectively supported at its radially inner and outer extents
by the turbine casing 17 and an inner support structure 18.
[0014] The aerofoil rotor blades 12 are mounted on a common disc 19 which is mounted for
rotation within the turbine 10. Each aerofoil rotor blade 12 comprises an aerofoil
portion 20 which, like the aerofoil portions 13 of the stator vanes 11, is situated
in the annular gas passage 14. Radially inner and outer platforms 21 and 22 respectively
on each blade 20 serve to define local portions of the gas passage 14.
[0015] Each aerofoil blade 12 is provided with a shank 23 radially inwardly of its inner
platform 21 which interconnects the remainder of the blade 12 with a firtree root
portion 24. The firtree portion 24 locates in a correspondingly shaped cut-out portion
25 provided in the periphery of the disc 19, thereby providing radial constraint for
the aerofoil blade 12. The shanks 23 are circumferentially narrower than their associated
firtree root portions 24 so that a circumferential gap 23a is defined between adjacent
shanks 23.
[0016] In order to provide axial constraint of each of the aerofoil blades 12, an annular
array of lockplates 40 is provided adjacent their firtree root portions 24. Each lockplate
40 is planar and locates at its radially outer extent in a radially inwardly directed
groove 41 defined by its adjacent aerofoil blade 12 and at its radially inner extent
in a radially outwardly directed annular groove 42 defined by the disc 19. The lockplates
40 are well known as such in the construction of turbines.
[0017] In operation, extremely hot gases flow through the annular gas passage 14. They act
upon the aerofoil portions 20 of the aerofoil blades 12 to bring about the rotation
of the turbine disc 19. Since the gases are extremely hot, internal air cooling of
the vanes 11 and the aerofoil blades 12 is necessary. Both the vanes 11 and the aerofoil
blades 12 are hollow in order to achieve this. In the case of the vanes 11, cooling
air derived from a suitable source is directed into their radially outer extents through
apertures 26 provided in their radially outer platforms 16. The air then flows through
the vanes 11 to exhaust therefrom through a large number of small apertures 27 provided
in the vane aerofoil portions 13 into the gas stream flowing through the annular gas
passage 14. This provides both convection cooling of the vane 11 interiors and film
cooling of their external aerofoil portion 13 surfaces.
[0018] Similarly, the aerofoil blades 12 are cooled by a flow of cooling air into their
interiors which is exhausted through a large number of small holes 28 in their aerofoil
portions 20. However, in this case, the cooling air is directed into the aerofoil
blade 12 interiors from their radially inner extents. The air flows in a radially
outward direction over the upstream surface 29 of the disc 19 to enter a plurality
of generally radially extending passages 30 in the disc 19 periphery. One passage
30 is associated with each firtree root cut-out portion 25 so that a flow of cooling
air is directed to the root portion 25 of each of the aerofoil blades 12. A passage
(not shown) in each root portion 24 directs cooling air into the blade 12 interior
to provide convection cooling of the blade 12. It then flows through the small holes
28 to provide film cooling of the aerofoil portion. The cooling air then mixes with
the gases flowing through the annular gas passage 14.
[0019] The above mentioned way of air cooling the vanes 11 and aerofoil blades 12 well known
as such.
[0020] In order to ensure that cooling air does not by-pass the blade feed passages 30 and
prematurely enter the hot gas stream flowing through the annular gas passage 14, an
annular seal 31 is provided between the upstream face 29 of the disc 19 and the downstream
face 32 of the fixed turbine structure 34 which supports the radially inner extents
of the vanes 11. The seal 31 is of the well known labyrinth type comprising a generally
axially extending element 35 carried by the disc 19 and a corresponding reception
element 36 carried by the fixed turbine support structure 34.
[0021] Unfortunately, labyrinth seals such as that described above are not as efficient
at providing a barrier to gas flow as would normally be desirable. Consequently, some
cooling air inevitably leaks through the labyrinth seal 31 into the region 37 between
the firtree root portions 24 and fixed turbine support structure 34. Under normal
circumstances, this leaked cooling air would pass into the annular gas passage 14
and have a prejudicial effect upon the gases operationally flowing through that passage
14. However, in accordance with the present invention, the leaked cooling air is utilised
in a more effective and efficient manner.
[0022] Each of the lock plates 40 is provided with an aperture 43. Each aperture 43 is partially
enclosed by a cowling 44 which is bonded to its associated lock plate 40a and is of
part-oval configuration in plan view. The centre portion 45 of each cowling 44 is
raised so as to define an outlet 46 adjacent one edge of its associated lock plate
40.
[0023] In operation, cooling air from the region 37 flows through circumferential gaps 23a
defined between the blade shanks 23. However, that cooling air then flows through
the apertures 43 in the lock plates 40. Each cowling 44 is so configured that the
cooling air flow is deflected in a generally circumferential direction which is opposite
to the direction of rotation 39 of the disc 19. Consequently, the deflected airflow
serves the same function as the airflow exhausted from the passages 37 in improving
overall turbine efficiency.
1. A turbine (10) comprising at least one rotatable disc (19) carrying an annular array
of aerofoil blades (12), each of said blades (12) having an aerofoil portion (20)
operationally located in an annular gas passage (14) extending through said turbine
(10) for the flow of gas through said turbine (10), means (30) being provided to direct
cooling air into passages provided internally of said aerofoil blades (12) to provide
cooling thereof, said cooling air operationally flowing, at least partially, in radially
outward directions over at least part of the upstream external surface (29) of said
disc (19) prior to a part thereof being diverted to provide cooling of said aerofoil
blades (12), a plurality of lock plates (40) being provided on the downstream side
of said disc (19) to provide locking of said blades (12) on said disc (19), means
being provided radially inwardly of said aerofoil portions (20) to direct at least
some of the remaining cooling air towards said lock plates (40) characterised in that each of said lock plates (40) has an aperture (43) therein, deflection means (44)
being provided on each of said lock plates (40) and associated with said aperture
(43) in said lockplate to deflect said cooling air directed towards said lock plates
(40) into a region downstream of said disc (40) in a direction having a circumferential
component generally opposite to that in which said disc (40) operationally rotates.
2. A turbine as claimed in claim 1 characterised in that each of said aerofoil blades (12) is provided with a shank (23) radially inwardly
of its aerofoil portion (20) said means to direct at least some of said remaining
cooling air towards said lock plates (40) comprises a plurality of passages (23a)
defined by the shanks of said aerofoil blades (12).
3. A turbine as claimed in claim 1 characterised in that each of said deflector means (44) is in the form of a cowling attached to its associated
lockplate (40).
1. Turbine (10) mit wenigstens einer drehbaren Scheibe (19), die eine ringförmige Anordnung
von Profil-Laufschaufeln (12) trägt, von denen jede einen Profilteil (20) aufweist,
der betriebsmäßig in einem ringförmigen Gaskanal (14) umläuft, der sich durch die
Turbine (10) erstreckt und von der Gasströmung durch die Turbine (10) durchströmt
wird, wobei Mittel (30) vorgesehen sind, um Kühltuft in Kanäle zu richten, die im
Inneren der Profil-Laufschaufeln (12) angeordnet sind, um diese zu kühlen und die
Kühlluft betriebsmäßig wenigstens teilweise radial nach außen über wenigstens einen
Teil der stromaufwärtigen äußeren Oberfläche (29) der Scheibe (19) strömt, bevor ein
Teil davon abgelenkt wird, um eine Kühlung der Profil-Laufschaufeln (12) zu bewirken,
und wobei mehrere Verriegelungsplatten (40) auf der stromabwärtigen Seite der Scheibe
(19) angeordnet sind, um die Laufschaufeln (12) auf der Scheibe (19) zu verriegeln,
und wobei Mittel radial innerhalb der Profilabschnitte (20) vorgesehen sind, um wenigstens
einen Teil der übrigen Kühlluft nach den Verriegelungsplatten (40) zu richten,
dadurch gekennzeichnet, daß jede Verriegelungsplatte (40) eine Öffnung (43) darin aufweist, daß Ablenkmittet
(44) auf jeder Verriegelungsplatte (40) vorgesehen und dieser Öffnung (43) in der
Verriegelungsplatte (40) zugeordnet sind, um die nach den Verriegelungsplatten (40)
gerichtete Strömung in einen Bereich stromab der Scheibe (19) in einer Richtung abzulenken,
welche eine in Umfangsrichtung verlaufende Komponente allgemein entgegengesetzt zu
jener Richtung hat, in der sich die Scheibe (19) betriebsmäßig dreht.
2. Turbine nach Anspruch 1,
dadurch gekennzeichnet, daß jede Profil-Laufschaufel (12) mit einem Schaft (23) radial innerhalb ihres Profilabschnitts
(20) versehen ist und daß die Mittel, die wenigstens einen Teil der verbleibenden
Kühlluft nach den Verriegelungsplatten (40) richten, aus mehreren Kanälen (23a) bestehen,
die durch die Schäfte der Profil-Laufschaufeln (12) definiert werden.
3. Turbine nach Anspruch 1,
dadurch gekennzeichnet, daß jedes Ablenkmittel (44) die Gestalt einer Verkleidung besitzt, die an der zugeordneten
Verriegelungsplatte (40) befestigt ist.
1. Turbine (10) comprenant au moins un disque rotatif (19) possédant une série en anneau
d'aubes profilées (12), chacune desdites aubes (12) ayant une partie profilée (20)
qui se situe en fonctionnement dans un passage de gaz en anneau (14) qui s'étend à
travers ladite turbine (10) pour laisser passer le flux de gaz dans la turbine (10)
en question, ce moyen (30) étant fourni pour diriger l'air de refroidissement dans
les passages munis intérieurement desdites aubes profilées (12) pour la refroidir,
ledit air de refroidissement se répand de manière opérationnelle, au moins en partie,
dans des directions radiales externes, sur au moins une partie de la surface externe
en amont (29) dudit disque (19) avant qu'une partie soit détournée pour refroidir
lesdites aubes profilées (12), une pluralité de tôles frein (40) installées sur le
côté en aval dudit disque (19) pour bloquer lesdites aubes (12) sur ledit disque (19),
ce moyen étant pourvu perpendiculairement à l'axe à l'intérieur desdites parties profilées
(20) pour diriger au moins les quantités d'air de refroidissement restantes vers lesdites
tôles frein (40), chacune desdites tôles frein (40) étant caractérisé par une ouverture, ce moyen de dérivation étant présent sur chacune desdites tôles frein
(40) et associé à ladite ouverture (43) de ladite tôle frein (40) pour dévier ledit
air de refroidissement vers lesdites tôles frein dans une région en aval dudit disque
(40) dans une direction ayant une composante circonférentielle généralement opposée
à cela et dans laquelle ledit disque (40) tourne de façon opérationnelle.
2. Turbine selon la revendication 1 caractérisée en ce que chacune desdites aubes profilées (12) est munie d'une tige (23) à l'intérieur de
ladite partie profilée (20), ledit moyen dirigeant au moins ledit air de refroidissement
restant à travers lesdites tôles frein (40) comporte une pluralité de passages (23a)
définis par les tiges desdites aubes profilées (12).
3. Turbine selon la revendication 1 caractérisée en ce que chacun desdits moyens de dérivation (44) est de la forme d'un capuchon solidaire
de sa tôle frein correspondante (40).