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Sommaire du brevet 2258206 

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Disponibilité de l'Abrégé et des Revendications

L'apparition de différences dans le texte et l'image des Revendications et de l'Abrégé dépend du moment auquel le document est publié. Les textes des Revendications et de l'Abrégé sont affichés :

  • lorsque la demande peut être examinée par le public;
  • lorsque le brevet est émis (délivrance).
(12) Brevet: (11) CA 2258206
(54) Titre français: CONFIGURATION DE CANAUX DE REFROIDISSEMENT POUR REFROIDIR LE BORD AVANT D'AILETTES DE TURBINE A GAZ
(54) Titre anglais: CONFIGURATION OF COOLING CHANNELS FOR COOLING THE TRAILING EDGE OF GAS TURBINE VANES
Statut: Réputé périmé
Données bibliographiques
(51) Classification internationale des brevets (CIB):
  • F01D 5/18 (2006.01)
  • F01D 9/02 (2006.01)
  • F01D 25/12 (2006.01)
(72) Inventeurs :
  • NORDLUND, RAYMOND S. (Etats-Unis d'Amérique)
  • HULTGREN, KENT G. (Etats-Unis d'Amérique)
  • SCOTT, ROBERT K. (Etats-Unis d'Amérique)
  • SINNOT, ZACHARY (Etats-Unis d'Amérique)
  • NORTH, WILLIAM E. (Etats-Unis d'Amérique)
  • WARD, STEVEN D. (Etats-Unis d'Amérique)
(73) Titulaires :
  • WESTINGHOUSE ELECTRIC CORPORATION (Etats-Unis d'Amérique)
(71) Demandeurs :
  • WESTINGHOUSE ELECTRIC CORPORATION (Etats-Unis d'Amérique)
(74) Agent: SMART & BIGGAR
(74) Co-agent:
(45) Délivré: 2006-06-27
(86) Date de dépôt PCT: 1998-03-25
(87) Mise à la disponibilité du public: 1998-10-22
Requête d'examen: 2003-03-12
Licence disponible: S.O.
(25) Langue des documents déposés: Anglais

Traité de coopération en matière de brevets (PCT): Oui
(86) Numéro de la demande PCT: PCT/US1998/006039
(87) Numéro de publication internationale PCT: WO1998/046860
(85) Entrée nationale: 1998-12-14

(30) Données de priorité de la demande:
Numéro de la demande Pays / territoire Date
08/843,414 Etats-Unis d'Amérique 1997-04-15

Abrégés

Abrégé français

Appareil servant à refroidir la partie bord avant d'une ailette de turbine à gaz. Deux passages s'étendant radialement et connectés à l'anneau de renforcement externe dirigent le fluide de refroidissement jusqu'à un plénum formé approximativement à mi-portée et adjacent au bord avant. Deux réseaux de passages pour le fluide de refroidissement s'étendent depuis le plénum. Un réseau s'étend radialement vers l'extérieur en direction de l'anneau de renforcement externe alors que l'autre réseau s'étend radialement vers l'intérieur en direction de l'anneau de renforcement interne. Le plénum distribue le fluide de refroidissement aux deux réseaux de passages de sorte qu'il s'écoule radialement vers l'intérieur et vers l'extérieur jusqu'à des collecteurs formés dans les anneaux de renforcement interne et externe. Les collecteurs envoient le fluide de refroidissement usé jusqu'à un passage de sortie.


Abrégé anglais





An apparatus for cooling the trailing edge portion
of a gas turbine vane. Two radially extending passages
connected to the outer shroud direct cooling fluid to a plenum
formed about mid-span adjacent the trailing edge. Two
arrays of cooling fluid passages extend from the plenum.
One array extends radially outward toward the outer shroud
while the other array extends radially inward toward the
inner shroud. The plenum distributes the cooling fluid to the
two arrays of passages so that it flows radially inward and
outward to manifolds formed in the inner and outer shrouds,
The manifolds direct the spent cooling fluid to a discharge
passage.

Revendications

Note : Les revendications sont présentées dans la langue officielle dans laquelle elles ont été soumises.





9

CLAIMS:

1. An airfoil for a turbomachine, comprising:
a) a leading edge and a trailing edge;
b) first and second ends, said first end
disposed radially outward from said second end;
c) first and second side walls;
d) a first passage formed between said first
and second sidewalls, said first passage having
an inlet for receiving a flow of a cooling fluid
directed to said airfoil;
e) a plenum formed between said side walls
and disposed between said first and second ends,
said plenum in flow communication with said first
passage;
f) a plurality of second passages in flow
communication with said plenum, said second
passages extending in a substantially radial
direction from said plenum toward said first end;
and
g) a plurality of third passages in flow
communication with said plenum, said third
passages extending in a substantially radial
direction from said plenum toward said second
end.

2. The airfoil according to claim 1, wherein
said plenum is disposed adjacent said trailing edge
approximately midway between said first and second ends.





10

3. The airfoil according to claim 1, wherein
said second and third passages form an array of passages
disposed adjacent said trailing edge.

4. The airfoil according to claim 1, further
comprising a first manifold for collecting cooling fluid
discharged from said second passages.

5. The airfoil according to claim 1, further
comprising an outlet for discharging said cooling fluid
from said airfoil, and means for directing said cooling
fluid collected by said first manifold to said airfoil
outlet.

6. The airfoil according to claim 5, wherein
said fluid directing means comprises a fourth passage in
flow communication with said first manifold.

7. The airfoil according to claim 6, further
comprising a first shroud affixed to one of said ends, and
wherein said fourth passage is formed in said first shroud.

8. The airfoil according to claim 6, wherein
said fourth passage extends in a direction substantially
perpendicular to the radial direction.

9. The airfoil according to claim 6, further
comprising a fifth passage formed between said first and
second walls.

10. The airfoil according to claim 9, further
comprising a rib extending between said first and second
sidewalls and separating said fifth passage from said first
passage.


11


11. The airfoil according to claim 9, wherein
said fourth passage is disposed so as to place said first
manifold in flow communication with said fifth passage.

12. The airfoil according to claim 7, further
comprising:
a) a second manifold for collecting cooling
fluid discharged from said third passages;
b) second cooling fluid directing means for
directing said cooling fluid collected by said
second manifold to said airfoil outlet.

13. The airfoil according to claim 12, wherein
said second cooling fluid directing means comprises a fifth
passage in flow communication with said second manifold,
and further comprising a second shroud affixed to the other
one of said ends, said fifth passage formed in said second
shroud.

14. The airfoil according to claim 1, wherein
said airfoil is part of a stationary vane.

15. A gas turbine vane, comprising:
a) a leading edge and a trailing edge;
b) first and second sidewalls;
c) inner and outer shrouds;
d) a cavity disposed between said first and
second sidewalls, said cavity having an inlet for
receiving a flow of cooling fluid directed to
said airfoil;
e) a plenum disposed between said cavity and
said trailing edge approximately midway between
said inner and outer shrouds, an opening formed
between said plenum and said cavity;
f) a first plurality of passages formed in
an array adjacent said trailing edge, said first
plurality of passages extending in a


12


substantially radially outward direction from
said plenum to said outer shroud; and
g) a second plurality of passages formed in
an array adjacent said trailing edge, said second
plurality of passages extending in a
substantially radially inward direction from said
plenum to said inner shroud.

16. The vane according to claim 15, further
comprising:
a) first and second manifolds formed in said
inner and outer shrouds, respectively;
b) said first plurality of passages
extending between said plenum and said first
manifold; and
c) said second plurality of passages
extending between said plenum and said second
manifold.

17. The vane according to claim 16, further
comprising:
a) means for discharging said cooling fluid
from said vane; and
b) third and fourth passages for placing
said first and second manifolds, respectively, in
flow communication with said cooling fluid
discharge means.


Description

Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.


CA 02258206 1998-12-14
WO 98/46860 PCT/US98/06039
CONFIGURATION OF COOLING CHANNELS FOR COOLING THE TRAILING EDGE OF GAS TURBINE
VANES
BACKGROUND OF THE INVENTION
The present invention relates to an airfoil for
use in a gas turbine, such as for a stationary vane. More
specifically, the present invention relates to an aifoil
having an improved cooling air flow path.
A gas turbine employs a plurality of stationary
vanes that are circumferentially arranged in rows in a
turbine section. Since such vanes axe exposed to the hot
gas discharging from the combustion section,~cooling of
these vanes is of the utmost importance. Typically,
cooling is accomplished by flowing cooling air through
cavities formed inside the vane airfoil.
According to one approach, cooling of the vane
airfoil is accomplished by incorporating one or more
tubular inserts into each of the airfoil cavities so that
passages surrounding the inserts are formed between the
inserts and the walls of the airfoil. The inserts have a
number of holes distributed around their periphery that
distribute the cooling air around these passages.
According to another approach, each airfoil
cavity includes a number of radially extending passages,
typically three, forming a serpentine array. Cooling air,
supplied to the vane outer shroud, enters the first passage
and flows radially inward until it reaches the vane inner
shroud. A first portion of the cooling air exits the vane
through the inner shroud and enters a cavity located
between adjacent rows of rotor discs. The cooling air in
the cavity serves to cool the faces of the discs. A second

CA 02258206 1998-12-14
WO 98/46860 PCT/US98/06039
2
portion of the cooling air reverses direction and flows
radially outward through the second passage until it
reaches the outer shroud, whereupon it changes direction
again and flows radially inward through the third passage.
Cooling of the trailing edge portion of the vane
is especially difficult because of the thinness of the
trailing edge portion. In traditional open loop cooling
systems, the cooling air is discharged from the vane
internal cavity into the hot gas flow path by axially
oriented passages in the trailing edge of the airfoil. In
closed loop systems, the trailing edge portion of the vane
airfoil may be cooled by directing the cooling air through
a channel that wraps around in the trailing edge in the
chord-wise direction. However, this approach results in a
thick trailing edge, which is aerodynamically undesirable,
and increased manufacturing complexity.
In another approach, the cooling air is directed
through span-wise radial holes extending between the inner
and outer shrouds, with the air flowing either radially
outward from the inner shroud to the outer shroud or
radially inward from the outer shroud to the inner shroud.
Unfortunately, this approach suffers from several
disadvantages. First, the cooling air can become
sufficiently heated by the time it reaches the ends of the
holes that its cooling effectiveness is inadequate, thereby
resulting in over-heating of the portion of the trailing
edge adjacent to the inner or outer shroud. Also, if the
diameter of the holes is relatively small, the length of
the holes results in an undesirably high pressure drop in
the cooling air. However, reducing the pressure drop by
increasing the diameter of the holes results in undesirably
thick trailing edges..
Span-wise radial holes are also difficult to
manufacture. If the airfoil is cast, the use of long,
small diameter span-wise radial holes can result it long,
unsupported, and therefore weak, casting cores. In
addition, such long cooling holes makes it difficult to

CA 02258206 1998-12-14
WO 98/46860 PCTlUS98/06039
3
maintain wall thickness tolerances, and results in a long
leaching time.
It is therefore desirable to provide a cooling
scheme for cooling the trailing edge portion of an airfoil
that overcomes the problems of previous approaches,
including the minimization of both the heat up of the
cooling fluid by the time it reaches the end of the cooling
path and the pressure drop experienced by the fluid.
SUMMARY OF THE INVENTION
Accordingly, it is the general object of the
current invention to provide a cooling scheme for cooling
the trailing edge portion of an airfoil that overcomes the
problems of previous approaches,, including the minimization
of both the heat up of the cooling fluid by the time it
reaches the end of the cooling path and the pressure drop
experienced by the fluid.
Briefly, this object, as well as other objects of
the current invention, is accomplished in an airfoil for a
gas turbine, comprising (i> a leading edge and a trailing
edge, (ii) first and second ends, the first end disposed
radially outward from the second end, (iii) first and
second side walls, (iv) a first passage formed between the
first and second sidewalls, the first passage having an
inlet for receiving a flow of a cooling fluid directed to
the airfoil, (v) a plenum disposed between the first and
second ends, the plenum in flow communication with the
first passage, (vi) a plurality of second passages in flow
communication with the plenum, the second passages
extending in a substantially radial direction from the
plenum toward the first end, (v) a plurality of third
passages in flow communication with the plenum, the third
passages extending in a substantially radial direction from
the plenum toward the second end.
In a preferred embodiment of the invention, the
plenum is disposed at about mid-height adjacent the
trailing edge of the airfoil.

CA 02258206 2006-03-29
66498-17
3a
In accordance with one aspect of this invention,
there is provided an airfoil for a turbomachine, comprising:
a) a leading edge and a trailing edge; b) first and second
ends, said first end disposed radially outward from said
second end; c) first and second side walls; d) a first
passage formed between said first and second sidewalls, said
first passage having an inlet for receiving a flow of a
cooling fluid directed to said airfoil; e) a plenum formed
between said side walls and disposed between said first and
second ends, said plenum in flow communication with said
first passage; f) a plurality of second passages in flow
communication with said plenum, said second passages
extending in a substantially radial direction from said
plenum toward said first end; and g) a plurality of third
passages in flow communication with said plenum, said third
passages extending in a substantially radial direction from
said plenum toward said second end.
In accordance with another aspect of this
invention, there is provided a gas turbine vane, comprising:
a) a leading edge and a trailing edge; b) first and second
sidewalk ; c) inner and outer shrouds; d) a cavity disposed
between said first and second sidewalls, said cavity having
an inlet for receiving a flow of cooling fluid directed to
said airfoil; e) a plenum disposed between said cavity and
said trailing edge approximately midway between said inner
and outer shrouds, an opening formed between said plenum and
said cavity; f) a first plurality of passages formed in an
array adjacent said trailing edge, said first plurality of
passages extending in a substantially radially outward
direction from said plenum to said outer shroud; and g) a

CA 02258206 2006-03-29
66498-17
3b
second plurality of passages formed in an array adjacent
said trailing edge, said second plurality of passages
extending in a substantially radially inward direction from
t said plenum to said inner shroud.

CA 02258206 1998-12-14
WO 98/46860 PCT/US98/06039
4
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a longitudinal cross-section through
a gas turbine vane of the current invention.
Figure 2 is a is transverse cross-section taken
through line II-II shown in Figure 1.
Figure 3 is a is transverse cross-section taken
through line III-III shown in Figure 1.
Figure 4 is an isometric view of a portion of the
trailing edge of the vane shown in Figure 1 in the vicinity
of the plenum.
DESCRIPTION OF THE PREFERRED EMBODIMENT
Referring to the drawings, there is shown in
Figures 1-4 a vane 1 having an airfoil according to the
current invention for use in the turbine section of a gas
turbine. The vane 1 is comprised of an airfoil 6 having an
inner shroud 2 on one end and an outer shroud 4 on the
other end. As shown best in Figure 2, the airfoil portion
6 of the vane 1 is formed by opposing side walls 9 and 11
that meet to form a leading edge 8 and a trailing edge 10.
The current invention concerns an apparatus for cooling the
airfoil 6, preferably the portion of the airfoil adjacent
the trailing edge 10.
The major portion of the airfoil 6 is hollow.
Transversely extending ribs 48, 50, and 52 divide the
hollow interior of the airfoil 6 into three cooling air
passages 32, 34, and 36. The first passage 32 is a cooling
air supply passage and is formed in the portion of the
airfoil 6 adjacent the leading edge 8. The second passage
34 is also a cooling supply passage but is formed in the
vicinity.of the trailing edge 6. A passage 17 in the inner
shroud 2 connects the passages 32 and 34. The third
passage 36 is formed in the mid-chord region of the airfoil
6 and forms a cooling air discharge passage.
Referring to Figure 1, a cooling fluid supply
pipe 13 is connected to the outer shroud 4. An opening 18
in the outer shroud 4 allows the supply pipe 13 to
communicate with a passage 16 formed within the outer

CA 02258206 1998-12-14
WO 98/46860 PCT/US98/06039
shroud. The outer shroud passage 16 is connected to
passages 32 and 34 in the airfoil 6.
As shown best in Figures 2 and 4, according to an
important aspect of the current invention, a cavity 42 is
5 formed between the side walls 9 and 11 that acts as a
plenum. The plenum 42 is preferably located at
approximately mid-height and adjacent the trailing edge 10
of the airfoil 6. An opening 40 in the rib.52 connects the
plenum 42 with the supply passage 34.
As shown best in Figures 1 and 3, a first array
of cooling fluid holes 38' extend radially outward from the
plenum 42 to a cooling fluid manifold 54 formed in the
outer shroud 4, with the inlets to the holes being at the
plenum and the outlets being at the manifold. As shown in
Figure 3, a passage 58 is formed in the outer shroud 4 that
extends generally perpendicularly to the radial direction.
The passage 58 extends from the manifold 54 around the
portion of the airfoil 6 projecting-into the outer shroud.
Openings 46 and 47 are formed in the portions of the side
walls 9 and 11, respectively, that extend into the outer
shroud 4. The openings 46 and 47 allow the passage 58 to
communicate with the discharge passage 36. As shown in
Figure 1, an outlet 30 is formed in the discharge passage
36 and is connected to a return pipe 14.
As shown best in Figures 1, 2 and 4, a second
array of cooling fluid holes 38", which are preferably
radially aligned with the cooling fluid holes 38', extend
radially inward from the plenum 42 to a cooling fluid
manifold 56 formed in the inner shroud 2, with the inlets
to the holes being at the plenum and the outlets being at
the manifold. A passage (not shown), similar to passage 58
in the outer shroud 4, is formed in the inner shroud 2 that
extends from the manifold 56 around the portion of the
airfoil 6 projecting into the inner shroud. Openings 44,
one of which is shown in Figure 1, which are similar to
openings 46 and 47 at the outer shroud 4, are formed in the
portions of the side walls 9 and 11, respectively, that

CA 02258206 1998-12-14
WO 98/46860 PCT/US98/06039
6
extend into the inner shroud 2. The openings 44 allow the
passage in the inner shroud 2 to communicate with the
discharge passage 36.
It should be understood that the inner and outer
shrouds may contain cooling passages, in addition to those
connecting the trailing edge cooling fluid manifolds 54 and
56 to the discharge passage 36, that aid in the cooling of
the shrouds themselves. However, such shroud cooling is
not part of the current invention, which concerns the
cooling of the airfoil 6 and, preferably, the portion of
the airfoil adjacent the trailing edge 10.
In operation, cooling fluid, which in the
preferred embodiment is compressed air 20, typically bled
from the compressor section of the gas turbine, is directed
to the vane outer shroud 4 by the supply pipe 13, as shown
in Figure 1. According to a preferred embodiment of the
invention, the vane 1 has cooling passages that are part of
a closed loop cooling air system. Thus, essentially all of
the cooling air supplied to the vane 1 is returned to the
cooling system.
Upon flowing through the opening 1B and entering
the passage 16 in the outer shroud 4, the cooling air 20 is
divided into two streams 22 and 24. The first cooling air
stream 22 flows radially inward through the trailing edge
supply passage 34 to the plenum 42 and, in so doing, cools
a portion of the side walls 9 and 11 of the airfoil 6.
The second cooling air stream 24 flows radially
inward through the leading edge supply passage 32 and cools
the leading edge a portion of the airfoil 6. The passage
17 in the inner shroud 2 then directs the cooling air 24
from the passage 32 to the passage 34, where it flows
radially outward (that is, toward the outer shroud 4) to
the plenum 42. In the plenum 42, the cooling air streams
22 and 24 combine and are then divided into numerous small
streams by the trailing edge cooling holes 38. As shown
best in Figures 2 and 4, the plenum is tapered as it
extends in the axial direction toward the trailing edge 10

CA 02258206 1998-12-14
WO 98/46860 PCT/US98/06039
7
of the airfoil 6. Such tapering provides the area
reduction necessary for uniform flow distribution among the
cooling holes 38.
A portion 28 of the combined flow of cooling air
22 and 24, flows radially outward (that is, toward the
outer shroud 4) from the plenum 42 through the holes 38' to
the manifold 54, thereby providing vigorous cooling of the
approximately upper half portion of the airfoil 6 adjacent
the trailing edge 10 that is located above the plenum 42.
In the manifold 54, the individual streams of cooling air
28 are collected and are then directed by passage 58 to the
openings 46 and 47, as shown in Figure 3. From the
openings 46 and 47, the cooling air 28 enters the discharge
passage 36 and flows radially outward to the exhaust pipe
14, as shown in Figure 1.
Similarly, a portion 26 of the combined flow of
cooling air 22 and 24, flows radially inward from the
plenum 42 through the holes 38" to the manifold 56, thereby
providing vigorous cooling of the approximately lower half
portion of the airfoil 6 adjacent the trailing edge l0
below the plenum 42. In the manifold 56, the individual
streams of cooling air 26 are collected and are then
directed by the inner shroud passage to the openings 44, as
discussed above with respect to the outer shroud 4. From
the openings 44, the cooling air 26 enters the discharge
passage 36 and flows radially outward to the exhaust pipe
14 and, in so doing, cools the mid-chord portion of the
side walls 9 and 11 of the airfoil 6. In the preferred
embodiment of the invention, the exhaust pipe 14 directs
the cooling air 29 to a cooler for recycling back to the
turbine.
The present~invention has numerous advantages
over traditional airfoil cooling schemes. First, since the
length of the cooling air passages 38 is effectively cut in
half, compared to span-wise holes that extend from the
inner shroud to the outer shroud, there is less chance of
overheating the coolant, which may be air or steam, for

CA 02258206 1998-12-14
WO 98/46860 PCT/US98/06039
8
example, by the time it reaches a shroud. Also, the
pressure drop through the passages 38 is reduced, thereby
allowing the use of holes 38 of minimum diameter. Small
diameter holes permit the use of a thin trailing edge 10,
which has aerodynamic advantages. The airfoil 6 is also
easier to manufacture since long runs of cooling holes are
avoided.
Although the current invention has been discussed
in connection with the.airfoil for a stationary vane in a
gas turbine, the invention is also applicable to other
types of components. In addition, although the invention
has been discussed with reference to a closed loop cooling
system utilizing compressed air, the invention is also
applicable to more conventional open loop systems as well
as to systems using other types of cooling fluids, such as
steam. Thus, the present invention may be embodied in
other specific forms without departing from the spirit or
essential attributes thereof and, accordingly, reference
should be made to the appended claims, rather than to the
foregoing specification, as indicating the scope of the
invention.

Dessin représentatif
Une figure unique qui représente un dessin illustrant l'invention.
États administratifs

Pour une meilleure compréhension de l'état de la demande ou brevet qui figure sur cette page, la rubrique Mise en garde , et les descriptions de Brevet , États administratifs , Taxes périodiques et Historique des paiements devraient être consultées.

États administratifs

Titre Date
Date de délivrance prévu 2006-06-27
(86) Date de dépôt PCT 1998-03-25
(87) Date de publication PCT 1998-10-22
(85) Entrée nationale 1998-12-14
Requête d'examen 2003-03-12
(45) Délivré 2006-06-27
Réputé périmé 2018-03-26

Historique d'abandonnement

Il n'y a pas d'historique d'abandonnement

Historique des paiements

Type de taxes Anniversaire Échéance Montant payé Date payée
Le dépôt d'une demande de brevet 300,00 $ 1998-12-14
Enregistrement de documents 100,00 $ 1999-05-05
Enregistrement de documents 100,00 $ 1999-05-05
Enregistrement de documents 100,00 $ 1999-05-05
Enregistrement de documents 100,00 $ 1999-05-05
Enregistrement de documents 100,00 $ 1999-05-05
Enregistrement de documents 100,00 $ 1999-05-05
Taxe de maintien en état - Demande - nouvelle loi 2 2000-03-27 100,00 $ 2000-02-18
Taxe de maintien en état - Demande - nouvelle loi 3 2001-03-26 100,00 $ 2001-02-21
Taxe de maintien en état - Demande - nouvelle loi 4 2002-03-25 100,00 $ 2002-02-25
Taxe de maintien en état - Demande - nouvelle loi 5 2003-03-25 150,00 $ 2003-02-17
Requête d'examen 400,00 $ 2003-03-12
Taxe de maintien en état - Demande - nouvelle loi 6 2004-03-25 200,00 $ 2004-02-12
Taxe de maintien en état - Demande - nouvelle loi 7 2005-03-25 200,00 $ 2005-02-11
Taxe de maintien en état - Demande - nouvelle loi 8 2006-03-27 200,00 $ 2006-02-10
Expiré 2019 - Modifications après acceptation 400,00 $ 2006-03-29
Taxe finale 300,00 $ 2006-04-04
Taxe de maintien en état - brevet - nouvelle loi 9 2007-03-26 200,00 $ 2007-02-14
Taxe de maintien en état - brevet - nouvelle loi 10 2008-03-25 250,00 $ 2008-02-15
Taxe de maintien en état - brevet - nouvelle loi 11 2009-03-25 250,00 $ 2009-02-04
Taxe de maintien en état - brevet - nouvelle loi 12 2010-03-25 250,00 $ 2010-02-10
Taxe de maintien en état - brevet - nouvelle loi 13 2011-03-25 250,00 $ 2011-02-08
Taxe de maintien en état - brevet - nouvelle loi 14 2012-03-26 250,00 $ 2012-02-08
Taxe de maintien en état - brevet - nouvelle loi 15 2013-03-25 450,00 $ 2013-02-08
Taxe de maintien en état - brevet - nouvelle loi 16 2014-03-25 450,00 $ 2014-02-07
Taxe de maintien en état - brevet - nouvelle loi 17 2015-03-25 450,00 $ 2015-02-04
Taxe de maintien en état - brevet - nouvelle loi 18 2016-03-29 450,00 $ 2016-02-08
Titulaires au dossier

Les titulaires actuels et antérieures au dossier sont affichés en ordre alphabétique.

Titulaires actuels au dossier
WESTINGHOUSE ELECTRIC CORPORATION
Titulaires antérieures au dossier
HULTGREN, KENT G.
NORDLUND, RAYMOND S.
NORTH, WILLIAM E.
SCOTT, ROBERT K.
SINNOT, ZACHARY
WARD, STEVEN D.
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Description du
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Date
(yyyy-mm-dd) 
Nombre de pages   Taille de l'image (Ko) 
Dessins représentatifs 1999-03-09 1 13
Page couverture 1999-03-09 2 68
Abrégé 1998-12-14 1 56
Dessins 1998-12-14 3 75
Description 1998-12-14 8 375
Revendications 1998-12-14 4 127
Page couverture 2006-05-31 1 50
Dessins représentatifs 2005-11-15 1 13
Description 2006-03-29 10 427
Correspondance 1999-02-16 1 33
PCT 1998-12-14 4 119
Cession 1998-12-14 3 99
PCT 1999-01-27 1 50
PCT 1999-01-27 1 51
Cession 1999-05-05 20 903
Correspondance 1999-05-05 9 514
Cession 1998-12-14 7 235
PCT 2000-06-14 1 61
Poursuite-Amendment 2003-03-12 1 44
Correspondance 2006-04-04 1 40
Poursuite-Amendment 2006-03-29 3 90
Poursuite-Amendment 2006-04-05 1 16