CA2206980C - Gas turbine blade retention - Google Patents
Gas turbine blade retention Download PDFInfo
- Publication number
- CA2206980C CA2206980C CA 2206980 CA2206980A CA2206980C CA 2206980 C CA2206980 C CA 2206980C CA 2206980 CA2206980 CA 2206980 CA 2206980 A CA2206980 A CA 2206980A CA 2206980 C CA2206980 C CA 2206980C
- Authority
- CA
- Canada
- Prior art keywords
- disc
- gas turbine
- blade
- blades
- strip
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Lifetime
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- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Gas turbine blades (14) are slid axially into disc (12) with a retention tang (22) on each blade abutting the disc. An axially extending space (18) between blade platforms (26) and the disc receives elongated strip (30). With prebent end (34) resiliently held against the disc the opposite end (36) is bent radially outward against the roots of adjacent blades. A bow (38) biases the blades outwardly, deterring vibration.
Description
Gas Turbine Blade Retention Technical Field The invention relates to retention of gas turbine blades on a disc, and in particular to a clip which retains, dampens and seals the arrangement.
Background of the Invention It is conventional to secure gas turbine blades to the disc of a gas turbine with dove tail fir tree grooves in the disc. A fir tree root on the blade engages these grooves.
Precise location of the blade in the radially outward direction is established by precise locations on the two fir trees. Therefore it is designed to bear against the support surface with the blade in its radially outermost position. Inboard clearances are of course required to permit insertion of the blade.
In such an arrangement some means are required to axially retain the blade at its desired position.
At high rpm's centrifugal force will establish the blade in its outer position.
However it is required that the blade have substantially the same position at balancing speed (1000 rpm) and also at tip grinding speed (100 rpm).
Sealing is required to deter gas passage from the gas path upstream of the blade, between blade platforms, to the space under the blade at the downstream side thereof.
Damping of the blades is also a benefit to reduce vibratory stresses of blades during operation.
Summary of the Invention In accordance with another aspect of the invention, there is provided a disc assembly for a gas turbine engine comprising: a gas turbine disc having a first side, a second side, an axis and a periphery; axially extending dove tail recesses in the periphery of said disc with dead load material between said recesses; a plurality of gas turbine blades, each blade having (a) a root conforming to and located within one of said recesses, (b) a retention tang on one side of said blade, said tang abutting said first side of said disc and (c) blade platforms extending circumferentially toward blade platforms of adjacent blades and terminating in closely spaced relation to said blade platforms of adjacent blades; spaces between said disc and said blade platforms, said spaces extending axially between adjacent blade platforms; and elongated retention strips located in said spaces, each of said strips having a first end engaging adjacent blades on said one side of each of said adjacent blades, each of said retention strips having a second end resiliently engaging said dead load material on said second side of said disc to axially bias said retention tangs of said adjacent blades against said first side of said disc to axially locate said blades.
In accordance with an aspect of the invention, there is provided a method of assembling a disc assembly for a gas turbine engine comprising the steps of:
sliding a first gas turbine blade axially in relation to the axis of a gas turbine disc, said disc having a first side and a second side, from said first side of said disc into engagement with said disc and against a stop; sliding a second gas turbine blade axially in relation to the axis of said disc from said first side of said disc into engagement with said disc and against another stop, adjacent said first gas turbine blade; inserting an elongated retention strip, said strip having a first end and a second end, axially in relation to the axis of said disc from said second side of said disc, between said disc and both said first and second blades, to bring a portion of said first end of said strip into resilient contact with said second side of said disc; applying a force to said strip to increase the resilient contact between said portion of said first end of said strip and said second side of said disc; bending said second end of said strip into contact with said first and second gas turbine blades on said first side of said disc while maintaining said applied force; and releasing said applied force leaving said strip in resilient contact with said disc and said blades.
Brief Description of the Drawings Figure 1 is a view of the disc, the gas turbine blades and the blade platform looking radially inward from outside the gas turbine stage;
Figure 2 is a view circumferentially taken through section 2-2 of Figure 1;
Figure 3 is an axial view of Figure 2 looking upstream;
Figure 4 is an axial view of Figure 2 looking downstream;
Figure 5 is a side view of the retention strip before insertion; and Figure 6 is a top view of the retention strip before insertion.
Description of the Preferred Embodiment Referring to Figure 1 the gas turbine blade retention arrangement 10 includes a gas turbine disc 12 and a plurality of gas turbine blades 14 located in gas flow 15.
Referring also to Figures 2, 3 and 4 it can be seen that there are a plurality of dove tail recesses 16 located around the periphery of the disc. These leave dead load material 18 between the recesses. Each gas turbine blade has a root 20 conforming to the dove tail recesses 16. Each root conforms to and is located in one of the recesses.
A
retention tang 22 is located on one side of each blade abutting the first side 24 of the 2a Vi'O 96118803 PCT/CA95100683 disc. The blades are inserted by sliding them into the recesses from this side unfit tang 22 stops movement of the blade.
Circumferentially extending platforms 26 are located on each blade. Axially extending space 28 is located between the disc and adjacent blade platforms.
An elongated retention strip 30 is located in this space. It is inserted by sliding it in from the second side 32 of the rim. The resilient tab 34 is formed on the retention strip prior to installation of the strip. The strip is inserted until resilient contact is made with surface 32. Additional force is then applied to further increase resilient contact.
While holding the strip in this location, tab 36 at the first end is bent upwardly or outwardly in contact with adjacent turbine blades. When the force is released resilient contact between resilient tab 34 in the face continues thereby maintaining a constant force on the gas turbine blades operating against the force applied on tab 22.
Only the extreme end 35 of tab 34 is in contact with the disc.
Figure 5 and 6 show the retention strip 30 in its formed condition prior to installation. End 34 which will be in resilient contact with the disc has already been bent. It is also noted that there is a bow 38 in the strip. Referring to Figure 2 this creates a force resiliently biasing the blades radially outward at location 40. This urges the blades outwardly maintaining them in position during tip grinding of the blades at 100 rpm approximately, and during balancing of the gas turbine section at about 1000 rpm.
This force against the blades combined with the resilient retention of the strip also dampens vibration as a blade to blade damper. The retention strip also tends to restrict flow through gap 42 where flow shown by arrow 44 in Figure 2 would otherwise pass form zone 46 in the gas passage upstream of the blade, through the gaps 42 to area 48 which is the space under the blade and downstream thereof.
Figure 6 is a top view of the retention strip 30 also showing the tab 36 in its unbent condition.
The invention retains the turbine blades in the turbine disc and also provides a seal where the blade is secured to the disc. It acts as a blade to blade damper, and also generates a radial load to aid in balancing and tip grinding.
Background of the Invention It is conventional to secure gas turbine blades to the disc of a gas turbine with dove tail fir tree grooves in the disc. A fir tree root on the blade engages these grooves.
Precise location of the blade in the radially outward direction is established by precise locations on the two fir trees. Therefore it is designed to bear against the support surface with the blade in its radially outermost position. Inboard clearances are of course required to permit insertion of the blade.
In such an arrangement some means are required to axially retain the blade at its desired position.
At high rpm's centrifugal force will establish the blade in its outer position.
However it is required that the blade have substantially the same position at balancing speed (1000 rpm) and also at tip grinding speed (100 rpm).
Sealing is required to deter gas passage from the gas path upstream of the blade, between blade platforms, to the space under the blade at the downstream side thereof.
Damping of the blades is also a benefit to reduce vibratory stresses of blades during operation.
Summary of the Invention In accordance with another aspect of the invention, there is provided a disc assembly for a gas turbine engine comprising: a gas turbine disc having a first side, a second side, an axis and a periphery; axially extending dove tail recesses in the periphery of said disc with dead load material between said recesses; a plurality of gas turbine blades, each blade having (a) a root conforming to and located within one of said recesses, (b) a retention tang on one side of said blade, said tang abutting said first side of said disc and (c) blade platforms extending circumferentially toward blade platforms of adjacent blades and terminating in closely spaced relation to said blade platforms of adjacent blades; spaces between said disc and said blade platforms, said spaces extending axially between adjacent blade platforms; and elongated retention strips located in said spaces, each of said strips having a first end engaging adjacent blades on said one side of each of said adjacent blades, each of said retention strips having a second end resiliently engaging said dead load material on said second side of said disc to axially bias said retention tangs of said adjacent blades against said first side of said disc to axially locate said blades.
In accordance with an aspect of the invention, there is provided a method of assembling a disc assembly for a gas turbine engine comprising the steps of:
sliding a first gas turbine blade axially in relation to the axis of a gas turbine disc, said disc having a first side and a second side, from said first side of said disc into engagement with said disc and against a stop; sliding a second gas turbine blade axially in relation to the axis of said disc from said first side of said disc into engagement with said disc and against another stop, adjacent said first gas turbine blade; inserting an elongated retention strip, said strip having a first end and a second end, axially in relation to the axis of said disc from said second side of said disc, between said disc and both said first and second blades, to bring a portion of said first end of said strip into resilient contact with said second side of said disc; applying a force to said strip to increase the resilient contact between said portion of said first end of said strip and said second side of said disc; bending said second end of said strip into contact with said first and second gas turbine blades on said first side of said disc while maintaining said applied force; and releasing said applied force leaving said strip in resilient contact with said disc and said blades.
Brief Description of the Drawings Figure 1 is a view of the disc, the gas turbine blades and the blade platform looking radially inward from outside the gas turbine stage;
Figure 2 is a view circumferentially taken through section 2-2 of Figure 1;
Figure 3 is an axial view of Figure 2 looking upstream;
Figure 4 is an axial view of Figure 2 looking downstream;
Figure 5 is a side view of the retention strip before insertion; and Figure 6 is a top view of the retention strip before insertion.
Description of the Preferred Embodiment Referring to Figure 1 the gas turbine blade retention arrangement 10 includes a gas turbine disc 12 and a plurality of gas turbine blades 14 located in gas flow 15.
Referring also to Figures 2, 3 and 4 it can be seen that there are a plurality of dove tail recesses 16 located around the periphery of the disc. These leave dead load material 18 between the recesses. Each gas turbine blade has a root 20 conforming to the dove tail recesses 16. Each root conforms to and is located in one of the recesses.
A
retention tang 22 is located on one side of each blade abutting the first side 24 of the 2a Vi'O 96118803 PCT/CA95100683 disc. The blades are inserted by sliding them into the recesses from this side unfit tang 22 stops movement of the blade.
Circumferentially extending platforms 26 are located on each blade. Axially extending space 28 is located between the disc and adjacent blade platforms.
An elongated retention strip 30 is located in this space. It is inserted by sliding it in from the second side 32 of the rim. The resilient tab 34 is formed on the retention strip prior to installation of the strip. The strip is inserted until resilient contact is made with surface 32. Additional force is then applied to further increase resilient contact.
While holding the strip in this location, tab 36 at the first end is bent upwardly or outwardly in contact with adjacent turbine blades. When the force is released resilient contact between resilient tab 34 in the face continues thereby maintaining a constant force on the gas turbine blades operating against the force applied on tab 22.
Only the extreme end 35 of tab 34 is in contact with the disc.
Figure 5 and 6 show the retention strip 30 in its formed condition prior to installation. End 34 which will be in resilient contact with the disc has already been bent. It is also noted that there is a bow 38 in the strip. Referring to Figure 2 this creates a force resiliently biasing the blades radially outward at location 40. This urges the blades outwardly maintaining them in position during tip grinding of the blades at 100 rpm approximately, and during balancing of the gas turbine section at about 1000 rpm.
This force against the blades combined with the resilient retention of the strip also dampens vibration as a blade to blade damper. The retention strip also tends to restrict flow through gap 42 where flow shown by arrow 44 in Figure 2 would otherwise pass form zone 46 in the gas passage upstream of the blade, through the gaps 42 to area 48 which is the space under the blade and downstream thereof.
Figure 6 is a top view of the retention strip 30 also showing the tab 36 in its unbent condition.
The invention retains the turbine blades in the turbine disc and also provides a seal where the blade is secured to the disc. It acts as a blade to blade damper, and also generates a radial load to aid in balancing and tip grinding.
S~J~STtTI~TE S~-1~~T
Claims (10)
1. A disc assembly for a gas turbine engine comprising:
a gas turbine disc having a first side, a second side, an axis and a periphery;
axially extending dove tail recesses in the periphery of said disc with dead load material between said recesses;
a plurality of gas turbine blades, each blade having (a) a root conforming to and located within one of said recesses, (b) a retention tang on one side of said blade, said tang abutting said first side of said disc and (c) blade platforms extending circumferentially toward blade platforms of adjacent blades and terminating in closely spaced relation to said blade platforms of adjacent blades;
spaces between said disc and said blade platforms, said spaces extending axially between adjacent blade platforms; and elongated retention strips located in said spaces, each of said strips having a first end engaging adjacent blades on said one side of each of said adjacent blades, each of said retention strips having a second end resiliently engaging said dead load material on said second side of said disc to axially bias said retention tangs of said adjacent blades against said first side of said disc to axially locate said blades.
a gas turbine disc having a first side, a second side, an axis and a periphery;
axially extending dove tail recesses in the periphery of said disc with dead load material between said recesses;
a plurality of gas turbine blades, each blade having (a) a root conforming to and located within one of said recesses, (b) a retention tang on one side of said blade, said tang abutting said first side of said disc and (c) blade platforms extending circumferentially toward blade platforms of adjacent blades and terminating in closely spaced relation to said blade platforms of adjacent blades;
spaces between said disc and said blade platforms, said spaces extending axially between adjacent blade platforms; and elongated retention strips located in said spaces, each of said strips having a first end engaging adjacent blades on said one side of each of said adjacent blades, each of said retention strips having a second end resiliently engaging said dead load material on said second side of said disc to axially bias said retention tangs of said adjacent blades against said first side of said disc to axially locate said blades.
2. A disc assembly for a gas turbine engine as claimed in claim 1 wherein said second end of said strip has an extreme end and wherein only said extreme end of said second end is in contact with said second side of said disc.
3. A disc assembly for a gas turbine engine as claimed in claim 1, said gas turbine engine being designed to cause a gas flow in a downstream direction toward said blades of said disc assembly, wherein said first face of said disc is downstream of said second face of said disc in said gas turbine engine.
4. A disc assembly for a gas turbine engine as claimed in claim 1 wherein said blade root has a fir tree configuration.
5. A disc assembly for a gas turbine engine comprising:
a gas turbine disc having a first side, a second side, an axis and a periphery;
axially extending dove tail recesses in the periphery of said disc with dead load material between said recesses;
a plurality of gas turbine blades, each blade having (a) a root conforming to and located within one of said recesses, (b) a retention tang on one side of said blade, said tang abutting said first side of said disc and (c) blade platforms extending circumferentially toward blade platforms of adjacent blades and terminating in closely spaced relation to said blade platforms of adjacent blades;
spaces between said disc and said blade platforms, said spaces extending axially between adjacent blade platforms; and elongated retention strips located in said spaces, each of said strips having a first end engaging adjacent blades on said one side of each of said adjacent blades, each of said retention strips having a second end resiliently engaging said dead load material on said second side of said disc to axially bias said retention tangs of said adjacent blades against said first side of said disc, each of said retention strips further resiliently biasing said adjacent blade platforms radially outwardly from said disc.
a gas turbine disc having a first side, a second side, an axis and a periphery;
axially extending dove tail recesses in the periphery of said disc with dead load material between said recesses;
a plurality of gas turbine blades, each blade having (a) a root conforming to and located within one of said recesses, (b) a retention tang on one side of said blade, said tang abutting said first side of said disc and (c) blade platforms extending circumferentially toward blade platforms of adjacent blades and terminating in closely spaced relation to said blade platforms of adjacent blades;
spaces between said disc and said blade platforms, said spaces extending axially between adjacent blade platforms; and elongated retention strips located in said spaces, each of said strips having a first end engaging adjacent blades on said one side of each of said adjacent blades, each of said retention strips having a second end resiliently engaging said dead load material on said second side of said disc to axially bias said retention tangs of said adjacent blades against said first side of said disc, each of said retention strips further resiliently biasing said adjacent blade platforms radially outwardly from said disc.
6. A disc assembly for a gas turbine engine as claimed in claim 5 wherein said second end of said strip has an extreme end and wherein only said extreme end of said second end is in contact with said second side of said disc.
7. A disc assembly for a gas turbine engine as claimed in claim 5, said gas turbine engine being designed to cause a gas flow in a downstream direction toward said blades of said disc assembly, wherein said first face of said disc is downstream of said second face in said gas turbine engine.
8. A disc assembly for a gas turbine engine as claimed in claim 5 wherein said blade root has a fir tree configuration.
9. A method of assembling a disc assembly for a gas turbine engine comprising the steps of:
sliding a first gas turbine blade axially in relation to the axis of a gas turbine disc, said disc having a first side and a second side, from said first side of said disc into engagement with said disc and against a stop;
sliding a second gas turbine blade axially in relation to the axis of said disc from said first side of said disc into engagement with said disc and against another stop, adjacent said first gas turbine blade;
inserting an elongated retention strip, said strip having a first end and a second end, axially in relation to the axis of said disc from said second side of said disc, between said disc and both said first and second blades, to bring a portion of said first end of said strip into resilient contact with said second side of said disc;
applying a force to said strip to increase the resilient contact between said portion of said first end of said strip and said second side of said disc;
bending said second end of said strip into contact with said first and second gas turbine blades on said first side of said disc while maintaining said applied force;
and releasing said applied force leaving said strip in resilient contact with said disc and said blades.
sliding a first gas turbine blade axially in relation to the axis of a gas turbine disc, said disc having a first side and a second side, from said first side of said disc into engagement with said disc and against a stop;
sliding a second gas turbine blade axially in relation to the axis of said disc from said first side of said disc into engagement with said disc and against another stop, adjacent said first gas turbine blade;
inserting an elongated retention strip, said strip having a first end and a second end, axially in relation to the axis of said disc from said second side of said disc, between said disc and both said first and second blades, to bring a portion of said first end of said strip into resilient contact with said second side of said disc;
applying a force to said strip to increase the resilient contact between said portion of said first end of said strip and said second side of said disc;
bending said second end of said strip into contact with said first and second gas turbine blades on said first side of said disc while maintaining said applied force;
and releasing said applied force leaving said strip in resilient contact with said disc and said blades.
10. The method as claimed in claim 9, further including the step prior to inserting said retention strip of:
introducing a bow to said retention strip for biasing said adjacent blades radially outwardly of said disc once said strip is inserted.
introducing a bow to said retention strip for biasing said adjacent blades radially outwardly of said disc once said strip is inserted.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US08/356,094 US5518369A (en) | 1994-12-15 | 1994-12-15 | Gas turbine blade retention |
| US08/356,094 | 1994-12-15 | ||
| PCT/CA1995/000683 WO1996018803A1 (en) | 1994-12-15 | 1995-12-07 | Gas turbine blade retention |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CA2206980A1 CA2206980A1 (en) | 1996-06-20 |
| CA2206980C true CA2206980C (en) | 2000-04-25 |
Family
ID=29550751
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA 2206980 Expired - Lifetime CA2206980C (en) | 1994-12-15 | 1995-12-07 | Gas turbine blade retention |
Country Status (1)
| Country | Link |
|---|---|
| CA (1) | CA2206980C (en) |
-
1995
- 1995-12-07 CA CA 2206980 patent/CA2206980C/en not_active Expired - Lifetime
Also Published As
| Publication number | Publication date |
|---|---|
| CA2206980A1 (en) | 1996-06-20 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EEER | Examination request | ||
| MKEX | Expiry |
Effective date: 20151207 |