JPS6125883B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a nozzle guide vane assembly for a gas turbine engine. The nozzle guide vane assembly is one of the most difficult to design in a gas turbine engine. This is because the nozzle guide vanes must withstand the highest temperatures of the engine and perform an efficient aerodynamic function on the hot gases flowing from the combustion chamber. Traditionally, the common practice has been to make the nozzle guide vane assembly an annular array of individual guide vanes;
Each guide vane has an airfoil and radially inner and outer platforms integral with the airfoil. This is not necessarily the best approach. Airfoil vanes have different required characteristics from the platform; in any case, the platform performs an aerodynamic function and a load-bearing function, and the shape of the platform has the characteristics necessary for these two functions. Usually it is a compromise of terms. For aerodynamic function, it must have a surface shape suitable for the desired flow of airflow, and for load-bearing function, it must have a surface shape suitable for the desired flow of airflow.
The material must have the required strength and rigidity, but if a relatively thick material is used to provide the required strength and rigidity, it will be difficult to mold the surface into the desired shape.

The present invention allows the airfoil vanes to be made separate from the platform, separating the load-bearing and aerodynamic functions of the platform and improving the design of the guide vane assembly. It is.

The nozzle guide vane assembly for a gas turbine engine of the present invention includes a series of circumferentially spaced airfoil vanes, a retaining structure that retains the airfoil vanes, and a retaining structure that is directly or indirectly connected to the retaining structure. a protrusion on each airfoil vane adapted to engage and radially retain the airfoil vane, and radially inner and outer platform members provided separately from the airfoil vane; Each platform member is comprised of a relatively thick and relatively stiff support skin and a relatively thin and relatively less stiff inner skin;
The support skin has at least one airfoil opening through which the airfoil blade passes, the airfoil opening supporting the airfoil blade against torsional, circumferential and axial loads; , the airfoil having at least one airfoil opening therethrough, the inner skin not substantially contributing to the support of the airfoil and forming a boundary for gas flow through the nozzle guide vane assembly. , the vanes are radially inflatably slidable within the airfoil openings of the support skin and the inner skin, and a sealing means associated with the inner skin is arranged between the inner skin and the airfoil vanes. is sealed.

In the present invention, the platform is constructed separately from the airfoil vanes and with a relatively thick support skin for performing exclusively load-bearing functions and a relatively thin inner skin for performing exclusively aerodynamic functions. As a result, each skin can be designed to perform the best possible function for each skin. There is no need to compromise the design to achieve these goals at the same time.

Embodiments of the present invention will be described in detail below with reference to the drawings.

FIG. 1 shows a gas turbine engine consisting of a casing 10, a compressor 11, a combustion device 12 and a turbine 13 arranged therein in the usual order, as well as a nozzle 14 formed by the casing 10. . A nozzle guide vane assembly 15 is provided to guide the hot gases exiting the combustion chamber 12 to the turbine 13.

FIG. 2 shows an enlarged cross-sectional view of nozzle guide vane assembly 15. FIG. The assembly 15 includes a plurality of airfoil blades 16 arranged circumferentially, each airfoil blade extending generally in a radial direction, and adjacent airfoil blades having a constant angular spacing. The radially inner and outer ends of each airfoil vane are provided with projecting pads 17 and 18, respectively. Inner pad 17 engages annular support member 19. The annular support member is carried by a frustoconical mounting flange 20 that forms part of the engine's fixed structure. Similarly, the outer pad 18 has a flange 2 projecting from the casing 22.
1 is engaged. The casing 22 is also
It is part of the engine's fixed structure. Padded 1
7, 18 and members 19, 21, the airfoil blade 16 is held radially stationary, but the flexibility between the member 19 and the flange 20 allows the airfoil blade to It is expandable relative to the fixed structure.

The airfoil blade 16 is provided with a land 23 at the outer end,
The land 23 is a slightly raised portion of the surface of the airfoil vane, completely circling the tip of the airfoil. The lands 23 fit into correspondingly shaped openings 24 in the support skin 25. The dimensions of the opening 24 are almost the same as those of the land 23, but there is no interference. In this way, the ends of the airfoil vanes are slidable in the radial direction relative to the skin 25, but are positioned in other directions.

The skin 25 extends rearwardly from an engagement portion in an annular groove 26 of a flange 27 projecting from the casing 22 . Immediately downstream of the airfoil 16, the support skin 25 is provided with an integral flange 28 which is secured to an end flange 30 formed in the casing 22 by a series of bolts 29. The skin 25 extends further downstream to the second
It terminates in a flange 31. The flange 31 projects radially outward and engages a structure not shown. An annular projection 32 is also formed at the downstream end of the skin 25, the purpose of which will be described later.

The skin 25 connects the airfoil vane 16 to a fixed structure,
In particular, it is strong enough to support all forces acting on the casing 22 in the plane of the skin 25 itself. Beneath the skin 25 is a thin inner skin 33. This skin also has an airfoil-shaped opening through which the airfoil vane 16 passes in a tight fit, but with no interference, and the underside of the skin has a smooth aerodynamic shape. None, forming the outer boundary of the gas passage through the guide vane assembly 15. To support the skin 33, a series of hooks 34 are provided at its front end, which together with the front end of the skin 25 fit into the groove 26. A second series of hooks 35 fit into corresponding grooves 36 formed in the inner surface of the skin 25 at the base of the flange 28. Finally, the downstream end of the skin 33 is provided with a third series of hooks 37 which engage the projections 32 of the skin 25.

Inner skin 33 extends beyond the ends of airfoil vane 16 and provides an outer boundary for flow through first stage turbine rotor 38 .

If the fit between the land 23 and the opening 24 in the skin 25 is not very tight, the leakage of high pressure cooling air from the space defined by the casing 22 and the skin 25 or the high temperature of the gas stream may occur. There is a risk of gas leaking into this space. In order to minimize this leakage, the skin 25 includes:
A raised lip 39 (FIG. 5) is provided which extends around the opening 24 and defines a rebate 40 at the inner edge. A sealing material is engaged on the inside of the rebate 40. The encapsulant may be any suitable elastic and heat resistant material, but in this embodiment silica string 4 is used.
1a is used. This string can be a tight fit with the rebate 40 to provide an effective gas seal as well as reduce the possibility of the airfoil vanes becoming scratched within the aperture 24.

3 and 4 show developed views of the upper surfaces of the skins 33 and 25, respectively, and the skin 33 has lines 41 and 25, respectively.
42 into several segments. The segments abut each other to form a complete ring. The abutment surface between each segment connects to a surface around one of the airfoil vanes 16. In this way, the length of the abutment surface of the skin (which requires sealing) is reduced by utilizing the openings required to pass the airfoil vanes through. As can be seen from FIG. 4, the skin 25 is provided with a large number of small holes 57. This small hole is a small hole for impingement cooling, and as is well known in the art, a high-pressure cooling fluid flows through this small hole and collides with the skin 33 to efficiently cool the skin 33. I can do it.

A radially outer platform made of skins 33, 25 has been described above. The inner platform is constructed in much the same way. In this case, the thick support skin 43 is provided with an opening 44 through which a land 45 similar to the land 23 of the airfoil 16 passes. Although the skin 43 is different in shape from the skin 25, it serves a similar function in supporting the load from the airfoil vane 16, and the airfoil vane also
can be slid in the radial direction relative to the In the case of skin 43, inwardly extending flange 4
5a is bolted to the fixed structure of the engine at 46. In this case, the fixing structure is flange 2
0 in the radial direction. The rear end of the skin 43 engages the end of the flange 20 by a dog connection 47. In this way, the skin 43 is held at both front and rear ends. Thin inner skin 48 is skin 4
It overlaps with 3. Like skin 33, skin 48 also supports very little load and is therefore very lightweight and very aerodynamically smooth. skin 4
8 has an airfoil-shaped opening 49 through which a land 45 passes. To secure the skin 48 at the front end, a series of hooks 50 are provided which engage under a lip 51 projecting forwardly from the inner skin 43. Skin 48 has a second series of hooks 52 at its rear end that fit into grooves 53 at the rear end of skin 43.

For exactly the same reasons as for the outer platform, it is necessary to seal the gap between the skin 43 and the land 45, and for that purpose the skin 43 is provided with a projection 54 which goes around the airfoil opening 49. A rebate 55 is formed on the inner edge of the protrusion (FIG. 5). This rebate also retains the silica laces 56 to provide the required seal.

Skin 4 forming the inner platform
3,48 has an impingement cooling structure identical to that described for the skin forming the outer platform.

The above-described arrangement makes it possible to separate the support function of the platform from the aerodynamic function, so that the two functions can be properly performed without compromise. The distension difference between the airfoil vanes and the platform, and the movement of the airfoil vanes, is made possible by the radial sliding possible between the airfoil vanes and the various skins. Also, the structure can be made relatively lightweight and parts can be easily replaced for repair.

7 to 10 show another embodiment having the same basic technical concept as the embodiment shown in FIGS. 2 to 6. This embodiment differs from the previous embodiment in the structure that holds the airfoil radially and provides a seal between the airfoil and the skin forming the planetform of the airfoil.

Airfoil vanes 60 are also between the inner and outer platform assemblies. The outer platform assembly includes an outer load-bearing skin 61 and a thin inner skin 6 corresponding to skins 25 and 33, respectively, of the FIG. 2 embodiment.
It consists of 2. The inner platform assembly similarly includes an outer load-bearing skin 63 and a thinner inner skin 64, corresponding to skins 43 and 48, respectively, of the FIG. 2 embodiment. The outer skin may also be perforated to act as an impingement plate. However, in this embodiment, the thin inner skin is provided with flanges 65, 66 around airfoil openings 67, 68 through which the ends of the airfoil vanes 60 pass.

This flange serves two purposes. First, the strength of the thin inner skin is increased so that it can withstand the large loads caused by the gas.

Second, the shoulders 69, 70 of the airfoil blades engage the flanges 65, 66 of the inner skin.
The radial position of the airfoil vanes is maintained by the fixed structure of the engine (casing 22, flange 20) via the inner skin. That is, in this embodiment, the shoulders 69, 70 (protrusions) of the airfoil blades are attached to the fixed retaining structure (casing 22, flange 20).
The radial retention of the airfoil blades is achieved by indirectly engaging the airfoil through the inner skin.

The shoulders and flanges provide positioning for the airfoil blades, but are such that the airfoil blades may have sufficient clearance to accommodate differential expansion.
That is, as can be seen from FIGS. 8 and 9, there is a gap between the flange 65 and the shoulder 69. The airfoil vanes are a sliding fit within the apertures 67, 68 so that they can move radially to accommodate differential expansion.

As in the previous embodiment, the airfoil vanes have openings 67,6
8, it is possible for leakage to occur between the airfoil vanes and the platform. Sealing means in the form of spring washers are therefore provided. The center holes of the spring washers 71, 72 are such that they fit very snugly with the outer surfaces of the corresponding mating portions of the airfoil vanes. FIG. 10 shows the shape of the washer 72. The shape of the washer 72 follows the outer surface of the corresponding portion of the airfoil vane and includes a protrusion 72 that applies the spring force of the washer to the platform.
A is provided.

Further spring washers 73, 74 are provided for sealing the downstream ends of the inner skins 62, 64 to the fixed structure (casing 22, flange 20), respectively. These spring washers are simple rings, and the cross section is such that the spring washers come into pressure contact with hooked flanges 75, 76 protruding from the skin and the fixed structure (casing 22, flange 20). . FIG. 8 shows that the dashed portion 77 of the washer 73 forms a partially conical ring when not elastically deformed. When this portion 77 is sandwiched between the skin 61 and the adjacent fixed structure (casing 22), it comes into pressure contact with the structure (casing 22) to effect sealing. Watusha 73
The main body of the flange 75 is also pressed into contact with the flange 75. Washers 73, 74 can therefore provide an effective seal to the downstream end of the platform.

The invention can be implemented in ways different from the embodiments described above. For example, the structural details for attaching each skin can be changed, and various parts of the assembly can be made of ceramic instead of conventional metal.

[Brief explanation of the drawing]

FIG. 1 is a side view of a gas turbine engine using the nozzle guide vane of the present invention. 2 is an enlarged sectional view of the nozzle guide vane assembly of FIG. 1; FIG. 3 and 4 are exploded views of the inner skin and support skin of the outer platform of the assembly of FIG. 2; FIG.
5 and 6 are enlarged cross-sectional views of a structure for sealing a platoform skin to an airfoil vane; FIGS.
FIG. 7 is a sectional view similar to FIG. 2 of another embodiment. 8 and 9 are enlarged views of the portions marked 8 and 9 of FIG. 7; Figure 10 is arrow 1 in Figure 7.
FIG. 3 is a bottom view of the inner platform looking in the direction 0; 15... Nozzle guide vane assembly, 16, 60...
... Airfoil vanes, 17, 18, 69, 70 ... Projections (pads, shoulders), 20, 22 ... Retaining structures (flanges, casings), 24, 44, 49, 67,
68... Airfoil opening, 25, 43, 61, 63...
Support skin, 33, 48, 62, 64... Inner skin, 41a, 56, 71, 72... Sealing means (silica string, spring washer).

Claims (1)

[Scope of Claims] 1. A nozzle guide vane assembly for a gas turbine engine, comprising a series of airfoil vanes arranged at angular intervals in the circumferential direction, a retaining structure for holding the airfoil vanes, and a retaining structure. a projection on each airfoil adapted to engage directly or indirectly to radially retain the airfoil, and radially inner and outer platforms separate from the airfoil; each platform member comprises a relatively thick and relatively stiff support skin and a relatively thin and relatively less stiff inner skin;
The support skin has at least one airfoil opening through which the airfoil blade passes, the airfoil opening supporting the airfoil blade against torsional, circumferential and axial loads; , the airfoil having at least one airfoil opening therethrough, the inner skin not substantially contributing to the support of the airfoil and forming a boundary for gas flow through the nozzle guide vane assembly. , the airfoil vanes are radially inflatably slidable within the airfoil openings of the support skin and the inner skin, and a sealing means associated with the inner skin comprises:
A nozzle guide vane assembly sealing between the inner skin and the airfoil vane. 2. The nozzle guide vane assembly according to claim 1, wherein the sealing means is an elastic member held from the inner skin and abutting against the outer surface of the airfoil to provide sealing. 3. The nozzle guide vane assembly according to claim 2, wherein the elastic member is a spring washer. 4. The nozzle guide vane assembly according to claim 2, wherein the elastic member is a fibrous heat-resistant material. 5. The nozzle guide vane assembly according to claim 4, wherein the elastic member is a silica string. 6. The nozzle guide vane assembly of claim 1, wherein the protrusion of the airfoil is a pad extending from the end of the airfoil, and the retaining structure is an adjacent fixed structure of the engine. Nozzle guide vane assembly. 7. The nozzle guide vane assembly of claim 1, wherein the protrusion of the airfoil engages the retaining structure indirectly through the inner skin to radially retain the airfoil. The nozzle guide vane assembly is a shoulder provided on the airfoil vane to hold it in place. 8. The nozzle guide vane assembly of claim 7, wherein the inner skin is provided with a flange that acts as an abutment for a shoulder of the airfoil vane. 9. The nozzle guide vane assembly according to claim 1, wherein the support skin has small holes and acts as an impingement plate. 10. The nozzle guide vane assembly of claim 1, wherein the support skin or inner skin is an assembly of a plurality of segments each engaging a plurality of airfoil vanes. 11. The nozzle guide vane assembly of claim 10, wherein the airfoil is sandwiched between adjacent segments constituting the support skin or the inner skin, and the end faces of the adjacent segments are the surfaces of the airfoil. The nozzle guide vane assembly is shaped so that it comes into contact with the respective nozzle guide vanes. 12. The nozzle guide vane assembly of claim 11, wherein a turbine rotor stage is disposed downstream thereof, and the inner skin of the radially outer platform extends downstream to define a flow path through the turbine rotor stage. a nozzle guide vane assembly forming a radially outer boundary;