JP2003322002A - Turbine airfoil part provided with metering plate for refreshing hole - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a cooling air circuit for a rotor blade and a stator blade in a turbine. <P>SOLUTION: A cooling circuit in an airfoil part 12 comprises a first, a middle, and a final passages 40, 41, and 42 extended in a radial direction, and the first passage 40 is fluid-communicated with a cooling air source from the external of the airfoil part 12. The final passage 42 is fluid-communicated with one of a front edge and a rear edge. A refreshing passage 66 is extended through a radial direction inner side wall as a boundary of a radial direction inner side part of the final passage 42 while fluid-communicated with the cooling air source. The refreshing passage 66 is disposed separately from the first passage 40 to be apart from the first passage, and it is isolated. In one embodiment, a metering plate 80 covers an inflow port 68 to the refreshing passage 66, and the metering plate 80 has a metering hole 82 over the inflow port. The metering hole 82 is adjustable. <P>COPYRIGHT: (C)2004,JPO

Description

Detailed Description of the Invention

[0001]

FIELD OF THE INVENTION The present invention relates generally to cooling air circuits for turbine rotor blades and stator vanes in gas turbine engines, and more particularly to metering the flow rate to the cooling circuit within the airfoil of the blades and vanes. It relates to a bottom bottom metering plate used for ringing.

[0002]

BACKGROUND OF THE INVENTION Gas turbine engines include a compressor that compresses air that is forced through a flow path into a combustor, where the air is mixed with fuel in the combustor and ignited to produce combustion gases. Combustion gases flow downstream through one or more turbine stages that extract energy from the combustion gases to drive compressors, for example, to drive fans that propel aircraft in flight. To generate additional output. A turbine stage includes a row of turbine rotor blades fixed to the outer circumference of a rotor disk, upstream of which is provided a fixed turbine nozzle having a plurality of stator blades. Combustion gases flow between the stator blades and between the turbine blades to extract energy that rotates the rotor disk. The temperature inside the gas turbine can exceed 2500 degrees Fahrenheit, and cooling the turbine blades is very important from the perspective of blade life. Without cooling, turbine blades deteriorate rapidly. It is highly desirable to improve the cooling of turbine blades, and in order to increase the cooling, those skilled in the art of cooling blades have made many efforts to devise improved geometries of internal cavities within turbine blades. I have been devoting myself. Due to the high temperature of the combustion gases, turbine blades and blades are typically cooled by a portion of the compressed air extracted from the compressor for this purpose. The diversion of any part of the compressor air from its use in the combustor necessarily reduces the overall efficiency of the engine. Therefore, it is desirable to cool the vanes and blades with as little compressed air as possible from the compressor.

Conventional turbine blades and blades include an airfoil over which combustion gases flow. Typically, the airfoil has therein one or more serpentine cooling passages through which compressor bleed air, cooling air, is flowed to cool the airfoil. The airfoil may include various turbulence generators to enhance the cooling effect, and cooling air is exhausted from the passages through various film cooling holes arranged around the outer surface of the airfoil. It In the pursuit of higher cooling effectiveness, modern blades eventually reach multi-path cooling circuits.

It is known to flow cooling air through a serpentine cooling air circuit and other passages inside the blades, which are in the process of moving the cooling air through the passages prior to impingement cooling on the blade leading edge. Preheat. The temperature differential across the blade leading edge is less than if cooling air was directed through the root of the blade to impinge cooling, resulting in lower thermal stress at the blade leading edge and increased blade life. This is because many parts of the blade midspan can be internally cooled before the cooling flow exits the radial leading edge cooling holes to film cool the exterior of the blade airfoil. Efficient use of cooling flow.

However, this technique also adversely affects the "backflow margin". Since the airflow travels through the internal passages of the blade, the pressure losses due to the turnaround and turbulence promotion cool to such a condition that under certain operating conditions hot gas is drawn into the blade leading edge. This will reduce the flow pressure. This undesired condition is called reflux. One way to increase the backflow margin is to increase the inlet pressure of the cooling air supplied to the blades. This approach is not always feasible because increasing the supply pressure will increase the leakage of the cooling flow to an undesirable state. To overcome this problem, a refreshing passage was provided in the leading edge passage that supplies cooling air for collision.
This refreshing passage is connected to the first passage portion of the cooling flow circuit at the blade root. This circuit supplies impingement cooling air to its leading edge cavity from its final passage, referred to herein as the leading edge feed passage. The refresh passage refreshes the airflow in this final flow path after the airflow has been preheated by passing through the rest of the circuit. US Patents 5,387,086 and 5,81
See 3,826.

The refreshing passage to the leading edge supply passage is
It is connected to the inlet flow path of the serpentine cooling circuit that passes through the root of the blade, so that it joins the flow rate of the serpentine cooling circuit. To adjust the flow rate and pressure to the leading edge supply channel,
It would be desirable to be able to regulate the cooling flow through the refreshing passages independently of the flow through the serpentine cooling circuit. This is particularly useful when casting blades are manufactured with a serpentine passage or impingement hole pressure drop higher than desired. It is also desirable for cooling circuit conditioning of blades, and is useful and necessary for blade hole and serpentine circuit degradation, such as caused by plugging and wear.

Known turbine airfoil cooling techniques include the use of internal cavities that form a serpentine cooling circuit. Specifically, meandering passages, leading edge impingement bridges, turbulence enhancers and turbulence generators, film holes, pin fins, trailing edge holes or bleed slots on the pressure side are utilized for blade cooling. It is desirable to provide improved blade cooling. In providing even better blade cooling, it is desirable to avoid a significant increase in blade manufacturing costs.

[0008]

SUMMARY OF THE INVENTION A gas turbine engine hollow airfoil is
From the radially inner bottom, which have pressure-side and suction-side walls joined to each other at the leading and trailing edges of the airfoil, spaced in the width direction and spaced chordwise An airfoil outer wall is provided that extends radially to a radially outer airfoil tip. A cooling circuit within the airfoil extends radially and has first, middle, and final flow passages each arranged in series, the first flow passage in fluid communication with a source of cooling air from outside the airfoil. is doing. The final flow path is in fluid communication with one of the leading or trailing edges. The refreshing passage extends through the radially inner wall that bounds the radially inner portion of the final flow path and is in fluid communication with the cooling air source. The refreshing passage is separate from the first flow passage, spaced apart from the first flow passage, and independent.

In one exemplary embodiment of the present invention, an edge cooling chamber disposed between the final flow passage and one of the leading and trailing edges, and between the final flow passage and the edge cooling chamber. And a cooling air discharge hole disposed through the rib extending in the radial direction. The edge cooling chamber can be a leading edge cooling chamber, the cooling air discharge holes can be impingement cooling holes, and the leading edge cooling holes can be derived from the edge cooling chamber through the outer wall around the leading edge. It can also be included. In another exemplary embodiment, the final flow path is bounded by a trailing edge and a cooling air exhaust hole is disposed therethrough that may be a trailing edge cooling slot.

The exemplary embodiment of the present invention further includes a metering plate covering the inlet to the refreshing passage,
The metering plate has a metering hole on the inlet,
The metering hole is adjustable. Another exemplary embodiment of the present invention is a gas turbine engine blade having a hollow airfoil extending radially outward from a root. The first flow path extends through the root portion and has an inlet on the bottom surface of the root portion. The refreshing passage extends through a radially inner wall that serves as a boundary of a radially inner portion of the final passage and a root portion. The inflow port to the refreshing passage is arranged on the bottom surface of the root portion, and is arranged separately from the first flow passage and apart from the first flow passage. The metering plate is arranged on the bottom surface of the base.

In another embodiment of the present invention, a forward flowing serpentine cooling circuit and a backward flowing serpentine cooling circuit are disposed within the airfoil. Each of the cooling circuits has first, middle and final channels extending radially, each arranged in series, each of the first channels extending through the root,
It has an entrance at the bottom of the root. The final flow path of the serpentine cooling circuit flowing forward is in fluid communication with the leading edge, and the final flow path of the serpentine cooling circuit flowing backward is in fluid communication with the trailing edge. The front and rear refreshing passages extend through the front and rear radial inner walls that bound the respective radially inner portions of the front and rear passages and extend through the root. The refreshing passage has an inflow port on the bottom surface of the root portion, and the inflow port is provided separately from the inlet of the first flow path and apart from the inlet.

The cooling circuit arrangement of the present invention allows the use of lower refrigerant supply pressures. The three-channel serpentine path is also less vulnerable to changes in pressure drop caused by casting characteristics, as compared to circuits with more channels and paths. Dedicated circuits or channels for cooling the leading and trailing edges are
Provides better internal cooling. The refreshing passage is
Because of the leading and trailing edges, cooler air is mixed, thus reducing the amount of refrigerant that must pass through the serpentine cooling circuit and reducing friction and fold-back flow losses. The refreshing passages make a lighter design practical with higher turbine temperatures and fewer cavities for cooling than previous ones. The present invention allows the airfoil and blade weight to be reduced, allowing greater cooling flow at the leading and trailing edges. The present invention can also be useful in protecting airfoils and blades from damage due to foreign objects, severe abrasion, or other damage that causes perforations in the meandering folds. Other circuits lose refrigerant at the holes, causing refrigerant starvation in the remaining serpentine cooling circuits. The refresh circuit provides flow to the root of each cavity to reduce thermal fatigue due to the loss of refrigerant from the serpentine circuit. With the airfoil and blade design of the present invention, if the pressure drop in the serpentine circuit or impingement hole as a result of casting is higher than desired,
The refresh holes can be adjusted to provide higher flow and pressure to the leading and trailing edges, thus increasing productivity and manufacturing yield. According to prior art designs, in this situation the parts had to be scrapped and had to be die reworked for the casting core. The present invention includes an adjustable metering plate so that it can be used to adjust the cooling flow to the leading and trailing edges. A metering plate with metering holes can be brazed on the inlet of the refreshing passage, allowing the use of robust cores and lightweight shanks in the casting, enabling adjustable metering of the refreshing flow rate. To do.

[0013]

The novel features which characterize the invention are set forth and pointed out in the appended claims. The present invention, together with other objects and advantages thereof, will be described more specifically with reference to the accompanying drawings.

FIG. 1 shows an exemplary turbine blade 10 for a gas turbine engine designed to operate in a hot gas stream flowing in an axial downstream direction F. The blade 10 includes a hollow airfoil 12 that extends radially outward from a root 14. The root portion 14 is the center line 11 of the engine.
It is used to secure the blade 10 to an engine rotor disk (not shown) disposed around the. Figure 2
As further shown in the cross-sectional view of airfoil 12 in FIG. 2, airfoils 12 are spaced laterally or laterally and are spaced apart from the upstream leading edge 20 and the chordwise leading edge. It includes an outer wall 15 having a pressure side wall 16 and a suction side wall 18 that are joined to each other along a downstream trailing edge 22. The airfoil 12 has a radial direction 24 that is the span direction of the airfoil 12.
Away from the engine centerline 11 and extending radially along a span S of the airfoil from a radially inner bottom 26 to a radially outer airfoil tip 28. The airfoil tip 28 is an outward extension of the outer wall 15 or squealer wall 2.
Shown as a squealer tip with 9, the squealer wall surrounds an outer tip wall 31 and extends radially outwardly from the tip wall forming a squealer tip cavity 33 therein. Tip cooling holes 59 that extend from inside the hollow airfoil 12 through the outer tip wall 31 to the squealer tip cavity 33 are used to cool the tip cavity. A radially inner bottom 26 is formed in the conventional platform 30 and defines the inner flow boundary of the blade 10 from which the root 14 extends.

During operation of blade 10, combustion gas 32
Are generated by a combustor (not shown), and the airfoil outer wall 1
5 on both the pressure side wall 16 and the suction side wall 18 respectively in the downstream direction. In the exemplary embodiment of the invention shown here, the airfoil 12 is designed to provide efficient cooling to better correspond to the distribution of heat load from the combustion gases 32 to the airfoil 12. The gas turbine blade 10 shown in FIGS. 1-3 is exemplary and the invention applies equally to turbine stator blades having similar airfoils with similar cooling.

With reference to FIG. 2, more specifically, a hollow airfoil 12 having an outer wall 15 is shown in cross-section,
Pressure side wall 16 and suction side wall 18 have leading edge 20 and trailing edge 22.
And are spaced apart from each other in the circumferential or lateral direction. The pressure side wall 16 and the suction side wall 18 are
They are joined together by a plurality of internal transverse ribs, generally designated 34, each extending between the pressure side wall 16 and the suction side wall 18. The first, second, third, and fourth ribs 1-4 of the transverse rib 34 each define a single forward-flowing, three-way serpentine cooling circuit 36, as shown in FIG. The fourth rib 4, the fifth rib 5, the sixth rib 6 and the trailing edge 22 define a single rear-flowing three-way serpentine cooling circuit 38.

In FIG. 3, the cooling circuit cut line 4 in FIG. 2 through the meandering cooling circuits 36 and 38 flowing forward and backward.
6 shows an airfoil 12 arranged flat along 6. In the meandering cooling circuit 36 that flows forward, the meandering cooling flow 35 in the cooling circuit 36 flows in the meandering cooling circuit 36 that flows forward in the trailing edge 22.
To the leading edge 20 in the forward chord flow direction 45. The cooling circuit 36 that flows forward is
An inlet 37 is provided on the bottom surface 49 of the root portion 14, and the meandering cooling flow 3 is provided behind the end portion 39 of the meandering cooling flow circuit that flows forward.
5 are arranged to flow forward from the trailing edge 22 to the leading edge 20 in a forward chord flow direction 45. In the meandering cooling circuit 38 that flows backward, the meandering cooling flow 35 in the meandering cooling circuit 38 that flows backward is directed backward in the meandering cooling flow circuit 38 that flows backward from the leading edge 20 to the trailing edge 22 in the rear chord flow direction. It is configured to flow to 43. The rearwardly flowing serpentine cooling flow circuit 38 comprises an inlet 37 at the bottom surface 49 of the root 14 and a serpentine cooling flow 35 in front of the rearwardly flowing serpentine cooling flow circuit end 39 from the leading edge 20 to the trailing edge 22. Is arranged so as to flow in the rear chord flow direction 43 toward the rear. This better accommodates the heat load applied from the combustion gases 32, allowing the serpentine cooling flow 35 to more efficiently match the heat load on the airfoil 12 and more effectively cool the airfoil. This is because.

The meandering cooling circuits 36 and 38, which flow forward and backward, are referred to as a three-pass circuit, each having three radially extending flow passages designated as a first flow passage 40, an intermediate flow passage 41, and a final flow passage 42. This is to prepare for. The present invention may have more than one intermediate channel as shown here as an exemplary embodiment. The first flow passage 40, the intermediate flow passage 41, and the final flow passage 42 of the meandering cooling circuit 36 flowing forward are spaced in the chord direction, which are shown as the first rib to the fourth rib 1 to 4 in FIG. 3, respectively. The lateral sides 47 are defined by, and are formed between, the internal ribs 34 that are arranged side by side, and their lateral sides 47 are bounded by the pressure side wall 16 and the suction side wall 18 (shown in FIG. 2). .

The first flow passages 40 of the meandering cooling circuits 36 and 38, which flow forwards and backwards, extend radially through the bottom portion 26 of the airfoil 12 and the root portion 14 of the blade 10 to a first radially outer portion. It extends upward in the radial direction to the turn-back passage 50. Meandering cooling circuits 36 and 38 that flow forward and backward
The first flow path 40 of the bottom surface 49 of the root portion 14 of the airfoil portion 12.
Begins at entrance 37. The first turn-back flow passage 50 turns the cooling air radially inward toward the intermediate cooling flow passage 41 (or the intermediate cooling flow passage when there is one or more intermediate cooling flow passages), 41 directs the cooling flow radially inwardly toward the radially inner second fold back channel 52 and then directs cooling air radially upward toward the final cooling channel 42. The final cooling flow path 42 and the serpentine cooling circuits 36 and 38 flowing forward and backward terminate in an outer tip wall 31 where one or more tip cooling holes 59 are vented to the serpentine cooling flow circuit. Can be used as The airfoil squealer tip is cooled by tip cooling holes 59 in the outer tip wall 31.

In the exemplary embodiment shown here, a leading edge cooling chamber 72 is arranged between the leading edge 20 of the outer wall 15 and the first rib 1. In the function of the exemplary embodiment, the discharge hole is the impingement cooling hole 74, which is disposed through the first rib 1 from the final flow path 42 of the meandering cooling circuit 36 flowing forward to the leading edge cooling chamber 72. . The impingement cooling hole 74 extends from the final flow path 42 of the meandering cooling circuit 36 flowing forward to the leading edge cooling chamber 72
Is supplied with cooling air, from which cooling air is flowed through the film cooling holes. The film cooling holes include one or more of the following: film cooling holes 60, 62, 64 in the showerhead, pressure side wall and suction side wall, respectively.

In the exemplary embodiment shown, trailing edge 22 is cooled by cooling air through exhaust holes in the form of trailing edge cooling slots 76 designed to provide convective cooling of trailing edge 22. These two devices are used to cool the leading edge 20 and the trailing edge 22, respectively.

The refreshing passage 66 is a final passage 42 of the meandering cooling circuits 36 and 38 which flow forward and backward.
Extends through the radially inner wall that bounds the radially inner portion 70 of the. The passage 66 for refreshing is at the base 1
4 and in fluid communication with a source of cooling air external to the airfoil 12, and a refreshing passage 66 is independent of the forward and aft serpentine cooling circuits 36 and 38 and operates separately. . The refreshing passage 66 has an inflow port 68 on the bottom surface 49 of the root portion 14, and the inflow port is arranged apart from the inlets 37 of the meandering cooling circuits 36 and 38 flowing forward and backward. Two metering plates 80 are arranged on the bottom surface 49 of the root 14, each covering the respective inlet 68 of the refreshing passage 66. Each of the metering plates has a centrally located metering hole 82 on the inlet. The metering plate is adjustable and can be used to adjust the amount of cooling flow to the leading and trailing edges. Adjustability is done by expanding the pores by expanding the pore size or pore area, or by using plates with different pore sizes. The metering plate 80 can be brazed onto the inlet 68 to the refreshing passage 66 after casting is complete and the blade or airfoil flow test is complete. Independently provided refreshing passages and metering plates allow the use of robust cores and lightweight shanks in casting, allowing for adjustable flow metering for refreshing.

Shown in FIG. 2 is a cross-sectional view of an exemplary airfoil of the present invention. The forward flowing serpentine cooling circuit 36 is shown as having no film cooling holes, while the backward flowing serpentine cooling circuit 38 is shown in the first, middle, and final circuits 40, 41, 42 of the positive pressure side film. It is shown to have cooling holes 62. Film cooling hole 6 on the positive pressure side wall
2 is arranged penetrating the outer wall 15 along the positive pressure side wall 16 of the outer wall 15 of the meandering cooling circuit 38 flowing backward. The film cooling holes are arranged downstream of the center line 11 of the engine and at a combined angle outward in the radial direction, penetrate the outer wall 15, and are led out from the flow path and the leading edge cooling chamber 72. The film cooling holes can be installed along both the pressure side wall 16 and the suction side wall 18 of the outer wall 15, respectively.

The airfoil 12 may comprise any conventional feature for enhancing cooling of the airfoil, such as turbulence generators or pins (not shown) well known in the art. Thermal barrier coatings (TBCs) known in the art may be used to improve the thermal characteristics of the airfoil 12.

Although the present invention has been described with reference to the exemplary turbine blade 10 shown, the present invention provides preferential cooling in the span direction to better accommodate the radial temperature distribution imparted by the combustion gases 32. It can also be used in turbine nozzle vanes with similar airfoils that would find advantages. The airfoils and blades of the serpentine cooling circuit that flow forward and aft can be readily manufactured using conventional casting techniques used in conventional multi-pass serpentine passages.

While we have described what is considered to be a preferred exemplary embodiment of the invention, other variations of the invention will be apparent to those skilled in the art from the teachings herein and, therefore, the invention. All such modifications that come within the spirit and scope of the invention are desired to be covered by the appended claims.

Reference numerals in the claims are provided for facilitating the understanding of the present invention and are not intended to limit the present invention.

[Brief description of drawings]

1 is a perspective view of a rotor blade of a gas turbine engine incorporated into an airfoil of the present invention. FIG.

2 is a schematic cross-sectional view of a spanwise midsection of an airfoil taken through line 2-2 of the airfoil of FIG. 1. FIG.

3 is a cross-sectional view of the exemplary gas turbine engine airfoil shown in FIGS. 1 and 2 arranged flat along the section line of FIG. 2 through a serpentine cooling circuit that flows in a downstream direction.

[Explanation of symbols]

1,2,3,4,5,6 ribs 12 airfoils 14 Root 20 leading edge 22 Trailing edge 26 bottom 28 Tip of airfoil 30 platforms 34 Internal rib 36 Meandering cooling circuit flowing forward 37 entrance 38 Meandering cooling circuit flowing backward 40 First flow path 41 Intermediate channel 42 Final flow path 49 Bottom of base 50 First return flow path 52 Second return flow path 59 Tip cooling hole 60 film cooling holes 66 Refreshing passage 68 Inlet 70 Inner part in radial direction of final flow path 72 Leading edge cooling room 74 Collision cooling hole 76 Trailing edge cooling slot 80 metering board 82 Metering hole

─────────────────────────────────────────────────── ───

[Procedure amendment]

[Submission date] June 19, 2002 (2002.6.1)
9)

[Procedure Amendment 1]

[Document name to be amended] Statement

[Name of item to be corrected] Claim 20

[Correction method] Change

[Correction content]

[Procedure Amendment 2]

[Document name to be corrected] Drawing

[Name of item to be corrected] Figure 1

[Correction method] Change

[Correction content]

[Figure 1]

[Procedure 3]

[Document name to be corrected] Drawing

[Name of item to be corrected] Figure 2

[Correction method] Change

[Correction content]

[Fig. 2]

[Procedure amendment 4]

[Document name to be corrected] Drawing

[Name of item to be corrected] Figure 3

[Correction method] Change

[Correction content]

[Figure 3]

Claims (28)

[Claims] 1. Pressure sidewalls that are joined together at the leading and trailing edges (20, 22) of airfoil (12) spaced apart in the width direction and spaced chordwise, and An airfoil outer wall (15) having suction sidewalls (16, 18) and extending radially from a radially inner bottom (26) to a radially outer airfoil tip (28); A cooling circuit within (12), the cooling circuit having first, middle and final flow paths (40, 41, 42) extending radially and each arranged in series, A first flow path (40) is in fluid communication with a source of cooling air from outside the airfoil (12) and the final flow path (42) is in fluid communication with one of the leading and trailing edges. The refreshing passage (66) has the final passage (42).
Extending through a radially inner wall that is a boundary of a radially inner portion of the fluid passage, and being in fluid communication with the cooling air source, wherein the refreshing passage is separate from the first flow passage (40), A hollow airfoil (12) for a gas turbine engine, wherein the hollow airfoil (12) is spaced apart from the first flow path and is independent. 2. An edge cooling chamber (72) disposed between the final flow passage (42) and the one of the leading edge and the trailing edge, and the final flow passage (42) and the edge cooling chamber (72). Airfoil (12) according to claim 1, characterized in that it further comprises a cooling air discharge hole arranged through a rib extending radially between it and 72). 3. An airfoil according to claim 2, characterized in that the edge cooling chamber (72) is a leading edge cooling chamber and the cooling air discharge holes are impingement cooling holes (74). (12). 4. The leading edge (2) from the edge cooling chamber (72).
Airfoil (12) according to claim 3, further comprising a leading edge cooling hole (74) extending through the outer wall (15) around the (0). 5. The final flow path (42) has the trailing edge (2).
Airfoil (12) according to claim 1, characterized in that 2) is the boundary. 6. The airfoil (12) of claim 5, further comprising a cooling air exhaust hole disposed through the trailing edge (22). 7. Airfoil (12) according to claim 6, characterized in that said cooling air discharge holes are trailing edge cooling slots (76). 8. A metering plate (80) covering an inlet (68) to said refreshing passage (66), said metering plate having a metering hole (82) on said inlet. An airfoil (12) according to claim 1, characterized in that 9. Airfoil (1) according to claim 8, characterized in that the metering hole (82) is adjustable.
2). 10. A hollow airfoil (12) extending radially outwardly from a root (14), the airfoils (12) being laterally spaced apart,
Chordwise having pressure side walls and suction side walls (16, 18) joined together at leading and trailing edges (20, 22) of the airfoil (12), said root Department (1
4) an airfoil outer wall (15) extending radially from a radially inner bottom (26) to a radially outer airfoil tip (28), wherein a cooling circuit is provided in the airfoil (12). First, middle and final flow paths (40, 41, 42) are provided, wherein the cooling circuit extends in a radial direction and is respectively arranged in series.
The first flow path (40) extends through the root portion (14), and has an inlet (3) at the bottom surface of the root portion (14).
7), wherein the final flow path (42) is in fluid communication with one of the leading edge and the trailing edge, and a radial inner wall that bounds a radially inner portion of the final flow path (42). A refreshing passage (66) is provided so as to extend through the root portion (14), and the refreshing passage (66) is formed in the root portion (14).
A gas turbine engine blade (1) having an inlet (68) on the bottom surface of the blade, the inlet being separate from the inlet (37) and spaced from the inlet.
0). 11. An edge cooling chamber (72) disposed between said final flow path (42) and said one of said leading and trailing edges.
And a cooling air discharge hole disposed through a rib extending in a radial direction between the final flow path (42) and the edge cooling chamber (72).
Blade (10) according to item 1. 12. Blade (1) according to claim 11, characterized in that the edge cooling chamber (72) is a leading edge cooling chamber and the cooling air discharge holes are impingement cooling holes (74).
0). 13. A leading edge cooling hole (74) further leading from the edge cooling chamber (72) through the outer wall (15) around the leading edge (20). Item 13. The blade (10) according to item 12. 14. Blade (10) according to claim 10, characterized in that the final channel (42) is bounded by the trailing edge (22). 15. The blade (10) of claim 14, further comprising a cooling air exhaust hole disposed through the trailing edge (22). 16. Blade (10) according to claim 15, characterized in that the cooling air outlet holes are trailing edge cooling slots (76). 17. A metering plate (80) is provided on the bottom surface of the root portion (14), and the metering plate (80) is provided.
Is the inlet (6) to the refreshing passage (66).
Blade (10) according to claim 10, characterized in that it covers 8) and the metering plate has a metering hole (82) on the inlet. 18. Blade (10) according to claim 17, characterized in that the metering hole (82) is adjustable. 19. A hollow airfoil (12) extending radially outwardly from a root (14), said airfoils (12) being laterally spaced apart,
Chordwise having pressure side walls and suction side walls (16, 18) joined together at leading and trailing edges (20, 22) of the airfoil (12), said root Department (1
4) with an outer airfoil wall (15) extending radially from a radially inner bottom (26) to a radially outer airfoil tip (28), flowing forward into said airfoil (12). Meandering cooling circuit (3
6) and a meandering cooling circuit (38) flowing backwards, each of said cooling circuits extending in the radial direction, each being arranged in series, first, middle and final flow paths (40, 41,
42), each of the first flow paths (40) extends through the root portion (14) and has an inlet (37) at the bottom surface of the root portion (14), The final flow path (42) of the meandering cooling circuit (36) flowing in the fluid communication with the leading edge (20), and the final flow path (42) of the meandering cooling circuit (38) flowing backwards is It is in fluid communication with the trailing edge (22) and extends through the root (14) and each of the front and rear radial inner walls that bound the radially inner portion of the front and rear final flow passages. Front and rear refreshing passages (66) are formed so as to extend in a horizontal direction, and the refreshing passages (66) include the root portion (1).
4) has an inlet on the bottom surface, and the inlet is the inlet (3
A blade (10) for a gas turbine engine, characterized in that it is arranged separately from the inlet (7). 20. The meandering cooling circuit (3) flowing in the forward direction.
6) a leading edge cooling chamber (72) arranged between the final flow passages (42), and the forward meandering cooling circuit (3).
6) The final flow path (42) and the leading edge cooling chamber (72)
20. The blade (10) of claim 19, further comprising an impingement cooling hole (74) disposed through a rib extending radially between and. 21. Further comprising a leading edge cooling hole (74) extending from the leading edge cooling chamber (72) through the outer wall (15) around the leading edge (20). Blade (10) according to claim 20. 22. Blade (10) according to claim 21, characterized in that the final channel (42) is bounded by the trailing edge (22) of the meandering cooling circuit (38) flowing backwards. ). 23. The blade (10) of claim 22, further comprising a cooling air exhaust hole disposed through the trailing edge (22). 24. Blade (10) according to claim 23, characterized in that the cooling air outlet holes are trailing edge cooling slots (76). 25. A metering plate (80) is provided on the bottom surface of the root portion (14), and the metering plate (8) is provided.
0) covers the inlet (68) to the refreshing passage (66) and the metering plate (80) has a metering hole (82) on the inlet. Item 24. A blade (10) according to item 24. 26. Blade (10) according to claim 25, characterized in that the metering hole (82) is adjustable. 27. A metering plate (80) is provided on the bottom surface of the root portion (14), and the metering plate (8) is provided.
0) covers the inlet (68) to the refreshing passage (66) and the metering plate (80) has a metering hole (82) on the inlet. Item 19. The blade (10) according to item 19. 28. Blade (10) according to claim 27, characterized in that the metering hole (82) is adjustable.