Method and device for reducing turbine blade temperature

The present invention generally relates to rotor blades of gas turbine engines, and more specifically to methods and devices for lowering turbine blade temperatures.

Gas turbine engine rotor blades generally include an aerofoil having front and trailing edges, a positive pressure side, and a negative pressure side. The positive and negative pressure sides are joined at the anterior and posterior edges of the aerofoil and extend radially between the root and tip of the aerofoil. During operation, the combustion gas that collides with the rotating rotor blades conducts heat to the aero foil. Continuous exposure to hot combustion gases for extended periods of time results in thermal fatigue of the aerofil.

To facilitate the prevention of damage caused by aerofoil due to exposure to hot combustion gases, known aerofoil includes an internal cooling circuit that allows a cooling fluid to flow through the aerofoil. Specifically, at least some known rotor blades flow compressor bleed air into a cavity formed between the side walls to convectively cool the side walls. Further, at least some known cooling circuits utilize shear jet cooling, where a plurality of shear jet openings allow cooling fluid to flow along the inner surface of the sidewall to facilitate cooling of the sidewall. Yet another cooling cavity can be achieved utilizing collision cooling, where the collision insert flushes the cooling fluid through the collision jet array against the inner surface of the front edge of the aerofoil along the front edge. It promotes the cooling of the aero foil. However, these circuits are inefficient as they allow the cooling fluid to flow through the center of the cavity, which is ineffective in removing heat from the walls of the aerofoil.<patcit num="1"><text>Japanese Unexamined Patent Publication No. 2004-028093</text></patcit>

<p> In one embodiment, a method of manufacturing a rotor blade for a gas turbine engine is provided. The rotor blades include an aerofoil having a first side wall and a second side wall that are joined together at the front and trailing edges to form a cavity between them. The method comprises forming at least partially extending rib walls between the first and second sidewalls, the rib walls forming at least one cooling circuit, and each cooling circuit comprising at least three cooling chambers. At least one row of openings in at least one rib wall extending between adjacent cooling chambers, with the first cooling chamber providing cooling fluid to the aerofil cavity, and the remaining chambers open. Includes a step of allowing fluid communication to be coupled to the first cooling chamber via.</p><p> In another aspect, an aerofil for a gas turbine engine is provided. The aerofoil includes a first side wall and a second side wall that are joined together at the anterior and posterior edges to form a cavity between them. A plurality of rib walls extend at least partially between the first and second side walls so that the plurality of rib walls form at least one cooling circuit having at least three cooling chambers. An opening in at least one row extends through at least one rib wall, with a first cooling chamber supplying cooling fluid to the cavity, and the remaining cooling chambers communicating fluid with the first cooling chamber through the openings. It is designed to be combined with.</p><p> In yet another embodiment, a gas turbine engine is provided. The gas turbine engine includes a plurality of rotor blades, each rotor blade being coupled together at a front edge, a trailing edge, and a leading edge and a trailing edge to form a cavity between them. A plurality of rib walls extending at least partially between the first and second side walls and the first and second side walls, and at least one row of openings extending through at least one rib wall. Has an aerofil containing. The plurality of rib walls form at least one cooling circuit having at least three cooling chambers, the first cooling chamber supplies the cooling fluid to the cavity, and the remaining cooling chambers are the first cooling chambers through the openings. It is designed to be connected by fluid communication with.</p>

FIG. 1 is a schematic view of a gas turbine engine 10 including a fan assembly 12, a high pressure compressor 14, and a combustor 16. The engine 10 also includes a high pressure turbine 18, a low pressure turbine 20, and a booster 22. The fan assembly 12 includes an array of fan blades 24 extending radially outward from the rotor disk 26. The engine 10 has a negative pressure side 28 and an exhaust side 30. In one embodiment, engine 10 is a CT7 engine commercially available from General Electric Aircraft Enngines located in Cincinnati, Ohio.

During operation, air flows through the fan assembly 12 and pressurized air is supplied to the high pressure compressor 14. The highly pressurized air is sent to the combustor 16. The airflow from the combustor 16 (not shown in FIG. 1) drives the turbines 18 and 20, which drive the fan assembly 12.

FIG. 2 is a perspective view of the rotor blade 40 that can be used in the gas turbine engine 10 (shown in FIG. 1). FIG. 3 is a cross-sectional view of the rotor blade 40. In one embodiment, a plurality of rotor blades 40 form a high pressure turbine rotor blade stage (not shown) of the gas turbine engine 10. Each rotor blade 40 includes a hollow aerofoil 42 and an integral dovetail 43 used to attach the aerofoil 42 to a rotor disk (not shown) in a known manner.

Aerofoil 42 includes a first side wall 44 and a second side wall 46. The first side wall 44 is convex and forms the negative pressure side of the aerofoil 42, and the second side wall 46 is concave and forms the positive pressure side of the aerofoil 42. The side walls 44 and 46 are joined together at the front edge 48 of the aerofoil 42 and the trailing edge 50 located downstream of the front edge 48 at axially spaced intervals. The aerofoil 42 includes a plurality of film openings 51, the plurality of film openings being arranged radially along the side wall 46 between the aerofoil tip 54 and the blade root 52. The cooling fluid is discharged from the film to promote the cooling of the outer surface of the aerofilm 42. The aerofoil 42 also includes a plurality of trailing edge slots 55 radially spaced between the aerofoil tip 54 and the blade root 52 along the trailing edge 50 to provide cooling fluid from the aerofoil 42. Is discharged to promote cooling of the trailing edge 50 of the aerofoil. Heat transfer enhanced by the film opening 51 and trailing edge slot 55 facilitates cooling along the outer surface 53 of the aerofoil.

The first and second side walls 44 and 46 extend longitudinally from the blade root 52 located adjacent to the dovetail 43 to the aerofoil tip 54, respectively, to form the radial outer boundary of the inner cavity 56. The cavity 56 is formed in the aerofoil 42 between the side wall 44 and the side wall 46. In an exemplary embodiment, the cavity 56 is divided into a plurality of cooling chambers 58 to form a cooling circuit 60 that covers a specific region of the aerofoil 42. In an exemplary embodiment, three cooling circuits 60 are provided. Specifically, in an exemplary embodiment, the cooling circuit 60 includes a front edge circuit 62 for cooling the front edge 48, a positive pressure side circuit 64 for cooling the positive pressure side wall 46, and a negative pressure side wall 44. Includes a negative pressure side circuit 66 for cooling. In another embodiment, the aerofoil 42 has approximately three cooling circuits 60.

The cavity 56 includes a plurality of ribbed walls 70 extending inward. Specifically, in an exemplary embodiment, the rib wall 70 extends radially between the aerofoil tip 54 and the blade root 52 to form a cooling chamber 58 with the aerofoil sidewalls 44 and 46. In another embodiment, the rib wall 70 extends only partially between the aerofoil tip 54 and the blade root. In an exemplary embodiment, the aerofil side walls 44 and / or 46 and the rib walls are approximately the same thickness T.₁Manufactured in. Thus, in an exemplary embodiment, each cooling chamber 58 is bounded by at least one aerofoil side wall 44 and / or 46, and / or at least one rib wall 70. Specifically, the cooling chamber 58 is formed by the inner surface 72 of at least one rib wall 70 and / or the inner surface 74 of at least one aerofoil side wall 44 and / or 46.

Each cooling circuit 60 includes at least one supply chamber 80, at least one transition chamber 82, and at least one discharge chamber 84. In an exemplary embodiment, chambers 80, 82 and / or 84 are separated from each other by a rib wall 70 and fluidly communicate with each other by an opening 88 or a row of slots formed in the rib wall 70 extending between adjacent chambers 58. Be combined. Each row of openings 88 is spaced across each rib wall between the blade root 52 and the aerofoil tip 54. Each transition chamber 82 and discharge chamber 84 extend substantially through the aerofoil 42 between the blade root 52 and the aerofoil tip 54. In an exemplary embodiment, the supply chamber 80 extends through the blade root 52 into the rotor blade dovetail portion 43, where the supply chamber 80 delivers cooling fluid, such as air, through the supply chamber 80, respectively. It is connected to the supply passage (not shown) that supplies the circuit 60 by fluid communication.

During operation, the cooling fluid supplied from the supply chamber 80 to each cooling circuit 60 is flushed into each transition chamber 82 and each discharge chamber 84 before being discharged to the ambient environment. Thus, the cooling circuit 60 supplies a nearly constant flow of cooling fluid through the chambers 80, 82 and 84 to the cavities 56 continuously dispersed along the inner surfaces of the aerofoil sidewalls 44 and 46. Eventually, the cooling fluid is discharged from the cooling passage through the film opening 51 and / or the trailing edge slot 55.

The positive pressure side circuit 64 utilizes a shear jet cooling method in which the high speed shear jet 100 orients the cooling fluid over the inner surface 74 of the side wall 46. In an exemplary embodiment, the positive pressure side circuit 64 includes a positive pressure side supply chamber 90, a positive pressure side transition chamber 92, and a positive pressure side discharge chamber 94. The opening 96 in the first row is formed in the rib wall 70 separating the supply chamber 90 and the transition chamber 94, and the opening 98 in the second row is in the rib wall 70 separating the transition chamber 92 and the discharge chamber 94. Is formed in. In an exemplary embodiment, openings 96 and 98 are adjacent positive pressure side wall inner surfaces 74 where cooling fluid discharged from openings 96 and / or 98 facilitates cooling of the aerofoil side wall 46, thereby the side walls. 46 The operating temperature will drop.

The cooling fluid is supplied to the positive pressure side circuit 64 via the supply chamber 90 and flows from the supply chamber 90 through the opening 96 into the transition chamber 92 as a shear jet indicated by an arrow 100. Specifically, the cooling fluid supplied to the transition chamber 92 flows along the inner surface 74 of the side wall to facilitate cooling of the aerofoil side wall 46. The shear jet 100 is discharged from the transition chamber 92 through the opening 98 into the discharge chamber 94, where the cooling fluid flows along the inner surface 74 of the side wall to facilitate another cooling of the aerofoil side wall 46. The shear jet 100 is then ejected from the aerofoil 42 through a row of film openings 51 extending through the positive pressure side wall 46.

In another embodiment, the positive pressure side circuit 64 includes a supply chamber 90 and a discharge chamber 94, but does not include any transition chamber 92. In yet another embodiment, the positive pressure side circuit 64 includes a plurality of transition chambers 92 coupled to each other by fluid communication between the supply chamber 90 and the discharge chamber 94. In yet another alternative embodiment, the positive pressure side circuit comprises two or more discharge chambers 94 coupled to each other in fluid communication, the positive pressure side film openings 51 draining the cooling fluid from the plurality of cooling chambers 58. To receive. In yet another alternative embodiment, the aerofoil inner surface 74 is a cooling enhancement element (not shown) that facilitates enhanced cooling and heat transfer, such as a turbulator, dimples, bumps or a combination thereof. including.

The negative pressure side circuit 66 utilizes a shear jet cooling method in which the high speed shear jet 130 orients the cooling fluid over the inner surface 74 of the negative pressure side wall 44. In an exemplary embodiment, the negative pressure side circuit 66 includes a negative pressure side supply chamber 110, two negative pressure side transition chambers 112 and 114, and two negative pressure side discharge chambers 116 and 118, wherein the openings 120, 122. , 124, and 126 are formed by rib walls 70 separating adjacent chambers 58 such as chambers 110, 112, 114, 116 and 118. Specifically, the opening 120 in the first row extends between the supply chamber 110 and the first transition chamber 112, and the opening 122 in the second row extends between the first transition chamber 112 and the second transition chamber 114. The opening 124 in the third row extends between the second transition chamber 114 and the first discharge chamber 116, and the opening 126 in the fourth row extends between the first discharge chamber 116 and the first discharge chamber 116. Extends to and from the discharge chamber 118 of 2. In an exemplary embodiment, openings 120, 122, 124 and 126 are adjacent to the negative pressure side inner surface 74 so that the cooling fluid drained from openings 120, 122, 124 and 126 facilitates cooling of the aerofoil side wall 44. Is placed.

The cooling fluid is supplied to the negative pressure side circuit 66 via the supply chamber 110 and flows from the supply chamber 110 through the opening 120 into the first transition chamber 112 as a shear jet 130. Specifically, the cooling fluid supplied to the first transition chamber 112 flows along the inner surface 74 of the side wall to facilitate cooling of the aerofoil side wall 44. The shear jet 130 is expelled from the transition chamber 112 through the opening 122 into the second transition chamber 114, where cooling fluid flows along the inner surface 74 of the side wall to facilitate another cooling of the side wall 44. The shear jet 130 is then discharged from the second transition chamber 114 through the opening 124 into the first discharge chamber 116. The cooling fluid flowing into the first discharge chamber 116 flows along the inner surface 74 of the side wall to promote the cooling of the aerofoil side wall 44. A portion 131 of the cooling fluid is then flushed from the negative pressure side circuit 66 through a row of positive pressure side film openings 132 extending through the positive pressure side wall 46. The remaining cooling fluid flows from the first discharge chamber 116 through the opening 126 into the second discharge chamber 118 as a shear jet 133. The cooling fluid flowing into the second discharge chamber 118 flows along the inner surface 74 of the side wall to facilitate another cooling of the aerofoil side wall 44. The shear jet 134 is then ejected from the aerofoil cavity 56 through a trailing edge slot 55 extending through the aerofoil 42 at the trailing edge 50.

In another embodiment, the negative pressure side circuit 66 includes supply chamber 110 and discharge chamber 118, but not transition chambers 112 and 114 and discharge chamber 116. In another alternative embodiment, the negative pressure side circuit 66 includes one transition chamber 112 or 114 coupled in fluid communication between the supply chamber 110 and the discharge chambers 116 and 118. In yet another alternative embodiment, the negative pressure side circuit 66 has one discharge chamber 116 or 118 fluidly coupled to the supply chamber 110 and the transition chambers 112 and 114, with only the negative pressure side film opening 132. To receive the cooling fluid discharged from one cooling chamber 116 or 118. In yet another alternative embodiment, the aerofoil inner surface 74 includes a cooling enhancement element that facilitates enhanced cooling and heat transfer, such as a turbulator, dimples, bumps or a combination thereof.

In an exemplary embodiment, the aerofoil 42 comprises a leading edge circuit 62 having a leading edge supply chamber 140 and a leading edge discharge chamber 142, using conventional cold bridge collision cooling methods. A row of openings 144 or slots is formed by rib walls 70 that separate the supply and discharge chambers 140 and 142, respectively. The cooling fluid is discharged from the supply chamber 140 through the opening 144 into the discharge chamber 142. The cooling fluid discharged from the opening 144 flows toward the front edge inner surface 74. The cooling fluid is deflected to the pressurized and negative pressure side walls 44 and 46, respectively, where the cooling fluid flows along the inner surface 74 of the side wall to facilitate another cooling of the side walls 44 and 46 of the aerofoil. The positive pressure side film cooling opening 146 and the negative pressure side film cooling opening 148 extend through the side walls 44 and 46, respectively. The cooling fluid is discharged from the cavity 56 through the film cooling openings 146 and 148, respectively, to facilitate another cooling of the sidewalls 44 and 46.

In an exemplary embodiment, the supply chambers 90, 110, and 140 extend from the supply passage (not shown) into the cavity 56 and are located adjacent to each other. In an exemplary embodiment, the supply chambers 90, 110, and 140 are located near the front edge 48 of the thickest portion of the aerofoil 42. This configuration provides a uniform temperature cooled by the thickest portion of Aerofoil 42.

In an exemplary embodiment, the aerofoil 42 includes a first purge chamber 150 and a second purge chamber 152. Purge chambers 150 and 152 are formed by rib walls 70 and are included within the aerofoil cavity 56 to provide structural support. The purge chambers 150 and 152 are not actively cooled by the cooling fluid as in the other chambers 58, but rather by a nearby wall collision process. Specifically, the rib wall 70 forming the purge chambers 150 and / or 152 further forms cooling chambers 58 such as chambers 90, 92, 94, 112, and 114. Therefore, when the cooling fluid is moved through the cooling chamber 58 as described above, heat is transferred from the rib wall 70 to the cooling fluid, so that the operating temperature of the rib wall 70 is lowered. Therefore, the purge chambers 150 and / or 152 are cooled by lowering the operating temperature of the rib wall 70. In another embodiment, the purge chambers 150 and 152 are further cooled when purge air is supplied to the purge chambers 150 and 152. In another alternative embodiment, two or more purge chambers are provided within the aerofoil cavity 56. In yet another alternative embodiment, less than two purge chambers are provided within the aerofoil cavity 56.

In an exemplary embodiment, the rotor blade 40 is manufactured by a casting method utilizing multiple cores integrally formed within a single piece core 166, and in an exemplary embodiment, three separate cores 160. , 162, and 164, respectively. The first core 160 is formed by a positive pressure side circuit 64 and a front edge circuit 62. The second core is formed by the negative pressure side circuit 66 and the third core 164 is formed by the purge chambers 150 and 152. The cores 160, 162, and 164 are incorporated together to form a single core in the casting method. Chambers 58, 150, and 152 are large enough to make ceramic cores 160, 162, and / or 164 in the cast.

The rotor blades described above are cost effective and highly reliable. The rotor blades include an aero foil having several cooling circuits aimed at cooling the front edge of the aero foil, the positive pressure side and the negative pressure side. Several cooling methods are used to cool the outer sidewall of the aerofoil, such as collision cooling, near-wall cooling and shear jet cooling. Yet another cooling enhancement element can be designed within the inner sidewall of the aerofoil. As a result, the cooler operating temperature within the rotor blades allows for a cost-effective and reliable way to extend the useful life of the rotor blades.

Having described the present invention with respect to various specific embodiments, it will be apparent to those skilled in the art that the invention can be practiced with the spirit and improvements within the scope of the claims. The reference numerals described in the claims are for the sake of comprehension and do not limit the technical scope of the invention to the examples.

<figref num="1">Schematic of an exemplary gas turbine engine.</figref><figref num="2">A perspective view of an exemplary rotor blade that can be used in the gas turbine engine shown in FIG.</figref><figref num="3">Sectional view of the rotor blade shown in FIG.</figref>

Code description

44 First side wall 46 Second side wall 48 Front edge 51 Film opening 53 Aerofoil outer surface 55 Rear edge slot 58 Cooling chamber 60 Cooling circuit 62 Front edge circuit 64 Positive pressure side circuit 66 Negative pressure side circuit 70 Rib wall 72 Rib inner surface 74 Side wall inner surface 80 Supply chamber 82 Transition chamber 84 Discharge chamber 88 Opening 90 Positive pressure side supply chamber 92 Positive pressure side transition chamber 94 Positive pressure side discharge chamber