Patent Application
20190024513
Turbine engines, and particularly gas or combustion turbine engines, are rotary engines that extract energy from a flow of pressurized combusted gases passing through the engine onto a multitude of rotating turbine blades.
Gas turbine engines for aircraft are designed to operate at high temperatures to maximize engine efficiency. It can be beneficial to cool engine components in high-heat environments, in addition to providing additional thermal protection for these components.
In one aspect, an airfoil for a turbine engine comprises a body having an outer wall defining a pressure side and a suction side opposite the pressure side, the outer wall enclosing an interior and extending axially between a truncated nose and a trailing edge, and a shield positioned upstream of the nose to define a leading edge for the airfoil, with the shield spaced from the nose to define a gap between the body and the shield.
In another aspect, an airfoil assembly for a turbine engine comprises radially spaced inner and outer bands and a plurality of airfoils extending between the inner and outer bands, with at least one of the airfoils comprising a body having an outer wall defining a pressure side and a suction side opposite the pressure side, the outer wall enclosing an interior and extending axially between a truncated nose and a trailing edge, and a shield positioned upstream of the nose to define a leading edge for the airfoil, with the shield spaced from the nose to define a gap between the body and the shield.
In yet another aspect, a method of cooling an airfoil in a turbine engine comprises introducing cooling air from an interior of the airfoil into a gap between a shield forming a leading edge of the airfoil and a truncated nose of an airfoil body forming pressure and suction sides of the airfoil, and flowing the cooling air from the gap along at least one of the pressure and suction sides.
In the drawings:
The described embodiments of the present disclosure are directed to a shield for an airfoil for a turbine engine. For purposes of illustration, the present disclosure will be described with respect to the turbine for an aircraft turbine engine. It will be understood, however, that the disclosure is not so limited and may have general applicability within an engine, including compressors, as well as in non-aircraft applications, such as other mobile applications and non-mobile industrial, commercial, and residential applications.
As used herein, the term “forward” or “upstream” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component. The term “aft” or “downstream” used in conjunction with “forward” or “upstream” refers to a direction toward the rear or outlet of the engine or being relatively closer to the engine outlet as compared to another component.
Additionally, as used herein, the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference.
All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader's understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of the disclosure. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and can include intermediate members between a collection of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to one another. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the drawings attached hereto can vary. Furthermore, as used herein, the term “set” or “a set” of elements can refer to any number of elements, including only one.
The fan section 18 includes a fan casing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 disposed radially about the centerline 12. The HP compressor 26, the combustor 30, and the HP turbine 34 form a core 44 of the engine 10, which generates combustion gases. The core 44 is surrounded by core casing 46, which can be coupled with the fan casing 40.
A HP shaft or spool 48 disposed coaxially about the centerline 12 of the engine 10 drivingly connects the HP turbine 34 to the HP compressor 26. A LP shaft or spool 50, which is disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20. The spools 48, 50 are rotatable about the engine centerline and couple to a plurality of rotatable elements, which can collectively define a rotor 51.
The LP compressor 24 and the HP compressor 26 respectively include a plurality of compressor stages 52, 54, in which a set of compressor blades 56, 58 rotate relative to a corresponding set of static compressor vanes 60, 62 to compress or pressurize the stream of fluid passing through the stage. In a single compressor stage 52, 54, multiple compressor blades 56, 58 can be provided in a ring and can extend radially outwardly relative to the centerline 12, from a blade platform to a blade tip, while the corresponding static compressor vanes 60, 62 are positioned upstream of and adjacent to the rotating blades 56, 58. It is noted that the number of blades, vanes, and compressor stages shown in
The blades 56, 58 for a stage of the compressor can be mounted to (or integral to) a disk 61, which is mounted to the corresponding one of the HP and LP spools 48, 50. The vanes 60, 62 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.
The HP turbine 34 and the LP turbine 36 respectively include a plurality of turbine stages 64, 66, in which a set of turbine blades 68, 70 are rotated relative to a corresponding set of static turbine vanes 72, 74 (also called a nozzle) to extract energy from the stream of fluid passing through the stage. In a single turbine stage 64, 66, multiple turbine blades 68, 70 can be provided in a ring and can extend radially outwardly relative to the centerline 12 while the corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the rotating blades 68, 70. It is noted that the number of blades, vanes, and turbine stages shown in
The blades 68, 70 for a stage of the turbine can be mounted to a disk 71, which is mounted to the corresponding one of the HP and LP spools 48, 50. The vanes 72, 74 for a stage of the compressor can be mounted to the core casing 46 in a circumferential arrangement.
Complementary to the rotor portion, the stationary portions of the engine 10, such as the static vanes 60, 62, 72, 74 among the compressor and turbine section 22, 32 are also referred to individually or collectively as a stator 63. As such, the stator 63 can refer to the combination of non-rotating elements throughout the engine 10.
In operation, the airflow exiting the fan section 18 is split such that a portion of the airflow is channeled into the LP compressor 24, which then supplies pressurized air 76 to the HP compressor 26, which further pressurizes the air. The pressurized air 76 from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby generating combustion gases. Some work is extracted from these gases by the HP turbine 34, which drives the HP compressor 26. The combustion gases are discharged into the LP turbine 36, which extracts additional work to drive the LP compressor 24, and the exhaust gas is ultimately discharged from the engine 10 via the exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20 and the LP compressor 24.
A portion of the pressurized airflow 76 can be drawn from the compressor section 22 as bleed air 77. The bleed air 77 can be drawn from the pressurized airflow 76 and provided to engine components requiring cooling. The temperature of pressurized airflow 76 entering the combustor 30 is significantly increased. As such, cooling provided by the bleed air 77 is necessary for operating of such engine components in the heightened temperature environments.
A remaining portion of the airflow 78 bypasses the LP compressor 24 and engine core 44 and exits the engine assembly 10 through a stationary vane row, and more particularly an outlet guide vane assembly 80, comprising a plurality of airfoil guide vanes 82, at the fan exhaust side 84. More specifically, a circumferential row of radially extending airfoil guide vanes 82 are utilized adjacent the fan section 18 to exert some directional control of the airflow 78.
Some of the air supplied by the fan 20 can bypass the engine core 44 and be used for cooling of portions, especially hot portions, of the engine 10, and/or used to cool or power other aspects of the aircraft. In the context of a turbine engine, the hot portions of the engine are normally downstream of the combustor 30, especially the turbine section 32, with the HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28. Other sources of cooling fluid can be, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.
The airfoil 105 can also include a shield 120 positioned upstream of the truncated nose 116, and the shield 120 can have a shape or profile that defines a leading edge 122 of the airfoil 105. The shield 120 can also be spaced from the nose 116 of the body 110 to define a gap 124 as shown. It is contemplated that the material used for the shield 120 can have a higher temperature capability than the material used for the body 110, wherein the temperature capability can be defined as the highest operating temperature contemplated for use for a given material in the turbine engine environment, and subjecting the material to temperatures higher than its temperature capability can cause effects such as oxidation, fatigue, or melting of the material. For example, the shield 120 can be made of ceramic matrix composite (CMC) or monolithic ceramic while the body 110 can be made of metal having a lower temperature capability than the insert material. However, such examples are not meant to be limiting and any material suitable for the turbine engine environment is contemplated for use in this disclosure. Furthermore, the shield 120 can include higher-temperature-capability material at the leading edge 122 only, or throughout the entire shield 120, in non-limiting examples.
It is contemplated that the shield 120 can be mounted between the inner band 102 and outer band 104 in order to define the gap 124. In one example, a tongue-and-groove mounting can be used where the inner band 102 can include a first groove 126A capable of receiving a first tongue 128A of the shield 120, and the outer band 104 can include a second groove 126B for receiving a second tongue 128B of the shield 120. It will be understood that other mountings can be used to secure the shield 120 within the airfoil assembly 100. Furthermore, the shield 120 can be mounted in any desired order during assembly; in a non-limiting example the shield 120 can be first mounted to the inner band 102 by inserting the first tongue 128A into the first groove 126A, and the outer band 104 can then be secured to the shield 120 by inserting the second tongue 128B into the second groove 126B.
Turning to
The outer wall 111 of the airfoil 105 can enclose an interior 138 as shown, and at least one airfoil cooling passage 139 within the interior 138 can supply cooling air to the airfoil 105. A set of cooling holes 140 can extend through the outer wall 111 and fluidly couple the interior 138 to the gap 124. It will be understood that slots or other transpiration cooling methods can be utilized in place of the set of cooling holes 140.
Furthermore, it is contemplated that the use of film holes 136 and cooling holes 140 can depend on a position of the airfoil 105 within the turbine engine 10 (
The gap 124 between the shield 120 and body 110 is illustrated in the example of
The downstream surface 132 of the shield 120 can further include a set of shield surface features 150, and the nose 116 of the body 110 can include a set of nose surface features 152 as shown. The surface features 150, 152 can be formed by casting, machining, or any other desired method as appropriate; in non-limiting examples the surface features 150, 152 can be shaped with a complementary geometry, or the nose surface features 152 can be essentially smooth while the shield surface features 150 can include a curved or wavy profile. It will be understood that any desired geometry or shaping can be used for the surface features 150, 152 on the shield 120 or nose 116.
In operation, a hot airflow H can impinge the apex 134 of the shield 120 and be redirected along the pressure side 112 or suction side 114 of the body 110 as shown in
The turbine engine 10 can further comprise another airfoil assembly 200 according to a second aspect of the disclosure. The airfoil assembly 200 is similar to the airfoil assembly 100, therefore, like parts will be identified with like numerals increased by 100, with it being understood that the description of the like parts of the first aspect applies to the second aspect, unless otherwise noted.
Turning to
A method of cooling the airfoil 105, 205 can include flowing cooling air through the film holes 136 in the shield 130 (
Aspects described in the present disclosure can provide for a variety of benefits. The shield's higher temperature capability can reduce the heat load on the airfoil and protect the airfoil nose from direct impingement of hot gases within the turbine engine, which can prolong the life of the airfoil. It can be appreciated that a reduction in heat load on the airfoil can provide for a reduced amount of cooling air supplied to interior cooling channels within the airfoil. Furthermore, the outer contour of the shield can be aerodynamically designed to replicate the surface of a non-shielded airfoil to maintain aerodynamic efficiency.
In addition, film hole production can be significantly less restrictive when film holes are created in the shield (such as by drilling, lasering, or electrical discharge machining, in non-limiting examples) prior to assembly, as radial surface angles can become restricted in regions proximate the inner and outer bands. Film effectiveness can be improved significantly with surface angle improvement, which further reduces needed cooling flow. Airflows also often contain dust particles that can accumulate within airfoil film holes or within the interior of the airfoil, which can block cooling air from passing through effectively, reduce the cooling benefit provided by the film holes, or reduce the effectiveness of the cooling air due to the insulating effects of accumulated dust. A reduction in supplied cooling air can therefore reduce the amount of dust encountering the airfoil, making the airfoil more dust-tolerant in operation.
It should be understood that application of the disclosed design is not limited to turbine engines with fan and booster sections, but is applicable to turbojets and turboshaft engines as well.
To the extent not already described, the different features and structures of the various embodiments can be used in combination, or in substitution with each other as desired. That one feature is not illustrated in all of the embodiments is not meant to be construed that it cannot be so illustrated, but is done for brevity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not the new embodiments are expressly described. All combinations or permutations of features described herein are covered by this disclosure.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
