Patent Grant
12203388
The present disclosure relates to gas turbine engines, and in particular, to turbine rotor blades.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a hot and high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
The turbine section includes turbine vanes to guide and direct the high-speed exhaust gas flow across turbine rotor blades in the turbine section. As the high-speed exhaust gas flow across the turbine rotor blades, the high-speed exhaust gas flow rotates the rotor blades to power the compressor section and/or the fan section. To withstand the high temperatures of the high-speed exhaust gas flow, the turbine vanes and turbine blades require cooling. Cooling air for cooling the turbine vanes and the turbine blades is generally bled from the compressor section and directed to the turbine vanes and the turbine blades. Various cooling schemes have been proposed to optimize the cooling of the turbine vanes and the turbine blades.
A turbine blade includes a platform with a top side and a bottom side opposite the top side. A root section extends from the bottom side of the platform and an airfoil section extends from the top side of the platform to a tip of the turbine blade. The airfoil section includes a leading edge extending from the top side of the platform to the tip. A trailing edge extends from the top side of the platform to the tip and is aft of the leading edge. A pressure side extends from the leading edge to the trailing edge and extends from the top side of the platform to the tip. A suction side extends from the leading edge to the trailing edge and extends from the top side of the platform to the tip. A tip wall is at the tip and extends from the leading edge to the trailing edge. A first core passage extends in a predominately straight direction radially outward from the root section to the tip wall between the leading edge and the trailing edge. An outer first tip flag passage extends in a predominately axial streamwise direction adjacent to the airfoil tip wall from at least one first core passage to a first flag outlet, approximate the airfoil trailing edge. A second tip flag passage extends in predominately an axial streamwise direction toward the leading edge from a second flag outlet approximate the airfoil trailing edge and is between the outer first tip flag passage and the root section. At least one second core passage is between the first core passage and the trailing edge. The second core passage is a serpentine passage that extends in a predominately straight radial direction from the root section to the second tip flag passage. The second core passage is fluidically connected in a predominately axial streamwise direction to the second tip flag passage opposite the second flag outlet approximate the airfoil trailing edge.
A turbine blade includes a base and a tip radially outward from the base in a radial direction. An airfoil section extends from the base to the tip. The airfoil section includes a leading edge extending radially outward from the base to the tip. A trailing edge extends radially outward from the base to the tip and is axially aft of the leading edge in an axial direction. An airfoil pressure side surface extends from the leading edge to the trailing edge and extends from the top side of the platform to the tip. An airfoil suction side surface extends from the leading edge to the trailing edge and extends from the top side of the platform to the tip. The convex suction side airfoil surface is opposite the concave pressure side airfoil surface in a circumferential direction. A tip wall is at the tip and extends axially from the leading edge to the trailing edge. At least one first core passage extends radially from the base to the tip wall between the leading edge and the trailing edge. A first flag wall is spaced radially inward from the tip wall and extends axially from a least one first core passage to the trailing edge. A first tip flag passage is between the tip wall and the first flag wall and extends predominately in an axial direction from the first core passage to a first flag outlet, approximate the airfoil trailing edge. A second flag wall is spaced radially inward from the first flag wall. The second flag wall extends in a predominate axial direction from the airfoil trailing edge toward the at least one first core passage. A second predominately axial tip flag passage is radially between the first flag wall and the second flag wall and extends toward the leading edge from a second flag outlet, approximate the airfoil trailing edge. A second core passage is predominately in an axial direction between the first core passage and the airfoil trailing edge. The at least one second core passage is a serpentine passage that extends from the base to the second tip flag passage. The at least one second core passage is fluidically connected to the second tip flag passage oriented in predominately an axial streamwise direction opposite to the second flag outlet approximate the airfoil trailing edge.
The present summary is provided only by way of example, and not limitation. Other aspects of the present disclosure will be appreciated in view of the entirety of the present disclosure, including the entire text, claims and accompanying figures.
While the above-identified drawing figures set forth one or more embodiments of the invention, other embodiments are also contemplated. In all cases, this disclosure presents the invention by way of representation and not limitation. It should be understood that numerous other modifications and embodiments can be devised by those skilled in the art, which fall within the scope and spirit of the principles of the invention. The figures may not be drawn to scale, and applications and embodiments of the present invention may include features and components not specifically shown in the drawings. Like reference numerals identify similar structural elements.
This disclosure relates to a turbine blade with a first outer tip flag passage oriented in a predominately axial direction adjacent to an outer tip surface of the turbine blade and a second tip flag passage that is radially located inboard under the first outer predominately axially oriented tip flag passage. At least one first core passage is radially oriented and fluidically connected to the first tip flag passage and extends directly from a root of the turbine blade to the predominately axially oriented outer first tip flag passage. Since the at least one first core passage supplies cooling air directly to the first outer predominately axially oriented tip flag passage, the cooling air in the at least one first radial core passage incurs minimal cooling air heat pickup prior to reaching the first outer predominately axially oriented cooling tip flag passage adjacent to the airfoil tip. Thus, the cooling air entering the first tip flag passage from the first core passage is primarily intended to cool the tip of the turbine airfoil blade. At least one second core passage is fluidically connected to the second predominately axially oriented cooling tip flag passage and fluidically extends from a fluidly connected series of predominately radially oriented cooling passages in a serpentine manner from the root of the turbine blade to the second tip flag passage. The at least one second core passage provides cooling air to a central portion of the turbine blade, and the second tip flag passage enables the cooling air flow capacity and mass flow rate in the at least one second core passage to be increased at a relatively high rate. As such the internal convective cooling performance of the at least one second core passage is increased due to the increased internal cavity Mach Numbers and Reynolds numbers. The turbine blade is discussed below with reference to the figures.
Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example, an industrial gas turbine; a reverse-flow gas turbine engine; and a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low-pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high-pressure turbine to drive a high-pressure compressor of the compressor section.
The example gas turbine engine 20 generally includes low speed spool 30 and high speed spool 32 mounted for rotation about center axis A of gas turbine engine 20 relative to engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
Low speed spool 30 generally includes inner shaft 40 that connects fan 42 and low-pressure (or first) compressor section 44 to low-pressure (or first) turbine section 46. Inner shaft 40 drives fan 42 through a speed change device, such as geared architecture 48, to drive fan 42 at a lower speed than low speed spool 30. High-speed spool 32 includes outer shaft 50 that interconnects high-pressure (or second) compressor section 52 and high-pressure (or second) turbine section 54. Inner shaft 40 and outer shaft 50 are concentric and rotate via bearing systems 38 about center axis A.
Combustor 56 is arranged between high-pressure compressor 52 and high-pressure turbine section 54. In one example, high-pressure turbine section 54 includes at least two stages to provide double stage high-pressure turbine section 54. In another example, high-pressure turbine section 54 includes only a single stage. As used herein, a “high-pressure” compressor or turbine experiences a higher pressure than a corresponding “low-pressure” compressor or turbine. The example low-pressure turbine section 46 has a pressure ratio that is greater than about 5. The pressure ratio of the example low-pressure turbine section 46 is measured prior to an inlet of low-pressure turbine section 46 as related to the pressure measured at the outlet of low-pressure turbine section 46 prior to an exhaust nozzle.
Mid-turbine frame 58 of engine static structure 36 can be arranged generally between high-pressure turbine section 54 and low-pressure turbine section 46. Mid-turbine frame 58 further supports bearing systems 38 in turbine section 28 as well as setting airflow entering the low-pressure turbine section 46. Mid-turbine frame 58 includes vanes 60, which are in the core flowpath and function as inlet guide vanes for low-pressure turbine section 46.
The gas flow in core flowpath C is compressed first by low-pressure compressor 44 and then by high-pressure compressor 52. The gas flow in core flowpath C is then mixed with fuel and ignited in combustor 56 to produce high speed exhaust gases that are then expanded through high-pressure turbine section 54 and low-pressure turbine section 46. As discussed below with reference to
In the example of
Top side 84a of platform 84 forms an inner endwall flow surface of turbine blade 78. Bottom side 84b is opposite top side 84a in the radial direction Y and is outside of core flowpath C. Root section 82 extends from bottom side 84b of platform 84. As shown in
Tip 88 of turbine blade 78 is radially outward from base 98 in the radial direction Y. Airfoil section 86 extends from top side 84a of platform 84 to tip 88 of turbine blade 78. Leading edge 90 extends radially outward from top side 84a of platform 84 in the radial direction Y to tip 88. Trailing edge 92 also extends radially outward from top side 84a of platform 84 to tip 88 and is aft of leading edge 90 in the axial direction X.
Pressure side 94 is a generally concave surface of airfoil section 86 that extends from leading edge 90 to trailing edge 92 and also extends from top side 84a of platform 84 to tip 88. Suction side 96 is a generally convex surface of airfoil section 86 that extends from leading edge 90 to the trailing edge 92 and extends from top side 84a of platform 84 to tip 88. Suction side 96 is opposite pressure side 94 in a circumferential direction Z, the circumferential direction Z generally being a direction of rotation of turbine blade 78 about center axis A of gas turbine engine 20 of
In the example of
Second flag wall 103 is spaced radially inward from first flag wall 102. Second flag wall 102 extends axially from trailing edge 92 toward first core passage 108. Second tip flag passage 106 is radially between first flag wall 102 and second flag wall 103 and extends toward axially toward leading edge 90 from second flag outlet 107 on trailing edge 92 to a wall dividing first core passage 108 from second core passage 110. Second core passage 110 is axially between first core passage 108 and trailing edge 92. Second core passage 110 is a serpentine passage extending from root section 82 to second tip flag passage 106. Second core passage 110 fluidically connects to second tip flag passage 106 axially opposite to second flag outlet 107 such that second core passage 110 and second tip flag passage 106 form a continuous passage from root section 82 to second flag outlet 107.
First up pass 124, first bend 126, down pass 128, second bend 129, and second up pass 130 together form the serpentine passage of second core passage 110. First up pass 124 extends from root section 82 toward first bend 126. First bend 126 is radially between root section 82 and second tip flag passage 106. In the example of
Leading edge core passage 112 extends straight and radially from root section 82 to tip wall 100. Leading edge core is between leading edge 90 and first core passage 108 in the axial direction X. A wall is axially between leading edge core passage 112 and first core passage 108 and separates leading edge core passage 112 from first core passage 108 and first tip flag passage 104. Leading edge cavities 116, also referred to as leading edge boxcar cavities 116, are formed axially between leading edge 90 and leading edge core passage 112. Leading edge cavities 116 are radially spaced apart from each other and aligned along leading edge 90. Leading-edge cross-over apertures 118 extend axially from leading edge core passage 112 to leading edge cavities 116 to fluidically connect leading edge cavities 116 with leading edge core passage 112.
Trailing edge core passage 114 extends straight and radially from root section 82 to second flag wall 103 and is axially between second core passage 110 and trailing edge 92 relative to the axial direction X. Trailing edge cavity 120 is formed axially between trailing edge core passage 114 and trailing edge 92 and radially between platform 84 and second flag wall 103. Trailing-edge cross-over apertures 121 extend axially from trailing edge core passage 114 to trailing edge cavity 120 to fluidically connect trailing edge cavity 120 with trailing edge core passage 114. Trailing edge outlets 122 are formed along trailing edge 92 and extend from trailing edge 92 to trailing edge cavity 120.
In the example of
First aperture 132 is formed in first flag wall 102 and extends from first tip flag passage 104 to second tip flag passage 106. In the example of
During operation of turbine blade 78, a supply of cooling air is bled from low-pressure compressor 44 and/or high-pressure compressor 52 (shown in
Cooling air that enters first core passage 108 at root section 82 flows directly up through first core passage 108 to tip wall 100, then turns into first tip flag passage 104 and flows through first tip flag passage 104 to first flag outlet 105. The relatively lower pressure at trailing edge 92 and first flag outlet 105 helps pull the cooling air across first tip flag passage 104 at a relatively fast rate and helps reduce the likelihood of turbulence or stagnation occurring at the turn between first core passage 108 and first tip flag passage 104. As the cooling air moves through first tip flag passage 104, the cooling air cools tip wall 100 and tip 88 of turbine blade 78. Since first core passage 108 is a straight passage with no turns between root section 82 and tip 88, the cooling air reaches the airfoil tip 88 quickly resulting from the relatively short distance the cooling air has to travel. The increase in the cooling air temperature is minimized by mitigating the heat flux and the convection that occurs between the hotter exterior airfoil wall surfaces to the cooling air. Thus, the cooling air temperature heat pickup is significantly reduced. As such the heat that the cooling air can absorb while traveling inside of the at least one first core passage 108 from root section 82 to tip 88 enables greater thermal cooling potential adjacent to the hot airfoil blade tip surface which results in lower operating metal temperatures and increased blade tip durability. During operation of turbine blade 78, tip 88 can be exposed to higher temperatures than any other part of turbine blade 78. Thus, supplying cooling air directly from root section 82 (where the cooling air is the coolest) to tip 88, and minimizing the amount of heat the cooling air absorbs in transit, can be very beneficial to cooling tip 88 extending the operation life of tip 88 and turbine blade 78. Some of the cooling air inside of first core passage 108 flows through fifth aperture 140 to help cool tip 88 and prevent stagnation from occurring in the turn between first core passage 108 and first tip flag passage 104. In examples of turbine blade 78 where tip 88 includes a squealer tip pocket or shelf, fifth aperture 140 can be used to supply cooling air from first core passage 108 to the squealer tip pocket or shelf. Fifth aperture 140 can also be large enough to purge dirt or particulate that happens to enter first core passage 108.
Cooling air that enters the at least one second core passage 110 at root section 82 first flows up through first up pass 124, then turns 180 degrees through first bend 126, then flows radially inward through down pass 128 to second bend 126, turns 180 degrees through second bend 129, and then flows radially outward through second up pass 130. After flowing through second up pass 130, the cooling air in second core passage 110 turns into the second predominately axially oriented tip flag passage 106 and flows axially aftward to second flag outlet 107 approximate the airfoil trailing edge 92. The relatively lower sink pressure at the airfoil trailing edge 92 and second tip flag outlet 107 helps to increase the flow capacity of the cooling air mass flow rate through the serpentine passages of the at least on second core passage 110. The increased mass flow rate enabled by the second axially oriented tip flag passage 106 mitigates the likelihood of internal cooling flow separation, recirculation, or stagnation occurring inside second core passage 110, resulting in significantly reduced internal convective heat transfer, cooling effectiveness, and thermal performance.
It shall be noted that in some embodiments that the flow capacity of second core passage 110 may further be increased by incorporating several film cooling hole apertures along the second predominately axial oriented tip flag passage 106. The addition of film cooling hole apertures increases the cooling mass flow rate in the at least one second core passage 110, which improves the internal convective heat transfer and thermal cooling effectiveness in the central portion of airfoil section 72 of turbine blade 78. The additional film cooling also mitigates local hot external heat flux that is present along the external pressure side airfoil surface, both approximate the second tip flag passage 106, and along the first outer most tip flag passage 104. As such further reductions in local operating metal temperature can be achieved thereby improving the durability of the turbine blade airfoil component. The additional film apertures in the second predominately axially oriented tip flag passage 106 also help mitigate the higher local metal temperatures resulting from the additional cooling air heat pickup observed in the longer second core passage 110.
The at least one second core passage 110 cools a central portion of turbine blade 78. As there are no dead ends inside of second core passage 110, the cooling air through second core passage 110 moves at relatively high flow rates and Mach numbers, which increases heat transfer and cooling of the central portion of turbine blade 78. Second tip flag passage 106 also spaces first bend 126 from tip 88, which decreases the overall length of first up pass 124. Decreasing the overall length of first up pass 124 reduces the amount of time and distance that the cooling air travels in first up pass 124, which reduces the amount of heat the cooling air absorbs before turning in first bend 126 and being directed back towards the cooler temperatures of root section 82. Some of the cooling air inside of first up pass 124 and first bend 126 flows through second aperture 134 and into second tip flag passage 106. The flow of cooling air through second aperture 134 can help the flow of cooling air through first up pass 124 and first bend 126 by reducing stagnation in first bend 126. Sixth aperture 142 can help the flow of cooling air through second bend 129 and second up pass 130 by injecting fresh cooling air from root section 82 into second bend 129. The injection of fresh cooling air from sixth aperture 142 can help cool the flow inside of second core passage 110 and can reduce stagnation at second bend 129. Since the cooling flow in second core passage 110 travels a longer, more circuitous route than the cooling flow in first core passage 108, the pressure in second tip flag passage 106 is lower than the pressure in first tip flag passage 104. This results in a small amount of cooling air inside of first tip flag passage 104 flowing through first aperture 132 into second tip flag passage 106 to prevent stagnation and separation from occurring in the turn between second up pass 130 and second tip flag passage 106. First aperture 132, second aperture 134, and sixth aperture 142 can each be large enough to purge dirt or particulate that happens to enter second core passage 110.
The cooling air that enters trailing edge core passage 114 flows up through leading edge core passage 112 to second flag wall 103. Most of the cooling air inside trailing edge core passage 114 flows through trailing-edge cross-over apertures 121 into trailing edge cavity 120. The cooling air inside of trailing edge cavity 120 can exit trailing edge cavity 120 via trailing edge outlets 122. Some of the cooling air inside of trailing edge core passage 114 flows through third aperture 136 and into second tip flag passage 106 to help reduce or prevent stagnation from occurring in the end of trailing edge core passage 114. Third aperture 136 can also be large enough to purge dirt or particulate that happens to enter trailing edge core passage 114.
First, second, and third apertures 132, 134, 136 can be sized appropriately to balance the amount of cooling flow travelling through the second core passage 110 with the cooling flow temperature exiting second tip flag passage 106 at second tip flag outlet 107. In other words, the larger the first, second, and third apertures 132, 134, and 136, the less cooling flow travelling through the entire serpentine of second core passage 110, but the colder the cooling flow temperature exiting the second tip flag outlet 107 due to a larger portion of the cooling flow exiting second tip flag outlet 107 travelling a shorter distance.
Although
Although not depicted it shall be recognized that internal cooling features such as trip strips, turbulators, circular/oblong pedestals, dimples, delta shaped features of various sizes and shapes may be incorporated and distributed to optimize internal pressure loss, local convective heat transfer and cooling effectiveness requirements to meet component durability requirements.
Although not depicted it shall be recognized that film cooling flow apertures may also be incorporated to further optimize and tailor both internal convective heat transfer and film cooling characteristics. The location, type, quantity, and spacing requirements may be tailored to mitigate turbine airfoil locations that are subjected to higher external heat flux due to external gas temperature distributions and aerodynamic design geometries and loading requirements.
Although not depicted it shall be recognized that the radial passages and axial tip flag passage cavity area distributions may be uniquely sized to meet internal convective cooling, pressure loss, based on allotted turbine blade cooling flow, stage efficiency, and turbine performance efficiency requirements.
Although not depicted it shall be recognized that the invention disclosed herein may also be applied to static turbine vane cooling design applications to mitigate locally high OD and ID airfoil metal temperatures to address local thermal, and thermal-mechanical structural limitations attributed to non-uniformities in vane airfoil and ID/OD platform operating metal temperatures and stresses resulting in thermal mechanical fatigue and creep bending failure mechanisms due to high external unsteady and steady gas pressure loads due upstream blade passing frequencies.
The following are non-exclusive descriptions of possible embodiments of the present invention.
In one example, a turbine blade includes a platform with a top side and a bottom side opposite the top side. A root section extends from the bottom side of the platform and an airfoil section extends from the top side of the platform to a tip of the turbine blade. The airfoil section includes a leading edge extending from the top side of the platform to the tip. A trailing edge extends from the top side of the platform to the tip and is aft of the leading edge. A pressure side extends from the leading edge to the trailing edge and extends from the top side of the platform to the tip. A suction side extends from the leading edge to the trailing edge and extends from the top side of the platform to the tip. A tip wall is at the tip and extends from the leading edge to the trailing edge. A first core passage extends straight from the root section to the tip wall between the leading edge and the trailing edge. A first tip flag passage extends adjacent to the tip wall from the first core passage to a first flag outlet on the trailing edge. A second tip flag passage extends toward the leading edge from a second flag outlet on the trailing edge and is between the first tip flag passage and the root section. A second core passage is between the first core passage and the trailing edge. The second core passage is a serpentine passage that extends from the root section to the second tip flag passage. The second core passage is fluidically connected to the second tip flag passage opposite the second flag outlet.
The turbine blade of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
In another example, a turbine blade includes a base and a tip radially outward from the base in a radial direction. An airfoil section extends from the base to the tip. The airfoil section includes a leading edge extending radially outward from the base to the tip. A trailing edge extends radially outward from the base to the tip and is axially aft of the leading edge in an axial direction. A pressure side extends from the leading edge to the trailing edge and extends from the base to the tip. A suction side extends from the leading edge to the trailing edge and extends from the top side of the platform to the tip. The suction side is opposite the pressure side in a circumferential direction. A tip wall is at the tip and extends axially from the leading edge to the trailing edge. A first core passage extends radially from the base to the tip wall between the leading edge and the trailing edge. A first flag wall is spaced radially inward from the tip wall and extends axially from the first core passage to the trailing edge. A first tip flag passage is between the tip wall and the first flag wall and extends axially from the first core passage to a first flag outlet on the trailing edge. A second flag wall is spaced radially inward from the first flag wall. The second flag wall extends axially from the trailing edge toward the first core passage. A second tip flag passage is radially between the first flag wall and the second flag wall and extends toward the leading edge from a second flag outlet on the trailing edge. A second core passage is axially between the first core passage and the trailing edge. The second core passage is a serpentine passage that extends from the base to the second tip flag passage. The second core passage is fluidically connected to the second tip flag passage axially opposite to the second flag outlet.
The turbine blade of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.
This invention was made with Government support under Contract N00019-21-G-0005 awarded by the United States Naval Air Systems Command. The Government has certain rights in this invention.
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