Patent Application
20200300107
This application relates to a ceramic matrix composite blade outer air seal.
Gas turbine engines are known and typically include a compressor compressing air and delivering it into a combustor. The air is mixed with fuel in the combustor and ignited. Products of the combustion pass downstream over turbine rotors, driving them to rotate.
It is desirable to ensure that the bulk of the products of combustion pass over turbine blades on the turbine rotor. As such, it is known to provide blade outer air seals radially outwardly of the blades. Blade outer air seals have been proposed made of ceramic matrix composite fiber layers.
In one exemplary embodiment, a blade outer air seal includes a base portion that extends between a first circumferential side and a second circumferential side and from a first axial side to a second axial side. A first wall is axially spaced from a second wall. The first and second walls extend from the base portion. An outer wall joins the first and second walls. The outer wall has a first edge and a second edge. Each of the edges have a first portion and a second portion arranged at a first angle relative to the first portion.
In a further embodiment of the above, the first and second circumferential edges are circumferentially inward of the first and second circumferential sides of the base portion.
In a further embodiment of any of the above, the first portion of the circumferential edges extends in a generally axial direction.
In a further embodiment of any of the above, the first angle is less than about 45°.
In a further embodiment of any of the above, the first angle is less than about 20°.
In a further embodiment of any of the above, the second portion is arranged axially forward of the first portion. A third portion is arranged axially aft of the first portion. The third portion is arranged at a second angle relative to the first portion.
In a further embodiment of any of the above, the second angle is smaller than the first angle.
In a further embodiment of any of the above, the base portion extends axially beyond the first wall.
In a further embodiment of any of the above, a slot extends through the outer wall.
In a further embodiment of any of the above, the first and second edges form mating surfaces configured to engage a support structure or carrier.
In a further embodiment of any of the above, the first wall, the second wall, and the outer wall have a same thickness.
In a further embodiment of any of the above, a film cooling hole extends through the base portion.
In a further embodiment of any of the above, the film cooling hole is between the first and second walls.
In a further embodiment of any of the above, the blade outer air seal is a ceramic matrix composite material.
In another exemplary embodiment, a turbine section for a gas turbine engine includes a turbine blade that extends radially outwardly to a radially outer tip and for rotation about an axis of rotation. A blade outer air seal has a plurality of segments mounted in a support structure via a carrier. The plurality of segments are arranged circumferentially about the axis of rotation and radially outward of the outer tip. Each segment has a first wall axially spaced from a second wall. The first and second walls are joined to a base portion and an outer wall. The outer wall has a first edge and a second edge. Each of the edges have a first portion and a second portion arranged at a first angle relative to the first portion.
In a further embodiment of any of the above, the first and second edges are engaged with the carrier.
In a further embodiment of any of the above, a wear liner is arranged within each segment. The wear liner has a radially extending tab engaged with the first portion.
In a further embodiment of any of the above, the base portion extends between first and second circumferential sides. The first and second circumferential edges are inward of first and second circumferential sides.
In a further embodiment of any of the above, the second portion is arranged axially forward of the first portion. A third portion is arranged axially aft of the first portion. The third portion is arranged at a second angle relative to the first portion. The second angle is smaller than the first angle.
In a further embodiment of any of the above, the blade outer air seal is a ceramic matrix composite material.
These and other features may be best understood from the following drawings and specification.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an 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, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in the exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive a fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in the exemplary gas turbine engine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 may be arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of the low pressure compressor, or aft of the combustor section 26 or even aft of turbine section 28, and fan 42 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7 ° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).
A turbine blade 102 has a radially outer tip 103 that is spaced from a blade outer air seal assembly 104 with a blade outer air seal (“BOAS”) 106. The BOAS 106 may be made up of a plurality of seal segments 105 that are circumferentially arranged in an annulus about the central axis A of the engine 20. The BOAS segments 105 may be monolithic bodies that are formed of a high thermal-resistance, low-toughness material, such as a ceramic matrix composite (“CMC”). The BOAS segments 105 are mounted to a BOAS support structure 110 via an intermediate carrier 112. The support structure 110 may be mounted to an engine structure, such as engine static structure 36. In some examples, the support structure 110 is integrated with engine static structure 36.
A wear liner 162 may be arranged between the seal segment 105 and the carrier 112 in some examples. A feather seal 160 may be used for sealing between circumferential ends C1, C2 of adjacent seal segments 105. The feather seal 160 may extend along the axial length of the BOAS segment 105.
The assembly 104 may include a front brush seal 164 and a dogbone or diamond seal 166 in some examples. These seals 164, 166 are engaged with the leading edge 99 of the BOAS segments 105, and help maintain the axial position of the BOAS 106. The seal 166 pushes the brush seal 164 axially forward and the BOAS segments 105 axially aft.
In the illustrated example, the BOAS segment 105 includes a first axial wall 120 and a second axial wall 122 that extend radially outward from a base portion 124. The first and second axial walls 120, 122 are axially spaced from one another. Each of the first and second axial walls 120, 122 extends along the base portion 124 in a generally circumferential direction along at least a portion of the seal segment 105. The base portion 124 extends between the leading edge 99 and the trailing edge 101 and defines a gas path on a radially inner side and a non-gas path on a radially outer side. An outer wall 126 extends between the first and second axial walls 120, 122. The outer wall 126 includes a generally constant thickness and constant position in the radial direction. The base portion 124, first and second axial walls 120, 122, and the outer wall 126 form a passage 138 that extends in a generally circumferential direction. In this disclosure, forward, aft, upstream, downstream, axial, radial, or circumferential is in relation to the engine axis A unless stated otherwise. The base portion 124 may extend axially forward and aft of the first and second walls 120, 122, and provides a flat surface for sealing of the BOAS leading and trailing edges 99, 101. For example, the base portion 124 includes a portion axially forward of the first axial wall 120 for engagement with seals 164, 166 (shown in
The outer wall 126 has first and second edges 130, 132. The edges 130, 132 have tapered portions. A first portion 131, 133 of the edges 130, 132, respectively, extends generally in the axial direction X. The first portions 131, 133 provide a flat face for engagement with the carrier 112, and help prevent rotation of the seal segment 105 relative to the carrier 112. Tapered portions upstream and downstream of the first portion 131, 133 are angled relative to the axial direction X. A second portion 134, 136 of the edges 130, 132, respectively, is upstream of the first portions 131, 133. The second portions 134, 136 are arranged at a first angle Θ1 with respect to the first portions 131, 133. A third portion 135, 137 of the edges 130, 132, respectively, is downstream of the first portions 131, 133. The third portions 135, 137 are arranged at a second angle Θ2 with respect to the first portions 131, 133. The second and third portions 134, 136, 135, 137 provide tapered faces, which may reduce stresses on the seal segment 105. In one example embodiment, the first and second angles Θ1, Θ2 are less than about 45° with respect to the axial direction X. In another embodiment, the first and second angles Θ1, Θ2 are less than about 20° with respect to the axial direction X. The first angle Θ1 may be greater than the second angle Θ2. In one example, the first angle Θ1 is about 20° and the second angle Θ2 is about 10°.
In the illustrated embodiment, the first portion 131, 133 is generally centered on the outer wall 126. However, in other embodiments, the first portion 131, 133 may be moved axially forward or aft, depending on the carrier 112 and wear liner 162 to address varying torque loads. In one example embodiment, the first portion 131, 133 has a length in the axial direction of about 0.30 inches (7.62 mm). The axial length of the first portion 131, 133 provides a surface for mating with the carrier 112.
Densification includes injecting material, such as a silicon carbide matrix material, into spaces between the fibers in the laminate plies. This may be utilized to provide 100% of the desired densification, or only some percentage. One hundred percent densification may be defined as the layers being completely saturated with the matrix and about the fibers. One hundred percent densification may be defined as the theoretical upper limit of layers being completely saturated with the matrix and about the fibers, such that no additional material may be deposited. In practice, 100% densification may be difficult to achieve. Although a CMC loop BOAS segment 105 is shown, other BOAS arrangements may be utilized within the scope of this disclosure
In an embodiment, the BOAS segment 105 is formed from two loops of CMC laminated plies. A first loop 144 comprises the inner-most layers relative to the respective passage 138. A second loop 146 is formed about the first loop 144 to form the outermost layers relative to the passage 138. In one example embodiment, the first and second loops 144, 146 are each formed from four laminated plies 142. A noodle region 145 may be formed between the first and second loops 144, 146. The noodle region 145 may be filled with a matrix material during densification, in some examples. In some examples, the base portion 124 may have additional reinforcement plies 143. The reinforcement plies 143 may reduce the size of the noodle regions 145, which strengthens the overall structure.
The transverse direction of the plies 142 helps evenly distribute stresses on the component. The shape of the seal segment 105 and the passage 138 allows for complex cooling arrangement and relatively low thermal stresses. The seal segment 105 also allows for multiple sealing surfaces and may accommodate different designs for the intermediate carrier 112. The loop construction of the seal segment 105 also minimizes delamination when the seal segment 105 is secured to the support structure 110 via the carrier 112.
In an example embodiment, the first wall 120, second wall 122, and outer wall 126 have a constant wall thickness of about 8 laminated plies 142, with each plie 142 having a thickness of about 0.011 inches (0.279 mm). This structure may reduce thermal gradient stress. Although 8 laminated plies are described, BOAS constructed of more or fewer plies may fall within the scope of this disclosure. In one example, the first and second loops 144, 146 are formed from laminates wrapped around a core mandrel. In some embodiments, after the laminate plies 142 are formed into a seal segment 105, additional features, such as edges 130, 132 are machined in to form mating surfaces and/or cooling holes. The seal segment 105 may be ultrasonically machined, for example.
The disclosed BOAS arrangement reduces stress on the seal segment 105 by providing edges 130, 132 to engage with the carrier 112. The edges 130, 132 have a flat portion 131, 133 to prevent rotation. The tapered portions 134, 135, 136, 137 of the edges 130, 132 reduce stresses on the seal segment 105. The edges 130, 132 also permit tooling access for machining cooling holes 141 into the base portion 124. The disclosed seal segment 105 permits cost effective manufacturing and assembly and allows the use of a ceramic BOAS. The ability to use a ceramic BOAS promotes a more stable assembly because ceramic materials are not as ductile as metallic materials. The disclosed CMC BOAS has simple features that are easily manufactured using CMC laminates.
In this disclosure, “generally axially” means a direction having a vector component in the axial direction that is greater than a vector component in the circumferential direction, “generally radially” means a direction having a vector component in the radial direction that is greater than a vector component in the axial direction and “generally circumferentially” means a direction having a vector component in the circumferential direction that is greater than a vector component in the axial direction.
Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.
