Aircraft gas turbine engine and method for operating same

Abstract

Provided is an aircraft gas turbine engine that is capable of appropriately adjusting the clearance between moving blade tips and a casing in accordance with the engine load. The aircraft gas turbine engine comprises: a casing; moving blades that are housed in the interior of the casing and rotate around the axis of a rotor by combustion gas; moving blade seal members that are fixed to the casing, provided at positions facing the tips of the moving blades, and seal the combustion gas with a predetermined clearance between said members and tips; a cooling flow passage that cools the moving blade seal members with cooling air guided into the interior of the casing; and an air flow rate adjustment part that adjusts the flow rate of cooling air flowing through the cooling flow passage in accordance with the engine load of the aircraft.

Description

Aircraft gas turbine engine and method of operating the same

[0001] The present disclosure relates to an aircraft gas turbine engine and a method of operating the same.

[0002] In aircraft gas turbine engines such as turbofan engines, an ACC (Active Clearance Control) system that controls the clearance between the tip of a rotor blade and a casing is used.

[0003] Patent Document 1 discloses, as an ACC system for a turbofan engine used in a helicopter, controlling the flow rate of cooling air for cooling a casing. Specifically, it discloses cooling by spraying cooling air from the outside of the casing to a position corresponding to a seal member (Blade Outer Air Seal: BOAS in the document) that faces the tip of the rotor blade via a clearance.

[0004] Patent Document 2 discloses supplying cooling air into the casing to cool a seal member (back plate 47 in the document) that faces the tip of the rotor blade via a clearance.

[0005] U.S. Patent Application Publication No. 2019 / 0078458, Japanese Patent Laid-Open No. 2011-174419

[0006] However, Patent Document 2 does not describe adjusting the flow rate of cooling air when cooling the seal member (back plate 47 in the document) from the inside of the casing.

[0007] On the other hand, since an aircraft performs ground taxiing before takeoff, takeoff, and cruise, the engine load (rating) changes greatly compared to that of a gas turbine engine for power generation, and the clearance between the tip of the rotor blade and the casing changes greatly. It is desired to exert appropriate performance as an aircraft gas turbine engine by adjusting the clearance in accordance with such a changing engine load.

[0008] This disclosure is made in view of these circumstances and aims to provide an aircraft gas turbine engine and a method for operating the same that can appropriately adjust the gap between the tip of the rotor blade and the casing according to the engine load.

[0009] An aircraft gas turbine engine according to one aspect of the present disclosure comprises a casing; rotor blades housed inside the casing and rotating around the axis of a rotor by combustion gases; rotor blade sealing members fixed to the casing and positioned opposite the tips of the rotor blades, sealing the combustion gases through a predetermined gap; a cooling passage for cooling the rotor blade sealing members with cooling air introduced into the casing; and an air flow rate adjustment unit that adjusts the flow rate of cooling air flowing through the cooling passage according to the engine load of the aircraft.

[0010] A method for operating an aircraft gas turbine engine according to one aspect of the present disclosure comprises a casing, rotor blades housed inside the casing and rotating around the axis of a rotor by combustion gas, rotor blade sealing members fixed to the casing and positioned opposite the tips of the rotor blades, sealing the combustion gas through a predetermined gap, and a cooling passage for cooling the rotor blade sealing members with cooling air introduced into the casing, wherein the flow rate of cooling air flowing through the cooling passage is adjusted according to the engine load of the aircraft.

[0011] The gap between the rotor blade tip and the casing can be appropriately adjusted according to the engine load.

[0012] This is a longitudinal cross-sectional view showing an aircraft gas turbine engine according to the first embodiment of this disclosure. This is a longitudinal cross-sectional view schematically showing the cooling structure around the casing. This is a longitudinal cross-sectional view showing an example of a structure for cooling the inside of the casing. This is a longitudinal cross-sectional view showing the control valve of Figure 3. This is a graph showing the engine load during driving, takeoff, and cruising, where (a) is the engine load, (b) is the amount of cooling air, and (c) is the casing temperature. This is a graph showing the diameter of each component and (e) is the gap during driving, takeoff, and cruising. This is a graph showing the gap after gap adjustment during driving, takeoff, and cruising. This is a graph of modification 1 during driving, takeoff, and cruising, where (a) is the engine load, (b) is the amount of cooling air, and (c) is the casing temperature. This is a graph of modification 1 during driving, takeoff, and cruising, where (d) is the diameter of each component and (e) is the gap. This is a graph of modification 1 during driving, takeoff, and cruising, where the gap after gap adjustment is shown. This is a graph of modification 2 during driving, takeoff, and cruising, where (a) is the engine load, (b) is the amount of cooling air, and (c) is the casing temperature. This is a graph of modification 2 during driving, takeoff, and cruising, where (d) is the diameter of each component and (e) is the gap. This is a graph of modification 2 during driving, takeoff, and cruising, showing the gap after gap adjustment. This is a longitudinal cross-sectional view showing an aircraft gas turbine engine according to the second embodiment of this disclosure. This is a graph during driving, takeoff, cruising, and landing, where (a) is the engine load, (b) is the amount of cooling air, and (c) is the casing temperature. This is a graph during driving, takeoff, cruising, and landing, where (d) is the engine load, (e) is the amount of cooling air, and (f) is the rotor temperature. This is a graph during driving, takeoff, cruising, and landing, showing the gap after gap adjustment.

[0013] Embodiments relating to this disclosure will be described below with reference to the drawings. [First Embodiment] The first embodiment of this disclosure will be described below. As shown in Figure 1, the aircraft gas turbine engine (hereinafter simply referred to as "gas turbine") 1 is used as a thrust generating device for an aircraft.

[0014] The gas turbine engine 1 comprises a compressor 10, a combustor 20, a turbine 30, a supercharger fan 40, a casing 50, an outer casing 60, and a shaft 70.

[0015] The compressor 10 is a device that generates compressed air by compressing external air EA flowing in from the front in the direction of the aircraft's movement. The compressor 10 has a plurality of rotor blades 11 that rotate around axis X and are connected to shaft 70, and generates compressed air by passing the incoming external air EA through the plurality of rotor blades 11.

[0016] The combustor 20 is a device that burns compressed air generated by the compressor 10 together with fuel to produce high-temperature, high-pressure combustion gas CA. The combustor 20 supplies the high-temperature, high-pressure combustion gas CA to the turbine 30, thereby rotating the turbine 30 around axis X. Multiple combustors 20 are installed around axis X.

[0017] The turbine 30 is a device driven by the combustion gas CA generated by the combustor 20. The turbine 30 includes a high-pressure turbine 31 and a low-pressure turbine 32 located downstream of the high-pressure turbine 31 in the flow direction of the combustion gas CA. The high-pressure turbine 31 has a plurality of high-pressure turbine blades 31a connected to the shaft 70. The low-pressure turbine 32 has a plurality of low-pressure turbine blades 32a connected to the shaft 70.

[0018] The low-pressure turbine blade 32a has a blade portion 32a1 and a platform portion 32a2 formed integrally with the blade portion 32a1. The blade portion 32a1 is connected to a disk 32a3 attached to the shaft 70 via the platform portion 32a2.

[0019] By guiding the combustion gas CA to the turbine 30, multiple high-pressure turbine blades 31a and multiple low-pressure turbine blades 32a rotate around axis X. The driving force that causes the high-pressure turbine blades 31a and low-pressure turbine blades 32a to rotate around axis X is transmitted to the compressor 10 via shaft 70. The compressor 10 generates compressed air by rotating the blades 11 using the power obtained from the turbine 30.

[0020] The external air EA, compressed by the supercharging fan 40, is supplied to the outside (outer periphery) of the casing 50 as external cooling air A2.

[0021] A portion of the air compressed by the compressor 10 is extracted and supplied to the inside of the casing 50 as internal cooling air A1.

[0022] As shown in Figure 2, external cooling air A2 is introduced, for example, through an introduction pipe 81 to a manifold 82 and ejected from multiple nozzles 83 toward the outside of the casing 50. External cooling air A2 is also introduced, for example, during the aircraft's cruising flight.

[0023] The internal cooling air A1 passes through the inside of the casing 50 and cools the rotor blade seal member 85, which is positioned opposite the tip of the low-pressure turbine rotor blade 32a. The rotor blade seal member 85 is fixed to the casing 50. The rotor blade seal member 85 is provided with a predetermined gap (clearance) d between it and the tip of the low-pressure turbine rotor blade 32a, and the combustion gas is sealed by adjusting this gap d.

[0024] Furthermore, between each low-pressure turbine rotor blade 32a, a low-pressure turbine stator blade 32b is provided, which is fixed to the casing 50.

[0025] Figure 3 shows an enlarged view of an example of a flow path through which the internal cooling air A1 flows. The internal cooling air A1 is drawn from the compressor 10, restricted by the control valve (air flow rate adjustment unit) 87, and then guided into the cooling flow path 90 inside the casing 50. The internal cooling air A1 cools the rotor blade sealing member 85 as it flows through the cooling flow path 90.

[0026] Unlike fixed throttling devices such as orifices, the regulating valve 87 has a variable valve opening (throttling amount) depending on the differential pressure across it.

[0027] As shown in Figure 4 as an example, the regulating valve 87 comprises a main body 87a and a valve body 87b. The main body 87a has a main through-hole 87a1 in the center through which internal cooling air A1 flows, and a seat surface 87a2 at its downstream end. The valve body 87b has a secondary through-hole 87b1 in the center, which is smaller in diameter than the main through-hole 87a1. The valve body 87b is biased toward the seat surface 87a2 of the main body 87a by a coil spring 87c. As a result, when the pressure difference applied before and after the regulating valve 87 is smaller than a predetermined value, the valve body 87b abuts against the seat surface 87a2, closing the surrounding flow path 87d formed between the main body 87a and the valve body 87b. In this case, the internal cooling air A1 flows out only through the secondary through-hole 87b1 of the valve body 87b after being restricted.

[0028] When the pressure difference across the regulating valve 87 exceeds a predetermined value, that is, when the engine load increases and the pressure of the compressed internal cooling air A1 increases, and the pressure applied to the valve body 87b overcomes the biasing force of the coil spring 87c, the valve body 87b separates from the seat surface 87a2. As a result, the internal cooling air A1 flows out not only from the sub-through hole 87b1 but also from the surrounding passage 87d (the state shown in Figure 4).

[0029] The timing of the opening of the surrounding flow path 87d of the control valve 87 and the valve opening degree can be adjusted, for example, by the spring force of the coil spring 87c.

[0030] In this way, the flow rate of the internal cooling air A1 can be increased by changing the throttle amount (valve opening) of the control valve 87 in response to an increase in the pressure of the internal cooling air A1, i.e., an increase in engine load. For example, the flow rate of the internal cooling air A1 can be adjusted according to the aircraft's ground conditions before takeoff, during takeoff, cruising, etc.

[0031] By using a passive flow control valve, such as the control valve 87 shown in Figure 4, the number of parts can be reduced, weight can be lowered, and reliability can be improved compared to actively controlling the system using an electric motor or the like. However, an active flow control valve may also be used. In this case, the engine load can be calculated from permanently measured values ​​such as engine speed and engine exhaust gas temperature.

[0032] Furthermore, the flow control valve is not limited to the control valve 87 shown in Figure 4; any type of valve whose opening degree is variable according to the differential pressure before and after the flow control valve is acceptable.

[0033] Next, using Figures 5A to 5C, we will explain the cooling of the inside of the casing 50 using internal cooling air A1 (cooling of the rotor blade seal member 85).

[0034] The upper graph (a) in Figure 5A shows the engine load according to the aircraft's operating conditions. At time t0, the gas turbine engine 1 is started, and the aircraft travels on the ground at a predetermined engine load from time t1 to t2. Then, at time t2, takeoff begins, and immediately after takeoff, from time t2 to t3, the engine load increases rapidly. From time t3 to t4, the aircraft maintains a high load for takeoff, and from time t4 to t5, the engine load is reduced before entering cruising mode.

[0035] The middle graph (b) in Figure 5A shows the change in the amount of cooling air flowing inside the casing 50. The vertical axis represents the amount of cooling air, and the horizontal axis represents time, as in the upper graph (a). In this figure, the solid line represents the base flow rate, and the dashed line represents the amount of cooling air in this embodiment (flow rate of internal cooling air A1). The same applies to the following graphs, where the solid line represents this embodiment and the dashed line represents the base flow rate.

[0036] Base flow rate is the flow rate when a predetermined throttling amount is used as a fixed value so that the temperature of the casing 50 is within the allowable range when the engine load reaches the design maximum load. Therefore, the base flow rate is the flow rate when a fixed throttling such as an orifice is used instead of the control valve 87 shown in Figure 4, and the flow rate is determined by the pressure before and after the fixed throttling. Accordingly, as shown by the solid line in graph (b) of Figure 5A, the base flow rate is the flow rate corresponding to the engine load (i.e., the extraction pressure from the compressor 10). Note that the "predetermined throttling amount" mentioned above is the throttling amount when a fixed throttling such as an orifice is used instead of the control valve 87 shown in Figure 4, and is generally set so that the temperature of the casing 50 is within the allowable range when the engine load reaches the design maximum load. "Design maximum load" refers to the condition in which the heat input load on the casing 50 is at its maximum, which is the engine load when the aircraft takes off under maximum payload conditions. Note that the allowable temperature of the casing 50 is set to a temperature that satisfies the durability according to the service life of the aircraft engine.

[0037] In contrast, the amount of cooling air in this embodiment is set to be smaller than the base flow rate when the device is on the ground (from time t1 to t2), as shown by the dashed line in graph (b). At this time, the valve body 87b of the regulating valve 87 shown in Figure 4 comes into contact with the seat surface 87a2, closing the surrounding flow path 87d and reducing the valve opening (increasing the throttling amount), so that only a small amount of cooling air flows through the sub-through hole 87b1.

[0038] Then, as shown in graph (b), during takeoff (from time t2 to t4), the amount of cooling air is increased until it reaches a flow rate greater than the base flow rate. At this time, the regulating valve 87 shown in Figure 4 has its valve body 87b separated from the seat surface 87a2, the surrounding flow path 87d opens, the valve opening degree increases (the throttling amount decreases), and the amount of cooling air flowing is corresponding to the pressure difference before and after the regulating valve 87.

[0039] Even during cruising (from time t5 onwards), the amount of cooling air in this embodiment is adjusted so that a flow rate greater than the base flow rate is maintained, similar to the flow rate during takeoff.

[0040] The lower graph (c) of FIG. 5A shows changes in the temperature of the casing 50. The vertical axis represents casing temperature, and the horizontal axis represents the same time as in the upper graph (a).

[0041] As shown in this figure, the casing temperature of the present embodiment is higher when the aircraft is on the ground (from time t1 to t2) than when the base flow rate is used. This is because the cooling air flow rate in the present embodiment is set smaller than the base flow rate, so the amount of cooling for the casing 50 is reduced.

[0042] The casing temperature during takeoff and cruise (after time t3) is lower in the present embodiment than when the base flow rate is used, because the cooling air flow rate in the present embodiment is set larger than the base flow rate. Therefore, in the present embodiment (broken line), as indicated by the temperature difference ΔT in graph (c), the difference in casing temperature between when on the ground and when taking off is smaller than that in the case of the base flow rate (solid line). During cruise, the casing temperature of the present embodiment (broken line) is kept lower than that in the case of the base flow rate (solid line).

[0043] The upper graph (d) of FIG. 5B shows changes in the diameter of each member after thermal deformation. The vertical axis represents the diameter of each member (after thermal deformation), and the horizontal axis represents the same time as graphs (a) to (c) of FIG. 5A. The broken line in graph (d) indicates the inner diameter of the casing (the inner diameter of the blade seal member 85 (see FIG. 3)) in the present embodiment, the solid line indicates the inner diameter of the casing (the inner diameter of the blade seal member 85) in the case of the base flow rate, and the dotted line indicates the outer diameter of the tip of the low-pressure turbine rotor blade 32a (see FIG. 3). Therefore, the difference between the broken line and the dotted line is the gap d (see FIG. 3) in the present embodiment, and the difference between the solid line and the dotted line is the gap d in the case of the base flow rate. This gap d is shown in the lower graph (e) of FIG. 5B.

[0044] The radial deformation of the low-pressure turbine rotor blade 32a increases not only due to thermal deformation caused by temperature rise, but also due to centrifugal force proportional to the engine rotation speed. Therefore, as shown in graph (d), immediately after takeoff (from time t2 to t3), the outer diameter of the low-pressure turbine rotor blade 32a rises sharply.

[0045] As can be seen from graph (d) and graph (e), during ground operation (from time t1 to t2), the casing temperature is higher in the present embodiment (broken line) than in the base flow rate case (solid line) (see graph (c) in FIG. 5A), and thermal expansion becomes larger. Therefore, the gap d is larger in the present embodiment than in the base flow rate case.

[0046] On the other hand, during cruising (after time t5), the casing temperature is lower in the present embodiment than in the base flow rate case (see graph (c) in FIG. 5A), and thermal expansion becomes smaller. Therefore, the gap d is smaller in the present embodiment than in the base flow rate case. This reduces leakage of combustion gas from the gap d and improves engine performance.

[0047] Immediately after takeoff (from time t2 to t3), in the present embodiment, the temperature is increased during ground operation to increase thermal expansion, so the inner diameter of the casing 50 is kept large. For this reason, at the pinching point P where the low-pressure turbine rotor blade 32a comes closest to the inner diameter of the casing 50 (that is, the rotor blade seal member 85), there is a marginal gap d1 (see graph (e)) compared to the base flow rate case where the gap d is substantially 0. Generally, turbine rotor blades rapidly extend in the radial direction due to centrifugal force and thermal expansion immediately after takeoff. On the other hand, since thermal deformation of the casing is slower than that of turbine rotor blades, the gap between the inner diameter of the casing and the outer diameter of the low-pressure turbine rotor blade becomes the smallest during takeoff. This is the pinching point where the low-pressure turbine rotor blade 32a comes closest to the inner diameter of the casing 50 (that is, the rotor blade seal member 85).

[0048] If the gap d at the time of design (in a cold state) is reduced so as to reduce this marginal gap d1, the gap d during operation can be further reduced. Specifically, the gap d is reduced so as to shift the solid line of graph (e) in FIG. 5B downward. Then, as shown by the alternate long and short dash line in graph (f) of FIG. 5C, by setting the gap d to be substantially 0 at the pinching point P, the gap d during takeoff after time t3 and during cruising can be further reduced. In particular, since the gap d during cruising, which has a significantly longer operation time than during ground operation and takeoff, can be reduced, engine performance can be improved.

[0049] This embodiment provides the following advantages: The flow rate of cooling air that cools the blade seal member 85, which forms a gap with the tip of the low-pressure turbine blade 32a, is adjusted according to the engine load. This allows the gap d between the tip of the low-pressure turbine blade 32a and the blade seal member 85, which is the inner diameter of the casing 50, to be appropriately adjusted according to the engine load.

[0050] By adjusting the flow rate of cooling air using the control valve 87 during ground operation, takeoff, and cruising, when the engine load changes significantly, an appropriate gap d can be provided under each operating condition.

[0051] While on the ground, cooling air is supplied to the rotor blade seal member 85 at a flow rate lower than the base flow rate. This reduces the cooling of the rotor blade seal member 85 compared to the base flow rate while on the ground before takeoff, promoting a temperature rise in the rotor blade seal member 85 and causing thermal expansion, thereby increasing the gap d between it and the tip of the low-pressure turbine rotor blade 32a. Therefore, even if the low-pressure turbine rotor blade 32a extends radially due to centrifugal force and thermal expansion during takeoff, the gap with the rotor blade seal member can be maintained. Consequently, the gap designed for the cold state can be made as small as possible.

[0052] During cruising, cooling air is supplied to the rotor blade seal member 85 at a flow rate greater than the base flow rate. This increases the cooling of the rotor blade seal member 85 compared to the base flow rate, thereby promoting cooling of the rotor blade seal member 85 and reducing the gap d between it and the tip of the low-pressure turbine rotor blade 32a, thereby improving engine performance.

[0053] The flow rate of cooling air supplied to the rotor blade seal member 85 is adjusted by a control valve 87 whose flow path area passively changes due to differential pressure. This allows for a simpler and lighter structure compared to electrically operated valves that actively control the valve opening, making it advantageous for aircraft applications.

[0054] <Modification 1> In the embodiment described above, the amount of cooling air was set to be less than the base flow rate while on the ground, and after takeoff, the cooling flow rate was set to be more than the base flow rate (see graph (b) in Figure 5A), but the disclosure is not limited thereto.

[0055] For example, it can be transformed as shown in Figures 6A to C. Graphs (a) to (f) in Figures 6A to C correspond to the respective graphs in Figures 5A to C.

[0056] As shown in graph (b) of Figure 6A, during ground flight (from time t1 to t2), the amount of cooling air is reduced to approximately 0, for example, below the base flow rate, similar to the embodiment described above. On the other hand, after takeoff (from time t3 onwards), unlike the embodiment described above, the amount of cooling air is set to the same as the base flow rate.

[0057] Even with this type of operation, as shown in graph (f) of Figure 6C, it is possible to design the gap d at the pinching point P to be approximately 0. This reduces the gap during cruising, which has a significantly longer operating time compared to during taxiing and takeoff, and improves engine performance.

[0058] <Modification 2> Alternatively, the amount of cooling air may be adjusted as shown in graph (b) of Figure 7A, relative to graph (b) of Figure 5A. Graphs (a) to (f) of Figures 7A to C correspond to the respective graphs of Figures 5A to C. Also, for comparison, the dashed lines shown in Figures 5A and B (cooling air amount in the above embodiment) are retained in Figures 7A to C. The cooling air amount in this modification is shown by a thick dashed line.

[0059] As shown in graph (b) of Figure 7A, the cooling air volume (thick dashed line) of this modified example is less than the base flow rate (solid line) when on the ground, similar to the dashed line in the above embodiment. On the other hand, when taking off, this modified example uses the same cooling air volume as the base flow rate (solid line), and when cruising, the cooling air volume is greater than both the base flow rate (solid line) and the above embodiment (thin dashed line).

[0060] As a result, in this modified example (thick dashed line), the inner diameter of the casing 50 deforms more significantly during takeoff than in the above embodiment (thin dashed line), but the inner diameter of the casing 50 can be made smaller during cruising than in the above embodiment (thin dashed line) (graph (c) in Figure 7A and graph (d) in Figure 7B). Furthermore, as shown in graph (f) in Figure 7C, the gap can be reduced during both takeoff and cruising.

[0061] In the embodiments and their modifications described above, the gap at the tip of the low-pressure turbine blade 32a was explained, but this disclosure is not limited thereto and can also be applied to the gap of the high-pressure turbine blade 31a. The same applies to the embodiments described below.

[0062] [Second Embodiment] Next, a second embodiment of the present disclosure will be described. In addition to the first embodiment which cools the inside of the casing 50, this embodiment also adjusts the flow rate of rotor cooling air which cools the inside of the rotor that supports the rotor blades.

[0063] Figure 8 shows an aircraft gas turbine engine 1 according to this embodiment. The gas turbine engine 1 shown in Figure 8 has some differences in its configuration from the gas turbine engine 1 shown in Figure 1, but its basic configuration is the same. Therefore, components that are the same as in the first embodiment are denoted by the same reference numerals and their descriptions are omitted.

[0064] The internal cooling air A1 used to cool the inside of the casing 50 is extracted air drawn from the intermediate stage of the compressor 10, similar to the first embodiment.

[0065] The rotor 35 supports the rotor blades 30a and rotates around axis X. Inside the rotor 35, a rotor cooling passage 35a is formed to cool the rotor 35 from the inside. Air extracted from the intermediate stage of the compressor 10 is introduced into the rotor cooling passage 35a as rotor cooling air A3. The flow rate of the rotor cooling air A3 is passively adjusted by a control valve 87, as illustrated in Figure 4, similar to the internal cooling air A1.

[0066] Next, using Figures 9A to 9C, we will explain the cooling of the inside of the casing 50 using internal cooling air A1 and rotor cooling air A3 (cooling of the rotor blade sealing member 85 (see Figure 3)).

[0067] Graphs (a) to (c) in Figure 9A show the internal cooling air A1 that cools the inside of the casing 50. The upper graph (a) in Figure 9A shows the engine load according to the aircraft's operating conditions. This graph (a), like graph (a) in Figure 5A, shows ground, takeoff, and cruising, and further shows landing from time t6 onwards.

[0068] The middle graph (b) in Figure 9A, similar to the graph (b) in Figure 5A, shows the amount of cooling air when the flow rate of internal cooling air A1 (dashed line) is changed relative to the base flow rate (solid line). Therefore, during travel (from time t1 to t2), the amount of cooling air decreases in accordance with the decrease in internal cooling air A1, and after takeoff (from time t3 onward), the amount of cooling air increases in accordance with the increase in internal cooling air A1.

[0069] The lower graph (c) in Figure 9A shows the casing temperature, similar to graph (c) in Figure 5A. In this embodiment, as shown by the temperature difference △T in graph (c), the difference in casing temperature between ground operation and takeoff (dashed line) is smaller than in the case of base flow rate (solid line). Also, the inner diameter of the casing 50 after thermal deformation (inner diameter of the rotor blade seal member 85) shows a similar change to the temperature change in graph (c), with the radius being larger on the ground than in the case of base flow rate, and smaller during takeoff and cruising than in the case of base flow rate.

[0070] Figure 9B shows a graph of the rotor 35 being cooled by the rotor cooling air A3. The upper graph (d) in Figure 9B is the same as graph (a) in Figure 9A.

[0071] The middle graph (e) in Figure 9B corresponds to graph (b) in Figure 9A and shows the amount of cooling air in response to changes in rotor cooling air A3. The flow rate of rotor cooling air A3 (dashed line) is increased to be greater than the base flow rate for the rotor (solid line) when on the ground, thereby increasing the amount of cooling air. During takeoff and cruising, the flow rate of rotor cooling air A3 is reduced to be less than the base flow rate for the rotor, thereby decreasing the amount of cooling air.

[0072] The rotor base flow rate is the flow rate when a predetermined throttling amount is used as a fixed value so that the temperature of the rotor 35 reaches an allowable value when the engine load is at the design maximum load. Therefore, the rotor base flow rate is the flow rate when a fixed throttling such as an orifice is used instead of the control valve 87 shown in Figure 4, and the flow rate is determined by the pressure before and after the fixed throttling. Accordingly, as shown by the solid line in the middle graph (e) of Figure 9B, the base flow rate is the flow rate corresponding to the engine load (i.e., the extraction pressure from the compressor 10).

[0073] As shown in graph (e) of Figure 9B, the rotor temperature shown in graph (f) of Figure 9B is achieved by adjusting the rotor cooling air A3 and changing the amount of cooling air. As shown in graph (f), when on the ground, the rotor 35 is cooled more than in the case of the base flow rate, so the rotor temperature is lower than in the case of the base flow rate. On the other hand, during takeoff and cruising, the amount of cooling to the rotor 35 is reduced compared to the case of the base flow rate, so the rotor temperature is higher than in the case of the base flow rate. Therefore, the radius of the tip of the rotor blade 30a attached to the rotor 35 after thermal deformation shows a change similar to the temperature change in graph (f), with the radius of the tip of the rotor blade 30a being smaller when on the ground than in the case of the base flow rate, and the radius of the tip of the rotor blade 30a being larger during takeoff and cruising than in the case of the base flow rate.

[0074] Graph (g) in Figure 9C shows the difference between the radius of the casing (inner diameter of the rotor blade seal member 85) and the radius of the tip of the rotor blade 30a, i.e., the gap d (see Figure 3). As is clear from the graph, when on the ground, the radius of the casing increases and the radius of the tip of the rotor blade 30a decreases, so the margin gap d1 (see graph (e) in Figure 5B) increases. If the gap is managed so that this margin gap d1 is approximately 0 during the design phase (cold state), then, as shown by the dashed line in graph (g), the gap d can be made smaller than in the case of base flow rate (solid line) over the entire range of ground, takeoff, and cruising.

[0075] The effects of this embodiment described above are as follows. On the ground, rotor cooling air A3 is supplied to the rotor 35 at a flow rate greater than the rotor base flow rate. This increases the cooling of the rotor 35 compared to the rotor base flow rate on the ground before takeoff, thereby reducing the temperature of the rotor 35, suppressing thermal expansion, and increasing the gap with the tip of the rotor blade 30a. Furthermore, even if the rotor blade 30a extends radially due to centrifugal force and thermal expansion during takeoff, the gap with the rotor blade sealing member 85 can be maintained. For this reason, the gap d designed for the cold state can be made as small as possible.

[0076] During cruising, rotor cooling air A3 is supplied to the inside of the rotor 35 at a flow rate lower than the rotor base flow rate. This reduces the cooling of the rotor blade seal member 85 compared to the rotor base flow rate, thereby suppressing the cooling of the rotor 35, reducing the gap d between the rotor blade 30a and its tip, and improving engine performance.

[0077] The aircraft gas turbine engines and their operating methods described in each of the embodiments above can be understood, for example, as follows.

[0078] An aircraft gas turbine engine (1) according to a first aspect of the present disclosure includes a casing (50), rotor blades (30, 32a) housed inside the casing and rotating around the axis of the rotor by combustion gas, a rotor blade sealing member (85) fixed to the casing and positioned opposite the tip of the rotor blade, sealing the combustion gas through a predetermined gap (d), a cooling passage (90) for cooling the rotor blade sealing member with cooling air (A1) introduced into the casing, and an air flow rate adjustment unit (87) that adjusts the flow rate of cooling air flowing through the cooling passage according to the engine load of the aircraft.

[0079] The flow rate of cooling air that cools the rotor blade seal member, which forms a gap with the tip of the rotor blade, is adjusted according to the engine load. This allows the gap between the tip of the rotor blade and the casing to be appropriately adjusted according to the engine load.

[0080] In the second aspect of the present disclosure, the aircraft gas turbine engine, in the first aspect, has an airflow adjustment unit that adjusts the flow rate of cooling air flowing through the cooling passage during grounding before takeoff, during takeoff, and during cruising.

[0081] By adjusting the flow rate of cooling air during ground operation, takeoff, and cruising, when engine load changes significantly, it is possible to provide an appropriate gap under each operating condition.

[0082] In the third aspect of the present disclosure, in the second aspect, the airflow adjustment unit reduces the flow rate by increasing a predetermined throttling amount during ground operation, which is such that the temperature of the casing reaches an allowable value when the engine load reaches the maximum design load.

[0083] The design throttling amount is defined as a predetermined throttling amount that specifies the flow rate of cooling air supplied to the rotor blade seal member so that the casing temperature remains within an acceptable range when the engine load reaches the design maximum load (e.g., 100% engine load). The flow rate of cooling air when this design throttling amount is used as a fixed value is called the base flow rate. On the ground, cooling air is supplied to the rotor blade seal member at a flow rate lower than the base flow rate. This reduces the cooling of the rotor blade seal member compared to the base flow rate on the ground before takeoff, promoting a temperature rise in the rotor blade seal member and causing thermal expansion, thereby increasing the gap with the tip of the rotor blade. Then, even when the rotor blade extends radially due to centrifugal force and thermal expansion during takeoff, the gap with the rotor blade seal member can be maintained. For this reason, the gap designed for cold conditions can be made as small as possible.

[0084] In the fourth aspect of the present disclosure, the aircraft gas turbine engine, in the second or third aspect, increases the airflow rate by reducing a predetermined throttling amount during cruising, which is the amount by which the casing temperature reaches an allowable value when the engine load reaches the design maximum load.

[0085] The design throttling amount is defined as a predetermined throttling amount that specifies the flow rate of cooling air supplied to the rotor blade seal member so that the casing temperature remains within an acceptable range when the engine load reaches the design maximum load (for example, 100% engine load). The flow rate of cooling air when this design throttling amount is used as a fixed value is called the base flow rate. During cruising, cooling air is supplied to the rotor blade seal member at a flow rate greater than the base flow rate. This increases the cooling of the rotor blade seal member beyond the base flow rate, thereby promoting cooling of the rotor blade seal member, reducing the gap with the tip of the rotor blade, and improving engine performance.

[0086] In the fifth aspect of the present disclosure, the aircraft gas turbine engine, in any of the first to fourth aspects described above, includes a flow control valve (87) in which the flow path area is passively changed by differential pressure in the air flow control unit.

[0087] The flow rate of cooling air supplied to the rotor blade seal member is adjusted using a flow control valve whose flow path area passively changes due to differential pressure. This allows for a simpler and lighter structure compared to electrically operated valves that actively control the valve opening, making it advantageous for aircraft applications.

[0088] An aircraft gas turbine engine according to a sixth aspect of the present disclosure, in any of the second to fifth aspects described above, is provided with a rotor cooling passage (35a) that cools the rotor (35) with rotor cooling air (A3) introduced into the rotor (35), and the air flow rate adjustment unit increases the flow rate of the rotor cooling air by reducing a predetermined throttling amount during ground operation, which is such that the rotor temperature reaches an allowable value when the engine load becomes the maximum design load.

[0089] The design throttle is defined as a predetermined throttle amount that specifies the flow rate of rotor cooling air supplied to the inside of the rotor so that the rotor temperature remains within an acceptable range when the engine load reaches the design maximum load (for example, 100% engine load). The flow rate of rotor cooling air when this design throttle amount is used as a fixed value is called the rotor base flow rate. On the ground, the rotor is supplied with cooling air at a flow rate greater than the rotor base flow rate. This increases the rotor cooling compared to the rotor base flow rate on the ground before takeoff, thereby reducing the rotor temperature, suppressing thermal expansion, and increasing the clearance between the rotor blades and their tips. Furthermore, even when the rotor blades extend radially due to centrifugal force and thermal expansion during takeoff, the clearance with the rotor blade seal can still be maintained. Therefore, the clearance designed for cold conditions can be made as small as possible.

[0090] An aircraft gas turbine engine according to a seventh aspect of the present disclosure, in any of the second to sixth aspects described above, is equipped with a rotor cooling passage for cooling the rotor with rotor cooling air introduced into the rotor, and the air flow rate adjustment unit reduces the flow rate of the rotor cooling air by increasing a predetermined throttle amount during cruising so that the rotor temperature reaches an allowable value when the engine load becomes the maximum design load.

[0091] The design throttle is defined as a predetermined throttle amount that specifies the flow rate of rotor cooling air supplied to the inside of the rotor so that the rotor temperature remains within an acceptable range when the engine load reaches the design maximum load (for example, 100% engine load). The flow rate of rotor cooling air when this design throttle amount is used as a fixed value is called the rotor base flow rate. During cruising, cooling air is supplied to the inside of the rotor at a flow rate lower than the rotor base flow rate. This reduces the cooling of the rotor blade seal area compared to the rotor base flow rate, thereby suppressing rotor cooling, reducing the gap between the blade tip and the rotor blade, and improving engine performance.

[0092] A method for operating an aircraft gas turbine engine according to a first aspect of the present disclosure comprises a casing, rotor blades housed inside the casing and rotating around the axis of a rotor by combustion gas, rotor blade sealing members fixed to the casing and positioned opposite the tips of the rotor blades, sealing the combustion gas through a predetermined gap, and a cooling passage for cooling the rotor blade sealing members with cooling air introduced into the casing, wherein the flow rate of cooling air flowing through the cooling passage is adjusted according to the engine load of the aircraft.

[0093] 1 Gas turbine engine (aircraft gas turbine engine) 10 Compressor 11 Blades 20 Combustor 30 Turbine 30a Blades 31 High-pressure turbine 31a High-pressure turbine blades 32 Low-pressure turbine 32a Low-pressure turbine blades 32a1 Blade section 32a2 Platform section 32a3 Disk 35 Rotor 35a Rotor cooling passage 40 Supercharger fan 50 Casing 60 Outer casing 70 Shaft 81 Inlet pipe 82 Manifold 83 Outlet section 85 Blade seal member 87 Control valve (airflow adjustment section) 87a Main body 87a1 Main through hole 87a2 Seat surface 87b Valve body 87b1 Sub-through hole 87c Coil spring 87d Surrounding passage 90 Cooling passage A1 Internal cooling air A2 External cooling air A3 Rotor cooling air CA Combustion gas EA External air d Gap d1 Clearance

Claims

1. An aircraft gas turbine engine comprising: a casing; rotor blades housed inside the casing and rotating around the axis of the rotor by combustion gases; rotor blade sealing members fixed to the casing and positioned opposite the tips of the rotor blades, sealing the combustion gases through a predetermined gap; a cooling passage for cooling the rotor blade sealing members with cooling air introduced into the casing; and an air flow rate adjustment unit that adjusts the flow rate of cooling air flowing through the cooling passage according to the engine load of the aircraft.

2. The aircraft gas turbine engine according to claim 1, wherein the airflow adjustment unit adjusts the flow rate of cooling air flowing through the cooling channel during grounding before takeoff, during takeoff, and during cruising.

3. The airflow adjustment unit reduces the flow rate by increasing a predetermined throttling amount during ground operation, which is such that the temperature of the casing reaches an allowable value when the engine load reaches the design maximum load.

4. The airflow adjustment unit increases the flow rate by reducing a predetermined throttling amount during cruising, which is the amount by which the temperature of the casing reaches an allowable value when the engine load becomes the maximum design load.

5. The aircraft gas turbine engine according to claim 3 or 4, wherein the air flow rate adjustment unit is provided with a flow rate adjustment valve whose flow path area is passively changed by differential pressure.

6. An aircraft gas turbine engine according to any one of claims 2 to 4, comprising a rotor cooling channel for cooling the rotor with rotor cooling air introduced into the rotor, wherein the air flow rate adjustment unit increases the flow rate of the rotor cooling air by reducing a predetermined throttling amount during ground operation such that the rotor temperature reaches an allowable value when the engine load becomes the maximum design load.

7. An aircraft gas turbine engine according to any one of claims 2 to 4, comprising a rotor cooling channel for cooling the rotor with rotor cooling air introduced into the rotor, wherein the air flow rate adjustment unit reduces the flow rate of the rotor cooling air by increasing a predetermined throttling amount during cruising so that the rotor temperature reaches an allowable value when the engine load becomes the maximum design load.

8. A method for operating an aircraft gas turbine engine comprising: a casing; rotor blades housed inside the casing and rotating around the axis of a rotor by combustion gases; rotor blade sealing members fixed to the casing and positioned opposite the tips of the rotor blades, sealing the combustion gases through a predetermined gap; and a cooling passage for cooling the rotor blade sealing members with cooling air introduced into the casing, wherein the method for operating an aircraft gas turbine engine adjusts the flow rate of cooling air flowing through the cooling passage according to the engine load of the aircraft.

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