A high speed turbomachine with a ducted combustion chamber and a method for operating the same

Abstract

The application discloses a high-speed turbine with a bypass combustion chamber, and belongs to the field of airplane turbine engines. The high-speed turbine comprises a core combustion chamber located in a core bypass, a bypass combustion chamber in a first outer bypass, and an afterburner in a second outer bypass. The bypass combustion chamber is arranged above the core combustion chamber and is separated from the core combustion chamber. The bypass combustion chamber is arranged below the afterburner and is separated from the afterburner. A core drive fan stage is arranged at the air inlet of the first outer bypass and the core bypass. A mixer and a low-pressure turbine are arranged at the air outlet of the first outer bypass and the core bypass. A high-pressure compressor and a high-pressure turbine are arranged on both sides of the core combustion chamber. The high-speed turbine with the bypass combustion chamber and the operation method thereof are simple in structure, easy to control, do not need additional weight, and can keep the airplane in a better performance under a high Mach number condition.

Description

Technical Field

[0001] This invention relates to the field of aircraft turbine engine technology, and in particular to a high-speed turbine with a ducted combustion chamber and its operating method. Background Technology

[0002] Turbine engines are crucial for achieving wide-speed-range flight in turbine-based combined cycle (TBCC) engines. Limited by combustion chamber temperature, the thrust of conventional turbine engines decreases rapidly with increasing flight speed. This insufficient power prevents combined cycle engines from switching from turbine mode to ramjet mode, thus hindering further speed increases. Furthermore, increasing the speed range of turbine engines means that the minimum speed within the operating envelope of ramjet engines can be increased, reducing the design complexity of subsonic / scramjet engines. Currently, there are three main approaches to improve the thrust performance of turbine engines at high Mach numbers:

[0003] 1) Variable cycle design, represented by engines such as the American J58 and RTA engines, and the Japanese HYPR90 engine, is characterized by the use of numerous geometrically adjustable components, such as compressor guide vanes, turbine guide vane throat area, front and rear bypass ejector inlet areas, and nozzle area. As flight speed increases, the bypass ratio of the engine is altered through the coordinated operation of these adjustable components, allowing more airflow to be allocated to the ramjet bypass duct to increase engine thrust. However, the increased number of adjustable components means that the engine control laws become more complex. To safely and stably meet the aircraft's power requirements, more robust control laws and corresponding controllers need to be designed.

[0004] 2) Jet precooling scheme: Water is injected into the engine intake to address the problem of excessively high total temperature of the incoming air at high Mach numbers through the heat absorption of liquid water evaporation. According to relevant literature, as flight speed increases, the required water injection volume increases from 3% at Mach 2 to 17% at Mach 4. Carrying a large amount of precooling water will sacrifice the engine's thrust-to-weight ratio. Furthermore, excessive water injection may lead to engine instability, thus requiring additional oxidizer, which will further reduce the engine's thrust-to-weight ratio. Regarding the impact on engine performance, the liquid water injected into the intake will not completely evaporate; liquid water entering the engine will affect its normal operation and cause thrust loss.

[0005] 3) The interstage combustion chamber design achieves staged expansion and intermediate reheat by arranging a combustion chamber between the high- and low-pressure turbines, thereby improving cycle power and cycle efficiency, and ultimately increasing engine thrust. According to relevant literature, adding an interstage combustion chamber requires increasing the amount of cooling air needed for the turbines, which can lead to a loss of engine performance. Compared to conventional combustion chambers, the tangential velocity of the turbine exit airflow affects the co-combustion in the interstage combustion chamber, resulting in decreased combustion efficiency. Furthermore, due to the limited axial dimension between the high- and low-pressure turbines, the residence time of the airflow in the interstage combustion chamber is less than the time required for complete combustion, further reducing combustion efficiency. Therefore, designing a compact and efficient interstage combustion chamber is the challenge of this design.

[0006] As can be seen from the above background, there is currently no wide-speed-range turbine engine that is simple in structure, easy to control, and does not sacrifice aircraft performance. Summary of the Invention

[0007] To address the aforementioned problems, this invention provides a high-speed turbine with a ducted combustion chamber and its operation method, which has a simple structure, is easy to control, does not require additional load, and enables the aircraft to maintain better performance under high Mach number conditions.

[0008] To achieve the above objectives, the present invention provides a high-speed turbine with a ducted combustion chamber, comprising a core combustion chamber located in a core duct, a ducted combustion chamber in a first outer duct, and an afterburner in a second outer duct. The ducted combustion chamber is located above and separated from the core combustion chamber, and is located below and separated from the afterburner. A core drive fan stage is provided at the airflow inlet of the first outer duct and the core duct. A mixer and a low-pressure turbine are provided at the airflow outlet of the first outer duct and the core duct. A high-pressure compressor and a high-pressure turbine are respectively provided on both sides of the core combustion chamber.

[0009] Preferably, a fan is installed at the airflow inlet of the high-speed turbine.

[0010] Preferably, the high-pressure compressor is positioned in the airflow inlet direction of the core duct, and the high-pressure turbine is positioned in the airflow outlet direction of the core duct.

[0011] Preferably, an inner nozzle is provided in the direction of the airflow discharge of the low-pressure turbine.

[0012] Preferably, an outer bypass nozzle is provided in the airflow discharge direction of the second outer bypass duct.

[0013] The present invention also provides a method for operating a high-speed turbine with a ducted combustion chamber, which changes accordingly with flight speed, including the following steps:

[0014] Operating steps at low Mach numbers:

[0015] Step 1: Due to the relatively large design of the bypass duct, part of the inlet airflow flows into the core bypass duct, and after being pressurized by the fan, core drive fan stage, and compressor stage, it is heated in the core combustion chamber, and then expanded and discharged through the high and low pressure turbines and the inner nozzle to generate thrust.

[0016] Step two: another part of the inlet airflow is pressurized by the fan and discharged from the outer bypass nozzle to generate thrust; due to the strong flow capacity of the high-pressure compressor, less airflow flows into the first outer bypass at this time, and the mixing and reheating effect of the duct combustion chamber on the core airflow is not obvious.

[0017] As flight speed increases, the high-speed turbine operates as follows:

[0018] Step 1: As flight speed increases, the high-pressure compressor's flow capacity decreases, the core bypass flow gradually decreases, and the proportion of thrust from the inner nozzle decreases.

[0019] Step 2: The second bypass airflow gradually increases due to the increase in the second split ratio, and the effect of the inlet ram press gradually becomes apparent as the speed increases, the total pressure ratio of the second bypass increases, and the thrust of the bypass nozzle increases.

[0020] Step 3: The decrease in the high-pressure compressor's flow capacity means that the airflow into the first bypass increases, the heating amount in the bypass combustion chamber increases, the reheat capacity of the core bypass increases, the core engine cycle work increases, which is conducive to increasing the exhaust velocity of the inner nozzle and maintaining the inner thrust.

[0021] As flight speed increases further, the high-speed turbine operates as follows:

[0022] Step 1: The flight Mach number is particularly high, the suction capacity of the core engine driving the fan stage and high-pressure compressor decreases, the flow rate of the core bypass duct accounts for a very small proportion of the total flow rate of the high-speed turbine, and the inner nozzle provides very little thrust.

[0023] Step two: At this point, the ramjet capability is further enhanced, and the inlet airflow mainly flows into the second bypass. The second bypass operates in a manner similar to a ramjet engine with fan boost, and the bypass thrust dominates the thrust of the high-speed turbine.

[0024] Step 3: At this point, the fan flow rate is high, while the low-pressure turbine flow rate is low. The first bypass airflow is heated by the bypass combustion chamber and then impacts the low-pressure turbine to provide sufficient compression work for the fan.

[0025] The present invention has the following beneficial effects:

[0026] 1. By setting up a ducted combustion chamber, this invention avoids the problems of excessive difficulty in combustion chamber design and low combustion chamber efficiency caused by insufficient size of the interstage combustion chamber;

[0027] 2. The high-speed turbine of the present invention has two nozzles, wherein the inner nozzle provides the main thrust at low Mach numbers and operates in a mode similar to that of a conventional turbofan engine; as the Mach number increases, more flow is allocated to the outer bypass duct, and the thrust ratio of the outer bypass nozzle gradually increases with the increase of Mach number, providing the majority of the thrust to the engine under high Mach number conditions, at which point the operating mode is similar to that of a ramjet engine.

[0028] 3. Under the same design parameters and component characteristics, by comparing with conventional engine configurations, the engine configuration proposed in this patent can improve the thrust of the aircraft and reduce the fuel consumption rate.

[0029] The technical solution of the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Attached Figure Description

[0030] Figure 1 This is a schematic diagram of the internal structure of a high-speed turbine according to an embodiment of the present invention;

[0031] Figure 2 This is a temperature entropy diagram of the thermodynamic cycle process of a conventional configuration turbine engine according to an embodiment of the present invention;

[0032] Figure 3 This is a temperature entropy diagram of the thermodynamic cycle process after airflow mixing and reheating in a high-speed turbine according to an embodiment of the present invention.

[0033] Figure 4 This is a temperature entropy diagram of the first bypass thermodynamic cycle process under high Mach number flight conditions of a high-speed turbine according to an embodiment of the present invention.

[0034] Figure 5 This is a schematic diagram illustrating the thrust variation of the high-speed turbine performance with Mach number according to an embodiment of the present invention.

[0035] Figure 6 This is a schematic diagram illustrating the variation of high-speed turbine performance with unit thrust at Mach number according to an embodiment of the present invention.

[0036] Figure 7 This is a schematic diagram illustrating the variation of the total bypass ratio with Mach number in the performance of a high-speed turbine according to an embodiment of the present invention.

[0037] Figure 8 This is a schematic diagram comparing the thrust variation of a high-speed turbine and a conventional hybrid afterburning turbofan engine with Mach number in an embodiment of the present invention.

[0038] Figure 9 This is a schematic diagram comparing the performance of a high-speed turbine and a conventional hybrid afterburning turbofan engine as a function of Mach number in an embodiment of the present invention.

[0039] Figure 10This is a schematic diagram comparing the fuel consumption rate of a high-speed turbine and a conventional hybrid afterburning turbofan engine as a function of Mach number, according to an embodiment of the present invention.

[0040] In the picture:

[0041] 1. Fan; 2. Core-driven fan stage; 3. High-pressure compressor; 4. Core combustor; 5. High-pressure turbine; 6. Mixer; 7. Low-pressure turbine; 8. Inner tube nozzle; 9. First outer bypass duct; 10. Second outer bypass duct; 11. Afterburner; 12. Bypass combustor; 13. Outer tube nozzle; 14. Core bypass duct. Detailed Implementation

[0042] To make the objectives, technical solutions, and advantages disclosed in the embodiments of the present invention clearer, the embodiments of the present invention will be further described in detail below with reference to the accompanying drawings and examples. It should be understood that the specific embodiments described herein are merely illustrative of the embodiments of the present invention and are not intended to limit the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the embodiments in this application without creative effort are within the scope of protection of this application. Examples of the embodiments are shown in the accompanying drawings, wherein the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout.

[0043] It should be noted that the terms "comprising" and "having," and any variations thereof, are intended to cover non-exclusive inclusion. For example, a process, method, system, product, or server that includes a series of steps or units is not necessarily limited to those steps or units that are explicitly listed, but may include other steps or units that are not explicitly listed or that are inherent to such process, method, product, or device.

[0044] Similar labels and letters in the following figures indicate similar items; therefore, once an item is defined in one figure, it does not need to be further defined and explained in subsequent figures.

[0045] In the description of this invention, it should be noted that the terms "upper," "lower," "inner," "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, or the orientation or positional relationship in which the product of this invention is usually placed when in use. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting this invention.

[0046] In the description of this invention, it should also be noted that, unless otherwise explicitly specified and limited, the terms "set," "install," and "connect" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0047] like Figure 1 As shown, the high-speed turbine with ducted combustion chambers of the present invention has three combustion chambers, including a core combustion chamber 4 located in the core duct 14, a ducted combustion chamber 12 in the first outer duct 9, and an afterburner combustion chamber 11 in the second outer duct 10.

[0048] A fan 1 is installed at the airflow inlet of the high-speed turbine. A ducted combustion chamber 12 is located above and separated from the core combustion chamber 4, and is located below and separated from the afterburner 11. A core drive fan stage 2 is installed at the airflow inlet of the first outer duct 9 and the core duct 14. A mixer 6 and a low-pressure turbine 7 are installed at the airflow outlet of the first outer duct 9 and the core duct 14. A high-pressure compressor 3 and a high-pressure turbine 5 are respectively installed on both sides of the core combustion chamber 4. The high-pressure compressor 3 is positioned in the airflow inlet direction of the core duct 14, and the high-pressure turbine 5 is positioned in the airflow outlet direction of the core duct 14.

[0049] An inner nozzle 8 is provided in the airflow discharge direction of the low-pressure turbine 7. An outer bypass nozzle 13 is provided in the airflow discharge direction of the second outer bypass duct 10. The high-speed turbine of this embodiment has two nozzles, wherein the inner nozzle 8 provides the main thrust at low Mach numbers, and its operating mode is similar to that of a conventional turbofan engine; as the Mach number increases, more flow is allocated to the outer bypass duct, and the thrust proportion of the outer bypass nozzle 13 gradually increases with the increase of Mach number, providing the majority of the thrust to the engine under high Mach number conditions, at which point its operating mode is similar to that of a ramjet engine.

[0050] The high-speed turbine inlet airflow is pressurized by fan 1 and core drive fan stage 2 (CDFS) and then diverted to the first bypass. After being heated by the bypass combustion chamber 12 (IBB), it is mixed with the outlet airflow of the high-pressure turbine 5. The mixing process of the mainstream is similar to the staged expansion and intermediate reheat cycle process of the interstage combustion chamber scheme, so as to improve cycle power and cycle efficiency and increase engine thrust.

[0051] The configuration in this embodiment has a relatively large bypass ratio. As the flight speed increases, the suction capacity of the high-pressure compressor 3 decreases, and the bypass ratio will further increase. At high Mach numbers, most of the airflow will flow into the second bypass. At this time, the bypass afterburner 11 is equivalent to a ramjet combustion chamber. The airflow in this bypass is expanded and discharged through the bypass nozzle 13, providing the main thrust for the engine.

[0052] The present invention discloses a method for operating a high-speed turbine with a ducted combustion chamber, which changes accordingly with flight speed, and includes the following steps:

[0053] Operating steps at low Mach numbers:

[0054] Step 1: Due to the relatively large design of the bypass duct, part of the inlet airflow flows into the core bypass duct, and after being pressurized by the fan, core drive fan stage, and compressor stage, it is heated in the core combustion chamber, and then expanded and discharged through the high and low pressure turbines and the inner nozzle to generate thrust.

[0055] In step two, another portion of the inlet airflow is pressurized by the fan and discharged through the outer bypass nozzle to generate thrust. Due to the strong flow capacity of the high-pressure compressor, less airflow flows into the first outer bypass at this time, and the mixing and reheating effect is not obvious.

[0056] As flight speed increases, the high-speed turbine operates as follows:

[0057] Step 1: As flight speed increases, the high-pressure compressor's flow capacity decreases, the core bypass flow gradually decreases, and the proportion of thrust from the inner nozzle decreases.

[0058] Step 2: The second bypass airflow gradually increases due to the increase in the second split ratio, and the effect of the inlet ram press gradually becomes apparent as the speed increases, the total pressure ratio of the second bypass increases, and the thrust of the bypass nozzle increases.

[0059] Step 3: The decrease in the high-pressure compressor's flow capacity means that the airflow into the first outer bypass increases, the heating amount in the bypass combustion chamber increases, the reheating capacity of the core bypass increases, the core engine cycle work increases, which is conducive to increasing the exhaust velocity of the inner nozzle and maintaining the inner thrust.

[0060] As flight speed increases further, the high-speed turbine operates as follows:

[0061] Step 1: The flight Mach number is particularly high, the suction capacity of the core engine driving the fan stage and high-pressure compressor decreases, the flow rate of the core bypass duct accounts for a very small proportion of the total flow rate of the high-speed turbine, and the inner nozzle provides very little thrust.

[0062] Step two: At this point, the ramjet capability is further enhanced, and the inlet airflow mainly flows into the second bypass. The second bypass operates in a manner similar to a ramjet engine with fan boost, and the bypass thrust dominates the thrust of the high-speed turbine.

[0063] Step 3: At this point, the fan flow rate is high, while the low-pressure turbine flow rate is low. The first bypass airflow is heated by the bypass combustion chamber and then impacts the low-pressure turbine to provide sufficient compression work for the fan.

[0064] Analyzing the thermodynamic cycle of each bypass duct, core bypass duct 14 (innermost) represents the Brayton cycle process of a conventional turbine engine, and its thermo-entropy diagram (TS) is shown below. Figure 2 As shown, there exists an optimal pressure ratio that maximizes cycle work as the Mach number increases. When the flight Mach number continues to increase, the total pressure ratio exceeds the optimal pressure ratio, and the cycle work decreases with increasing Mach number, resulting in a drop in thrust.

[0065] After the ducted combustion chamber 12 is used for mixed combustion, the airflow in the core duct 14 expands after passing through the high-pressure turbine 5 (HPT), and is then mixed and reheated with the high-temperature airflow in the first outer duct 9 in the mixer 6, which can generate more cycle work, such as... Figure 3 As shown.

[0066] Under high Mach number flight conditions, compared to the core bypass duct 14, the first outer bypass duct 9 lacks the boost from the high-pressure compressor 3, resulting in a pressure ratio closer to the optimal pressure ratio and thus greater cycle work. Figure 4 As shown.

[0067] Therefore, from a thermodynamic cycle perspective, through afterburning in the bypass combustion chamber 12 and the first outer bypass duct 9 with a lower pressure ratio, the engine concept proposed in this invention can generate more cycle work under high Mach number conditions, thus maintaining high thrust performance even under high-speed flight conditions. The corresponding design parameters are as follows:

[0068]

[0069]

[0070] By establishing a component-level overall performance simulation model, the variation of high-speed turbine performance with Mach number is calculated as follows: Figure 5 , Figure 6 , Figure 7 As shown, the engine's thrust per unit volume remains high even below Mach 2.5. Furthermore, as the Mach number increases, the bypass ratio increases from 3.28 to 20, meaning the outer bypass nozzle 13 provides the majority of the engine's thrust under high Mach number conditions.

[0071] A conventional hybrid afterburning turbofan engine (MTF) with the same design parameters was selected for performance comparison during acceleration. The engine performance variation with Mach number was obtained as follows: Figure 8 , Figure 9 , Figure 10 As shown.

[0072] As can be seen, under the same modeling method and design parameters, conventional hybrid afterburning turbofan engines are limited by the temperature of the main combustion chamber and can only operate up to Mach 2.5. However, the high-speed turbine (MCE) concept proposed in this invention can provide heating to the low-pressure system through afterburner combustion in the bypass combustion chamber 12, thereby decoupling the low-pressure system (fan 1 and low-pressure turbine 7) from the core engine (CDFS, compressor, core combustion chamber 4 and high-pressure turbine 5), and no longer being limited by the temperature of the core combustion chamber 4, enabling it to operate up to Mach 3.5.

[0073] In terms of engine performance, the high-speed turbine concept proposed in this invention has comparable thrust performance to conventional hybrid afterburning turbofan engines at low Mach numbers, and even better performance at high Mach numbers (thrust is 32% higher than conventional hybrid afterburning turbofan engines at Mach 2.5), and has a significant advantage in fuel consumption performance throughout the acceleration process.

[0074] Therefore, the present invention employs the above-mentioned high-speed turbine with ducted combustion chamber and its operation method, which has a simple structure, is easy to control, does not require additional load, and enables the aircraft to maintain better performance under high Mach number conditions.

[0075] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit them. Although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art should understand that modifications or equivalent substitutions can still be made to the technical solutions of the present invention, and these modifications or equivalent substitutions cannot cause the modified technical solutions to deviate from the spirit and scope of the technical solutions of the present invention.

Claims

1. A high-speed turbine with a ducted combustion chamber, characterized in that: It includes a core duct, a first outer duct, and a second outer duct arranged radially from the inside out. The core duct contains a core combustion chamber, the first outer duct contains a duct combustion chamber, and the second outer duct contains an afterburner. The duct combustion chamber is located above the core combustion chamber and is separated from it. The duct combustion chamber is located below the afterburner and is separated from it. A core drive fan stage is located at the airflow inlet of the first outer duct and the core duct. A mixer and a low-pressure turbine are located at the airflow outlet of the first outer duct and the core duct. A high-pressure compressor and a high-pressure turbine are located on both sides of the core combustion chamber, respectively. A fan is installed at the airflow inlet of the high-speed turbine; An internal nozzle is provided in the direction of the airflow discharge of the low-pressure turbine; An outer bypass nozzle is installed in the airflow discharge direction of the second outer bypass duct; A portion of the airflow from the high-speed turbine inlet flows into the first bypass duct and the core bypass duct via the core drive fan stage, while the other portion enters the second bypass duct.

2. A high-speed turbine with a ducted combustion chamber according to claim 1, characterized in that: The high-pressure compressor is positioned in the direction of airflow in the core duct, while the high-pressure turbine is positioned in the direction of airflow out of the core duct.

3. A method for operating a high-speed turbine with a ducted combustion chamber as described in claim 1 or 2, characterized in that: Includes the following steps: Operating steps of a high-speed turbine at low Mach numbers: Step 1: Part of the inlet airflow flows into the core bypass duct, and after being pressurized by the fan, core drive fan stage, and high-pressure compressor stage, it is heated in the core combustion chamber, and then expanded and discharged through the high-pressure turbine, low-pressure turbine and internal nozzle to generate thrust. Step two: Another portion of the inlet airflow is pressurized by the fan and then discharged through the outer bypass nozzle to generate thrust; As flight speed increases, the high-speed turbine operates as follows: Step 1: As flight speed increases, the high-pressure compressor's flow capacity decreases, the core bypass flow decreases, and the proportion of thrust from the inner nozzle decreases. Step 2: The airflow in the second bypass duct increases due to the increase in the second split ratio, and the effect of the inlet ram press becomes apparent as the velocity increases, thus increasing the total pressure ratio of the second bypass duct and the thrust ratio of the bypass nozzle. Step 3: The high-pressure compressor's flow capacity decreases, the airflow into the first outer bypass increases, the heating amount in the bypass combustion chamber increases, and the cycle work of the high-speed turbine core increases. As flight speed increases further, the high-speed turbine operates as follows: Step 1: As the flight Mach number further increases, the suction capacity of the high-speed turbine core engine driving the fan stage and high-pressure compressor decreases, the core bypass flow accounts for a small proportion of the total flow of the high-speed turbine, and the inner nozzle provides less thrust. Step two, the ramming capability is further enhanced, the inlet airflow flows into the second outer bypass, and the outer bypass thrust dominates the thrust of the high-speed turbine; Step 3: The fan has a large flow rate, while the low-pressure turbine has a small flow rate. The first bypass airflow is heated by the bypass combustion chamber and then impacts the low-pressure turbine to provide compression work for the fan.

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