An aeroengine adopting a knock combustion chamber and a performance design method thereof

Abstract

The application belongs to the technical field of aero-engine design, and particularly relates to an aero-engine adopting a knock combustion chamber, comprising: a booster stage is additionally arranged behind a compressor, the booster stage is located inside the knock combustion chamber and is in parallel with the knock combustion chamber, the airflow pressure after being boosted by the booster stage is equivalent to the airflow pressure at the outlet of the knock combustion chamber, and the high-pressure turbine is cooled by using pipeline bleed air after the booster stage; a front mixing chamber is additionally arranged behind the knock combustion chamber and the booster stage, the front mixing chamber is located before the high-pressure turbine, the airflow flowing out from the knock combustion chamber and the booster stage enters the front mixing chamber to occur mixing, and the airflow at the outlet of the knock combustion chamber is processed into a quasi-steady and constant state, in addition, the application also relates to a performance design method of the aero-engine adopting the knock combustion chamber.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aero-engine design, and particularly relates to an aero-engine adopting a detonation combustion chamber and a performance design method thereof. BACKGROUND

[0002] With the development of aero-engine technology, the turbofan engine technology based on the constant pressure combustion cycle has been very mature, and the aerodynamic performance and structural performance of the turbofan engine gradually approach the limit. Only by seeking a new cycle or combustion mode can a significant breakthrough in performance be achieved.

[0003] The introduction of detonation combustion into the main combustion chamber of an aero-engine and the adoption of a detonation combustion chamber can significantly improve the cycle efficiency of the aero-engine. Based on the self-pressurization characteristics of detonation combustion, the number of compressor stages and the pressure ratio thereof can be reduced, the main combustion chamber inlet temperature can be reduced, more heat can be added to the main combustion chamber, and the energy extracted from the high-temperature gas by the turbine is also reduced due to the reduced pressure ratio of the compressor. The saved gas energy can be expanded by the nozzle at high speed to produce more thrust, thereby improving the thrust of the aero-engine and reducing the specific fuel consumption.

[0004] However, the current aero-engine designed with a detonation combustion chamber has the following defects:

[0005] 1) The cooling gas of the high-pressure turbine comes from the compressor. The detonation combustion chamber is located after the compressor and before the high-pressure turbine. Due to the self-pressurization characteristics of detonation combustion, the pressure of the detonation combustion chamber outlet airflow is higher than that of the cooling gas of the high-pressure turbine, and it is difficult to effectively cool the high-pressure turbine;

[0006] 2) Detonation combustion is supersonic combustion, and the detonation combustion chamber outlet airflow is a supersonic airflow with a Mach number greater than 1, and the airflow state is unsteady. The unsteady supersonic airflow directly impacts the subsequent high-pressure turbine, which usually causes a sharp decrease in the efficiency of the high-pressure turbine;

[0007] 3) There is a lack of reliable performance design method. When designing, the existing aero-engine performance calculation process is still used, resulting in a large parameter matching error of the aero-engine.

[0008] In view of the above technical defects, the present application is proposed.

[0009] It should be noted that the disclosure of the above background art content is only used to assist in understanding the inventive concept and technical solutions of the present application, and it does not necessarily belong to the prior art of the present application. In the absence of explicit evidence that the above content has been disclosed on the filing date of the present application, the above background art should not be used to evaluate the novelty and inventiveness of the present application. SUMMARY

[0010] The object of the present application is to provide an aero-engine with a detonation combustion chamber and a performance design method thereof to overcome or alleviate at least one aspect of the known technical defects.

[0011] The technical solution of the present application is:

[0012] In one aspect, the present application provides an aero-engine with a detonation combustion chamber, comprising:

[0013] A booster stage is added after the compressor, the booster stage is located inside the detonation combustion chamber and is connected in parallel with the detonation combustion chamber, the pressure of the airflow after the booster stage is equal to the pressure of the airflow at the outlet of the detonation combustion chamber, and the airflow after the booster stage is used to cool the high-pressure turbine through a pipeline;

[0014] A premixing chamber is added after the detonation combustion chamber and the booster stage, the premixing chamber is located before the high-pressure turbine, and the airflow from the detonation combustion chamber and the booster stage enters the premixing chamber to mix and treat the airflow at the outlet of the detonation combustion chamber into a quasi-steady and constant state.

[0015] According to at least one embodiment of the present application, in the aero-engine with a detonation combustion chamber described above, the pressure of the airflow after the booster stage is slightly greater than the pressure of the airflow at the outlet of the detonation combustion chamber.

[0016] In another aspect, the present application provides a performance design method for an aero-engine with a detonation combustion chamber, comprising:

[0017] Step one, set flight conditions, including far-field atmospheric pressure P0, far-field atmospheric temperature T0, far-field atmospheric speed Ma, and far-field atmospheric height H;

[0018] Step two, set the fan duct ratio B and the fan pressure ratio π;

[0019] Step three, based on P0, T0, Ma, H, B, and π, calculate the fan outer bypass outlet pressure P22, the fan outer bypass outlet temperature T22, the fan inner bypass outlet pressure P2.5, and the fan inner bypass outlet temperature T2.5;

[0020] Step four, based on P2.5 and T2.5, calculate the compressor outlet pressure P3 and the compressor outlet temperature T3;

[0021] Step five, set the detonation combustion chamber fuel-air ratio far, based on P3 and T3, calculate the detonation combustion chamber outlet pressure P3.5 II and the detonation combustion chamber outlet temperature T3.5 II;

[0022] Step six, set the pressure ratio π of the booster stage, based on P3 and T3, calculate the outlet pressure P3.5 of the booster stage and the outlet temperature T3.5 of the booster stage;

[0023] Step seven, judging whether P3.5II and P3.5 are equal, if not, returning to step six to reset π, if yes, proceeding to step eight;

[0024] Step eight, based on P3.5II, T3.5II, P3.5, T3.5, calculating the front mixing chamber outlet pressure P4 and the front mixing chamber outlet temperature T4;

[0025] Step nine, judging whether T4 and the high pressure turbine front total temperature design value T4des are equal, if not, returning to step five to reset far, if yes, proceeding to step ten;

[0026] Step ten, based on P4, T4, calculating the high pressure turbine outlet pressure P4.5 and the high pressure turbine outlet temperature T4.5;

[0027] Step eleven, based on P4.5, T4.5, calculating the low pressure turbine outlet pressure P5 and the low pressure turbine outlet temperature T5;

[0028] Step twelve, based on P22, T22, calculating the engine outer bypass outlet pressure P5II and the engine outer bypass outlet temperature T5II;

[0029] Step thirteen, judging whether P5II and P5 are equal, if not, returning to step two to reset B and / or π, if yes, proceeding to step fourteen;

[0030] Step fourteen, based on P5II, T5II, P5, T5, calculating the rear mixing chamber outlet pressure P6 and the rear mixing chamber outlet temperature T6;

[0031] Step fifteen, based on P6, T6, calculating the afterburner outlet pressure P7 and the afterburner outlet temperature T7;

[0032] Step sixteen, based on P7, T7, calculating the nozzle outlet pressure P9 and the nozzle outlet temperature T9;

[0033] Step seventeen, based on the calculated parameters, calculating the aeroengine thrust Fn and the specific fuel consumption sfc. BRIEF DESCRIPTION OF DRAWINGS

[0034] Figure 1 is a schematic diagram of an aeroengine provided by an embodiment of the present application and adopting a detonation combustion chamber;

[0035] Figure 2 is a schematic diagram of an aeroengine performance design method provided by an embodiment of the present application and adopting a detonation combustion chamber.

[0036] For better illustrating the embodiments, some components in the drawings can be omitted, enlarged or reduced, and the size of the actual product is not represented. In addition, the drawings are only used for illustrative description and cannot be understood as limitation to the present application. DETAILED DESCRIPTION

[0037] In order to make the technical solutions of the present application and the advantages thereof clearer, the technical solutions of the present application will be further clearly and completely described in the following with reference to the drawings. It should be understood that the specific embodiments described herein are only some embodiments of the present application, which are used to explain the present application but not to limit the present application. It should be noted that, for the purpose of description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design. In the case of no conflict, the embodiments in the present application and the technical features in the embodiments can be combined to obtain new embodiments.

[0038] In addition, unless otherwise defined, the technical terms or scientific terms used in the present application description should be the general meaning understood by the general technical personnel in the field to which the present application belongs. The words indicating the direction or position relationship such as "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer" and the like used in the present application description are only used to indicate the relative direction or position relationship, and not to imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation, and the relative position relationship can also change accordingly when the absolute position of the described object changes, therefore, it cannot be understood as a limitation to the present application. The terms "first", "second", "third" and the like used in the present application description are only for the purpose of description, to distinguish different components, and cannot be understood as indicating or implying relative importance. The terms "one", "an" or "the" and the like used in the present application description should not be understood as an absolute limitation on the quantity, but should be understood as the existence of at least one. The terms "include" or "contain" and the like used in the present application description mean that the elements or objects appearing before the word cover the elements or objects listed after the word and their equivalents, and do not exclude other elements or objects.

[0039] In addition, it should be further noted that, unless otherwise explicitly specified and limited, the terms "mount", "connect", "connect" and the like used in the present application description should be understood in a broad sense, for example, the connection can be fixed connection, or detachable connection, or integral connection; can be mechanical connection, or electrical connection; can be directly connected, or indirectly connected through intermediate medium, or the connection between two elements, and the person skilled in the art can understand the specific meaning of the terms in the present application according to the specific circumstances.

[0040] The present application will be further described in detail below with reference to the drawings. Figures 1-2 The present application will be further described in detail below with reference to the drawings.

[0041] For the aero-engine using the detonation combustion chamber, the pressure of the detonation combustion chamber outlet section is higher than the pressure of the cooling gas of the high-pressure turbine, and it is difficult to effectively cool the high-pressure turbine, the aero-engine using the detonation combustion chamber provided by the application adds a coaxial booster stage after the compressor, the booster stage is located inside the detonation combustion chamber and is parallel to the detonation combustion chamber, the airflow after the compressor is divided into two parts, one part enters the detonation combustion chamber for combustion, and the other part enters the booster stage for pressure boosting, the pressure of the airflow after the pressure boosting of the booster stage can be designed to be equivalent to the pressure of the airflow at the outlet of the detonation combustion chamber, and the high-pressure turbine is cooled by using the pipeline bleed air after the booster stage, so that effective cooling of the high-pressure turbine can be realized, and the distribution ratio of the two parts of the airflow after the compressor can be determined according to the actual demand of turbine cooling.

[0042] For the aero-engine using the detonation combustion chamber, the unsteady supersonic airflow at the outlet of the detonation combustion chamber directly impacts the high-pressure turbine, resulting in a sharp decline in the efficiency of the high-pressure turbine, the aero-engine using the detonation combustion chamber provided by the application adds a premixing chamber after the detonation combustion chamber and the booster stage, the premixing chamber is located before the high-pressure turbine, and the airflow flowing out of the detonation combustion chamber and the booster stage enters the premixing chamber for mixing, so that the airflow at the outlet of the detonation combustion chamber can be processed into a quasi-steady and constant state, thereby reducing the adverse effects on the efficiency of the high-pressure turbine, and the pressure ratio of the booster stage is determined according to the pressure balance principle at the inlet of the premixing chamber, and the pressure of the airflow at the outlet of the booster stage can be designed to be slightly greater than the pressure of the airflow at the outlet of the detonation combustion chamber, specifically, it can be 0.01MPa greater.

[0043] The aero-engine using the detonation combustion chamber provided by the application adds a booster stage inside the detonation combustion chamber after the compressor, provides cooling gas for the high-pressure turbine, and ensures effective cooling of the high-pressure turbine, and adds a premixing chamber before the high-pressure turbine after the detonation combustion chamber and the booster stage, and uses the premixing chamber for flow regulation to ensure the efficiency of the high-pressure turbine, while keeping the main configuration of the aero-engine unchanged.

[0044] In the above-mentioned aero-engine using the detonation combustion chamber, to distinguish the mixing chamber located after the low-pressure turbine and the outer duct and before the afterburner from the premixing chamber, it is referred to as a post-mixing chamber.

[0045] The above-mentioned aero-engine using the detonation combustion chamber is mainly the design of the overall configuration, the connection and cooperation between the components can be designed according to the existing technology, and detailed description will not be repeated.

[0046] Figure 1 The schematic diagram of the aero-engine using the detonation combustion chamber, each section is marked with a reference sign to indicate the performance design section, and the specific meaning is as follows:

[0047]

[0048]

[0049] To solve the problem that the overall aero-engine with detonation combustion chamber lacks reliable performance design method, resulting in large parameter matching error of the aero-engine, the application provides an aero-engine performance design method with detonation combustion chamber, which specifically comprises the following steps.

[0050] Step one, set flight conditions, including remote atmospheric pressure P0, remote atmospheric temperature T0, remote atmospheric speed Ma, remote atmospheric height H.

[0051] Step two, set fan duct ratio B and fan pressure ratio π.

[0052] Step three, based on P0, T0, Ma, H, B and π, the fan calculation module is used to calculate the fan outer duct outlet pressure P22, the fan outer duct outlet temperature T22, the fan inner duct outlet pressure P2.5 and the fan inner duct outlet temperature T2.5.

[0053] Step four, based on P2.5 and T2.5, the compressor calculation module is used to calculate the compressor outlet pressure P3 and the compressor outlet temperature T3.

[0054] Step five, set the detonation combustion chamber oil-gas ratio far, based on P3 and T3, the detonation combustion chamber calculation module is used to calculate the detonation combustion chamber outlet pressure P3.5II and the detonation combustion chamber outlet temperature T3.5II.

[0055] Step six, set the pressure ratio π of the pressure booster, based on P3 and T3, the pressure booster calculation module is used to calculate the pressure booster outlet pressure P3.5 and the pressure booster outlet temperature T3.5.

[0056] Step seven, judge whether the values of P3.5II and P3.5 are equal, if not, return to step six and reset π, if yes, proceed to step eight.

[0057] Step eight, based on P3.5II, T3.5II, P3.5 and T3.5, the pre-mixing chamber calculation module is used to calculate the pre-mixing chamber outlet pressure P4 and the pre-mixing chamber outlet temperature T4.

[0058] Step nine, judge whether the value of T4 is equal to the high-pressure turbine total temperature design value T4des, if not, return to step five and reset far, if yes, proceed to step ten.

[0059] Step ten, based on P4 and T4, the high-pressure turbine calculation module is used to calculate the high-pressure turbine outlet pressure P4.5 and the high-pressure turbine outlet temperature T4.5.

[0060] Step eleven, based on P4.5, T4.5, the low pressure turbine calculation module is used to calculate the low pressure turbine outlet pressure P5 and the low pressure turbine outlet temperature T5.

[0061] Step twelve, based on P22, T22, the outer bypass calculation module is used to calculate the engine outer bypass outlet pressure P5II and the engine outer bypass outlet temperature T5II.

[0062] Step thirteen, it is judged whether P5II and P5 are equal, if not, return to step two, reset B and / or π, if yes, proceed to step fourteen.

[0063] Step fourteen, based on P5II, T5II, P5, T5, the after mixing chamber calculation module is used to calculate the after mixing chamber outlet pressure P6 and the after mixing chamber outlet temperature T6.

[0064] Step fifteen, based on P6, T6, the afterburner calculation module is used to calculate the afterburner outlet pressure P7 and the afterburner outlet temperature T7.

[0065] Step sixteen, based on P7, T7, the nozzle calculation module is used to calculate the nozzle outlet pressure P9 and the nozzle outlet temperature T9.

[0066] Step seventeen, based on P9, T9 and the calculated parameters, the aero-engine performance calculation module is used to calculate the aero-engine thrust Fn and the specific fuel consumption sfc.

[0067] The aero-engine performance design method disclosed above using the detonation combustion chamber is designed to use the calculation modules of each component of the aero-engine to separately calculate each component, and through simple iteration, the aero-engine thrust Fn and the specific fuel consumption sfc are quickly calculated, which is convenient and efficient.

[0068] The embodiments in the specification are described in a progressive manner, and each embodiment focuses on the difference from other embodiments. The same or similar parts between the embodiments can be referred to each other.

[0069] The technical solution of the present application has been described in combination with the preferred embodiments shown in the drawings. It should be understood by those skilled in the art that the protection scope of the present application is obviously not limited to these specific embodiments. Without deviating from the principles of the present application, those skilled in the art can make equivalent changes or replacements to the related technical features, and the technical solutions after the changes or replacements will fall within the protection scope of the present application.

Claims

1. A performance design method for an aero-engine employing a detonation combustor, wherein the aero-engine employing a detonation combustor has an additional booster stage after the compressor, the booster stage is located inside the detonation combustor and connected in parallel with the detonation combustor, the booster stage pressurizes the airflow pressure to be equivalent to the outlet airflow pressure of the detonation combustor, and the high-pressure turbine is cooled by bleed air from the booster stage via pipeline. A pre-mixing chamber is added after the detonation combustion chamber and the turbocharger stage. The pre-mixing chamber is located before the high-pressure turbine. The airflow from the detonation combustion chamber and the turbocharger stage enters the pre-mixing chamber and is mixed, so that the airflow at the outlet of the detonation combustion chamber is processed into a near steady state or a stable state. The performance design method for an aero-engine employing a detonation combustion chamber is characterized in that, include: Step 1: Set flight conditions, including distant atmospheric pressure P0, distant atmospheric temperature T0, distant atmospheric speed Ma, and distant atmospheric altitude H; Step 2: Set the fan bypass ratio B and the fan boost ratio π; Step 3: Based on P0, T0, Ma, H, B, and π, calculate the fan outer duct outlet pressure P22, fan outer duct outlet temperature T22, fan inner duct outlet pressure P2.5, and fan inner duct outlet temperature T2.

5. Step 4: Based on P2.5 and T2.5, calculate the compressor outlet pressure P3 and compressor outlet temperature T3. Step 5: Set the air-fuel ratio far in the detonation combustion chamber. Based on P3 and T3, calculate the outlet pressure P3.5Ⅱ and the outlet temperature T3.5Ⅱ of the detonation combustion chamber. Step 6: Set the boost ratio π of the boost stage. Based on P3 and T3, calculate the boost stage outlet pressure P3.5 and boost stage outlet temperature T3.

5. Step 7: Determine whether the values ​​of P3.5Ⅱ and P3.5 are equal. If not, return to step 6 and reset π. If yes, proceed to step 8. Step 8: Based on P3.5Ⅱ, T3.5Ⅱ, P3.5, and T3.5, calculate the pre-mixing chamber outlet pressure P4 and the pre-mixing chamber outlet temperature T4. Step 9: Determine whether T4 and the total temperature design value T4des before the high-pressure turbine are equal. If not, return to step 5 and reset far. If yes, proceed to step 10. Step 10: Based on P4 and T4, calculate the high-pressure turbine outlet pressure P4.5 and the high-pressure turbine outlet temperature T4.

5. Step 11: Based on P4.5 and T4.5, calculate the low-pressure turbine outlet pressure P5 and the low-pressure turbine outlet temperature T5. Step 12: Based on P22 and T22, calculate the engine bypass outlet pressure P5Ⅱ and the engine bypass outlet temperature T5Ⅱ. Step 13: Determine whether P5Ⅱ and P5 are equal. If not, return to step 2 and reset B and / or π. If yes, proceed to step 14. Step 14: Based on P5Ⅱ, T5Ⅱ, P5, and T5, calculate the outlet pressure P6 and outlet temperature T6 of the after-mixing chamber. Step 15: Based on P6 and T6, calculate the afterburner outlet pressure P7 and afterburner outlet temperature T7. Step 16: Based on P7 and T7, calculate the nozzle outlet pressure P9 and nozzle outlet temperature T9. Step 17: Based on the calculated parameters, calculate the aero-engine thrust Fn and fuel consumption rate sfc.

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