A method and device for afterburner fuel supply at the right boundary of the flight envelope of an aviation turbofan engine

Abstract

The present application belongs to the field of engine control technology and relates to a method and device for afterburner fueling at the right boundary of the flight envelope of an aviation turbofan engine. The method comprises: step S1, determining the nozzle throat area adjustment range for each operating point at the right boundary of the flight envelope; step S2, discretizing multiple nozzle throat area points; step S3, determining the maximum low-pressure relative speed; step S4, calculating the engine inlet air flow rate and the bypass air flow rate; step S5, determining the air flow rate involved in combustion; step S6, determining the total fuel flow rate; step S7, determining the afterburner fuel flow rate; step S8, calculating the full afterburner thrust; step S9, determining the nozzle throat area point corresponding to the maximum value of the full afterburner thrust, and using the nozzle throat area, afterburner fuel flow rate, and the maximum low-pressure relative speed corresponding to the nozzle throat area point as the afterburner fueling control law for the operating point. The present application reduces the horizontal acceleration time of the aircraft and improves the aircraft's maneuverability.

Description

Technical Field

[0001] The present application belongs to the field of engine control technology, and in particular relates to a method and device for afterburner fuel supply at the right boundary of the flight envelope of an aviation turbofan engine. Background Art

[0002] To reduce aircraft detectability, current turbofan engine designs typically utilize a significant amount of bypass air for nozzle cooling, compared to earlier designs. This bypass air, used for nozzle cooling, cannot participate in afterburner combustion, and the cooling air flow rate is typically related to the overall engine bypass ratio. Without adjustable bypass area, the bypass ratio changes inversely with the nozzle throat area. In other words, increasing the nozzle throat area increases the overall engine bypass ratio, and the amount of nozzle cooling air—the amount of air that cannot participate in afterburner combustion—also increases.

[0003] At the right edge of an aircraft turbofan engine's flight envelope, inlet pressure is high. At this point, increasing the engine's inlet air flow, and thus thrust, is typically achieved by expanding the nozzle throat area. Existing technical solutions only consider the increased air flow's impact on overall thrust, failing to consider how this increased air flow affects the afterburner's air flow. This results in suboptimal overall engine matching and, consequently, suboptimal overall thrust, leading to lower thrust at the right edge of the flight envelope. Furthermore, this low thrust at the right edge of the flight envelope results in longer horizontal acceleration times, reducing aircraft maneuverability. Summary of the Invention

[0004] In order to solve the above problems, the present application provides a method and device for afterburner fuel supply at the right boundary of the flight envelope of an aviation turbofan engine.

[0005] In a first aspect, the present application provides a method for afterburner fueling at the right boundary of the flight envelope of an aviation turbofan engine, mainly comprising:

[0006] Step S1: determining the nozzle throat area adjustment range for each operating point on the right boundary of the flight envelope;

[0007] Step S2, discretizing a plurality of nozzle throat area points within the nozzle throat area adjustment range;

[0008] Step S3, determining the maximum low-pressure relative speed corresponding to each nozzle throat area point;

[0009] Step S4, calculating the engine inlet air flow rate and the bypass air flow rate at each nozzle throat area point according to the nozzle throat area and the maximum low-pressure relative speed;

[0010] Step S5: determining the air flow rate involved in combustion based on the engine inlet air flow rate and the bypass air flow rate;

[0011] Step S6: determining the total fuel flow rate based on the air flow rate involved in combustion;

[0012] Step S7, determining the afterburner fuel flow rate according to the total fuel flow rate;

[0013] Step S8, calculating the full afterburner thrust according to the afterburner fuel flow rate;

[0014] Step S9: Determine the nozzle throat area point corresponding to the maximum value of the full afterburner thrust, and use the maximum values ​​of the nozzle throat area, afterburner fuel flow rate, and low-pressure relative speed corresponding to the nozzle throat area point as the afterburner fuel supply control law for the operating point.

[0015] Preferably, in step S2, the nozzle throat area adjustment range is discretized in steps of 1%.

[0016] Preferably, in step S3, the maximum low-pressure relative speed is determined by the following two conditions:

[0017] The maximum low-pressure relative speed is not greater than the maximum low-pressure relative speed allowed by the engine design at that operating point; and

[0018] The engine turbine inlet temperature calculated from the maximum low-pressure relative speed is not greater than the maximum turbine inlet temperature allowed by the engine design at that operating point.

[0019] Preferably, step S5 further comprises:

[0020] Step S51: Calculate the cooling air flow rate Wb1 that does not participate in combustion based on the ratio X% of the nozzle cooling air to the external bypass air. N *X%, of which Wb1 N is the bypass gas flow rate at the Nth nozzle throat area point;

[0021] Step S52: Determine the air flow rate involved in combustion as Wa1 N -Wb1 N *X%, of which Wa1 N is the engine inlet air flow rate at the Nth nozzle throat area point.

[0022] Preferably, in step S6, the total fuel flow is determined as: (Wa1 N -Wb1 N *X%) / Y; where Y is the air flow rate per kilogram of fuel consumed.

[0023] Preferably, step S7 further includes:

[0024] Step S71, calculating the main fuel flow required for each nozzle throat area point according to the maximum low-pressure relative speed;

[0025] Step S72: Subtract the main fuel flow from the total fuel flow to obtain the afterburner fuel flow.

[0026] A second aspect of the present application provides an afterburner fuel supply device for an aviation turbofan engine at the right boundary of the flight envelope, mainly comprising:

[0027] A nozzle throat area adjustment range determination module is used to determine the nozzle throat area adjustment range for each operating point on the right boundary of the flight envelope;

[0028] A nozzle throat area point discrete module is used to discretize a plurality of nozzle throat area points within the nozzle throat area adjustment range;

[0029] A low-pressure relative speed maximum value determination module is used to determine the low-pressure relative speed maximum value corresponding to each nozzle throat area point;

[0030] The engine inlet air flow and bypass air flow calculation module is used to calculate the engine inlet air flow and bypass air flow at each nozzle throat area point based on the nozzle throat area and the maximum low-pressure relative speed;

[0031] An air flow determination module for participating in combustion, used to determine the air flow participating in combustion based on the engine inlet air flow and the external bypass air flow;

[0032] A total fuel flow determination module, used for determining the total fuel flow according to the air flow involved in combustion;

[0033] An afterburner fuel flow determination module, configured to determine the afterburner fuel flow according to the total fuel flow;

[0034] A full afterburner thrust calculation module is used to calculate the full afterburner thrust based on the afterburner fuel flow rate;

[0035] The afterburner fuel supply control law selection module is used to determine the nozzle throat area point corresponding to the maximum value of the full afterburner thrust, and the nozzle throat area, afterburner fuel flow rate and low-pressure relative speed maximum value corresponding to the nozzle throat area point are used as the afterburner fuel supply control law for this operating point.

[0036] Preferably, in the nozzle throat area point discretization module, the nozzle throat area adjustment range is discretized according to a step size of 1%.

[0037] Preferably, the module for determining the air flow involved in combustion includes:

[0038] The cooling air flow calculation unit that does not participate in combustion is used to calculate the cooling air flow Wb1 that does not participate in combustion based on the ratio X% of the nozzle cooling air to the external bypass air. N *X%, of which Wb1 N is the bypass gas flow rate at the Nth nozzle throat area point;

[0039] The air flow calculation unit involved in combustion is used to determine the air flow involved in combustion as Wa1 N -Wb1 N *X%, of which Wa1 N is the engine inlet air flow rate at the Nth nozzle throat area point.

[0040] Preferably, the afterburner fuel flow determination module includes:

[0041] A main fuel flow calculation unit is used to calculate the main fuel flow required at each nozzle throat area point based on the maximum low-pressure relative speed;

[0042] The afterburner fuel flow calculation unit is used to subtract the main fuel flow from the total fuel flow to obtain the afterburner fuel flow.

[0043] This application achieves the optimal design of the whole aircraft thrust at the right boundary of the flight envelope, and improves the whole aircraft thrust compared to the existing technical solutions. This application reduces the horizontal acceleration time of the aircraft and improves the maneuverability of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Figure 1 This is a flow chart of a preferred embodiment of the afterburner fueling method for the right boundary of the flight envelope of an aviation turbofan engine of the present application. DETAILED DESCRIPTION

[0045] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the implementation of this application will be described in more detail below in conjunction with the drawings in the implementation of this application. In the drawings, the same or similar numbers throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and should not be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in conjunction with the drawings.

[0046] The first aspect of the present application provides a method for afterburner fuel supply at the right boundary of the flight envelope of an aviation turbofan engine, such as Figure 1 As shown, it mainly includes:

[0047] Step S1: determining the nozzle throat area adjustment range for each operating point on the right boundary of the flight envelope;

[0048] Step S2, discretizing a plurality of nozzle throat area points within the nozzle throat area adjustment range;

[0049] Step S3, determining the maximum low-pressure relative speed corresponding to each nozzle throat area point;

[0050] Step S4, calculating the engine inlet air flow rate and the bypass air flow rate at each nozzle throat area point according to the nozzle throat area and the maximum low-pressure relative speed;

[0051] Step S5: determining the air flow rate involved in combustion based on the engine inlet air flow rate and the bypass air flow rate;

[0052] Step S6: determining the total fuel flow rate based on the air flow rate involved in combustion;

[0053] Step S7, determining the afterburner fuel flow rate according to the total fuel flow rate;

[0054] Step S8, calculating the full afterburner thrust according to the afterburner fuel flow rate;

[0055] Step S9: Determine the nozzle throat area point corresponding to the maximum value of the full afterburner thrust, and use the maximum values ​​of the nozzle throat area, afterburner fuel flow rate, and low-pressure relative speed corresponding to the nozzle throat area point as the afterburner fuel supply control law for the operating point.

[0056] In step S1, according to the characteristics of the aviation turbofan engine, when the engine nozzle throat area is gradually enlarged, the overall bypass ratio increases and the nozzle cooling air increases. Assume that the minimum nozzle throat area allowed by the engine design at a certain operating point on the right boundary of the flight is A8 min , the maximum nozzle throat area is A8 max , in step S2, at A8 min -A8 max N nozzle throat area points are discretized within the range of A81, A82...A8 N , where A81 = A8 min , A8 N =A8 max .

[0057] In some optional embodiments, in step S2, the nozzle throat area adjustment range is discretized in steps of 1%. In this embodiment, theoretically, the shorter the step size, the more accurate the final optimal nozzle throat area is. The longer the step size, the faster the calculation speed. After comprehensive consideration, this application discretizes the nozzle throat area adjustment range into 100 nozzle throat area points in steps of 1% for subsequent calculations.

[0058] In step S3, for each nozzle throat area point, the maximum value of the low-pressure relative speed n11, n12...n1 can be given. NIn some optional embodiments, in step S3, the maximum low-pressure relative speed is determined by the following two conditions: the maximum low-pressure relative speed is not greater than the maximum low-pressure relative speed n1 allowed by the engine design at the operating point. max and the engine turbine inlet temperature T41, T42 ... T4 calculated by the maximum value of the low-pressure relative speed N Not greater than the maximum turbine inlet temperature T4 allowed by the engine design at this operating point max .

[0059] Then in step S4, the low pressure relative speed maximum values ​​n11, n12...n1 N And the corresponding nozzle throat areas A81, A82...A8 N Substitute into the whole machine performance calculation program to calculate the inlet air flow Wa11, Wa12...Wa1 N and the external bypass air flow Wb11, Wb12...Wb1 N , the engine inlet air flow rate increases with the increase of nozzle throat area.

[0060] With the engine inlet air flow and the external bypass air flow, the air flow participating in the combustion can be calculated in step S5. In some optional embodiments, step S5 further includes:

[0061] Step S51: Calculate the cooling air flow rate Wb1 that does not participate in combustion based on the ratio X% of the nozzle cooling air to the external bypass air. N *X%, of which Wb1 N is the bypass gas flow rate at the Nth nozzle throat area point;

[0062] Step S52: Determine the air flow rate involved in combustion as Wa1 N -Wb1 N *X%, of which Wa1 N is the engine inlet air flow rate at the Nth nozzle throat area point.

[0063] In this embodiment, as the nozzle throat area increases, the proportion of the entire machine's external air flow increases, resulting in a gradual increase in the proportion of the nozzle cooling air flow to the engine's inlet air flow.

[0064] With the air flow, the corresponding amount of fuel consumed can be determined. In some optional embodiments, in step S6, the total fuel flow is determined as: (Wa1 N -Wb1 N *X%) / Y; where Y is the air flow rate per kilogram of fuel consumed.

[0065] After the total fuel flow is calculated in step S6, the afterburner fuel flow allocated to the afterburner can be determined in step S7. In some optional embodiments, step S7 further includes:

[0066] Step S71, calculating the main fuel flow required for each nozzle throat area point according to the maximum low-pressure relative speed;

[0067] Step S72: Subtract the main fuel flow from the total fuel flow to obtain the afterburner fuel flow.

[0068] In this embodiment, it is assumed that n11, n12, ..., n1 corresponding to the engine N The main fuel flow required by the speed is Wfb1, Wfb2...Wfb N , then the required afterburner fuel flow is (Wa11-Wb11*X%) / Y-Wfb1, (Wa12-Wb12*X%) / Y-Wfb2……(Wa1 N -Wb1 N *X%) / Y-Wfb N .

[0069] Finally, in step S8 and step S9, the afterburner fuel flow is (Wa11-Wb11*X%) / Y-Wfb1, (Wa12-Wb12*X%) / Y-Wfb2...(Wa1 N -Wb1 N *X%) / Y-Wfb N Substituting them into the overall performance calculation program, different full afterburner thrusts are obtained. The A8 area, afterburner fuel supply, and low-pressure relative speed corresponding to the maximum value are selected, which is the optimal solution.

[0070] The design method of the afterburner fuel supply law at the right boundary of the flight envelope of an aviation turbofan engine considering the nozzle cooling air proposed in this application can achieve the optimal design of the whole aircraft thrust at the right boundary of the flight envelope, reduce the horizontal acceleration time of the aircraft, and improve the maneuverability of the aircraft.

[0071] In a second aspect, the present application provides an aviation turbofan engine flight envelope right boundary afterburner fuel supply device corresponding to the above method, mainly comprising:

[0072] A nozzle throat area adjustment range determination module is used to determine the nozzle throat area adjustment range for each operating point on the right boundary of the flight envelope;

[0073] A nozzle throat area point discrete module is used to discretize a plurality of nozzle throat area points within the nozzle throat area adjustment range;

[0074] A low-pressure relative speed maximum value determination module is used to determine the low-pressure relative speed maximum value corresponding to each nozzle throat area point;

[0075] The engine inlet air flow and bypass air flow calculation module is used to calculate the engine inlet air flow and bypass air flow at each nozzle throat area point based on the nozzle throat area and the maximum low-pressure relative speed;

[0076] An air flow determination module for participating in combustion, used to determine the air flow participating in combustion based on the engine inlet air flow and the external bypass air flow;

[0077] A total fuel flow determination module, used for determining the total fuel flow according to the air flow involved in combustion;

[0078] An afterburner fuel flow determination module, configured to determine the afterburner fuel flow according to the total fuel flow;

[0079] A full afterburner thrust calculation module is used to calculate the full afterburner thrust based on the afterburner fuel flow rate;

[0080] The afterburner fuel supply control law selection module is used to determine the nozzle throat area point corresponding to the maximum value of the full afterburner thrust, and the nozzle throat area, afterburner fuel flow rate and low-pressure relative speed maximum value corresponding to the nozzle throat area point are used as the afterburner fuel supply control law for this operating point.

[0081] In some optional implementations, in the nozzle throat area point discretization module, the nozzle throat area adjustment range is discretized according to a step size of 1%.

[0082] In some optional implementations, the module for determining the air flow involved in combustion includes:

[0083] The cooling air flow calculation unit that does not participate in combustion is used to calculate the cooling air flow Wb1 that does not participate in combustion based on the ratio X% of the nozzle cooling air to the external bypass air. N *X%, of which Wb1 N is the bypass gas flow rate at the Nth nozzle throat area point;

[0084] The air flow calculation unit involved in combustion is used to determine the air flow involved in combustion as Wa1 N -Wb1 N *X%, of which Wa1 N is the engine inlet air flow rate at the Nth nozzle throat area point.

[0085] In some optional implementations, the afterburner fuel flow determination module includes:

[0086] A main fuel flow calculation unit is used to calculate the main fuel flow required at each nozzle throat area point based on the maximum low-pressure relative speed;

[0087] The afterburner fuel flow calculation unit is used to subtract the main fuel flow from the total fuel flow to obtain the afterburner fuel flow.

[0088] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A method for afterburner fueling at the right boundary of the flight envelope of an aviation turbofan engine, characterized in that: include: Step S1: determining the nozzle throat area adjustment range for each operating point on the right boundary of the flight envelope; Step S2, discretizing a plurality of nozzle throat area points within the nozzle throat area adjustment range; Step S3, determining the maximum low-pressure relative speed corresponding to each nozzle throat area point; Step S4, calculating the engine inlet air flow rate and the bypass air flow rate at each nozzle throat area point according to the nozzle throat area and the maximum low-pressure relative speed; Step S5: determining the air flow rate involved in combustion based on the engine inlet air flow rate and the bypass air flow rate; Step S6: determining the total fuel flow rate according to the air flow rate involved in combustion; Step S7, determining the afterburner fuel flow rate according to the total fuel flow rate; Step S8, calculating the full afterburner thrust according to the afterburner fuel flow rate; Step S9: Determine the nozzle throat area point corresponding to the maximum value of the full afterburner thrust, and use the maximum values ​​of the nozzle throat area, afterburner fuel flow rate, and low-pressure relative speed corresponding to the nozzle throat area point as the afterburner fuel supply control law for the operating point.

2. The afterburner fueling method for the right boundary of the flight envelope of an aviation turbofan engine according to claim 1, characterized in that: In step S2, the nozzle throat area adjustment range is discretized in steps of 1%.

3. The afterburner fueling method for the right boundary of the flight envelope of an aviation turbofan engine according to claim 1, characterized in that: In step S3, the maximum low-pressure relative speed is determined by the following two conditions: The maximum low-pressure relative speed is not greater than the maximum low-pressure relative speed allowed by the engine design at that operating point; as well as The engine turbine inlet temperature calculated from the maximum low-pressure relative speed is not greater than the maximum turbine inlet temperature allowed by the engine design at that operating point.

4. The afterburner fueling method for the right boundary of the flight envelope of an aviation turbofan engine according to claim 1, characterized in that: Step S5 further comprises: Step S51: Calculate the cooling air flow rate Wb1 that does not participate in combustion based on the ratio X% of the nozzle cooling air to the external bypass air. N *X%, of which Wb1 N is the bypass gas flow rate at the Nth nozzle throat area point; Step S52: Determine the air flow rate involved in combustion as Wa1 N -Wb1 N *X%, of which Wa1 N is the engine inlet air flow rate at the Nth nozzle throat area point.

5. The afterburner fueling method for the right boundary of the flight envelope of an aviation turbofan engine according to claim 4, characterized in that: In step S6, the total fuel flow is determined as: (Wa1 N -Wb1 N *X%) / Y; where Y is the air flow rate per kilogram of fuel consumed.

6. The afterburner fueling method for the right boundary of the flight envelope of an aviation turbofan engine according to claim 5, characterized in that: Step S7 further includes: Step S71, calculating the main fuel flow required for each nozzle throat area point according to the maximum low-pressure relative speed; Step S72: Subtract the main fuel flow from the total fuel flow to obtain the afterburner fuel flow.

7. An afterburner fuel supply device for the right boundary of the flight envelope of an aviation turbofan engine, characterized in that: include: A nozzle throat area adjustment range determination module is used to determine the nozzle throat area adjustment range for each operating point on the right boundary of the flight envelope; A nozzle throat area point discrete module is used to discretize a plurality of nozzle throat area points within the nozzle throat area adjustment range; A low-pressure relative speed maximum value determination module is used to determine the low-pressure relative speed maximum value corresponding to each nozzle throat area point; The engine inlet air flow and bypass air flow calculation module is used to calculate the engine inlet air flow and bypass air flow at each nozzle throat area point based on the nozzle throat area and the maximum low-pressure relative speed; An air flow determination module for participating in combustion, used to determine the air flow participating in combustion based on the engine inlet air flow and the external bypass air flow; A total fuel flow determination module, used for determining the total fuel flow according to the air flow involved in combustion; An afterburner fuel flow determination module, configured to determine the afterburner fuel flow according to the total fuel flow; A full afterburner thrust calculation module is used to calculate the full afterburner thrust based on the afterburner fuel flow rate; The afterburner fuel supply control law selection module is used to determine the nozzle throat area point corresponding to the maximum value of the full afterburner thrust, and the nozzle throat area, afterburner fuel flow rate and low-pressure relative speed maximum value corresponding to the nozzle throat area point are used as the afterburner fuel supply control law for this operating point.

8. The afterburner fuel supply device for the right boundary of the flight envelope of an aviation turbofan engine according to claim 7, characterized in that: In the nozzle throat area point discretization module, the nozzle throat area adjustment range is discretized according to a step size of 1%.

9. The afterburner fuel supply device for the right boundary of the flight envelope of an aviation turbofan engine according to claim 7, characterized in that: The air flow determination module involved in combustion includes: The cooling air flow calculation unit that does not participate in combustion is used to calculate the cooling air flow Wb1 that does not participate in combustion based on the ratio X% of the nozzle cooling air to the external bypass air. N *X%, of which Wb1 N is the bypass gas flow rate at the Nth nozzle throat area point; The air flow calculation unit involved in combustion is used to determine the air flow involved in combustion as Wa1 N -Wb1 N *X%, of which Wa1 N is the engine inlet air flow rate at the Nth nozzle throat area point.

10. The afterburner fuel supply device for the right boundary of the flight envelope of an aviation turbofan engine according to claim 7, characterized in that: The afterburner fuel flow determination module includes: A main fuel flow calculation unit is used to calculate the main fuel flow required at each nozzle throat area point based on the maximum low-pressure relative speed; The afterburner fuel flow calculation unit is used to subtract the main fuel flow from the total fuel flow to obtain the afterburner fuel flow.

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