Aircraft airspeed calculation method, airspeed calculation system, airspeed calculation device, storage medium and program product

Abstract

The invention discloses an airspeed calculation method of an aircraft, which can improve the precision of airspeed calculation in a whole flight stage. The airspeed calculation method comprises the following steps: calculating a first dynamic pressure at least according to the wing area, the lift coefficient and the lift of the aircraft; second dynamic pressure is calculated according to the engine operation parameters of the aircraft; calculating a first static pressure according to the flight height of the aircraft and an atmospheric statics equation; calculating a second static pressure according to a pressure sensor arranged on the engine; fusing the first dynamic pressure and the second dynamic pressure to form a fused dynamic pressure, and fusing the first static pressure and the second static pressure to form a fused static pressure; and calculating the airspeed of the aircraft according to the fusion dynamic pressure and the fusion static pressure.

Description

Technical Field

[0001] The present invention relates to the field of aerospace, and in particular to an airspeed calculation method, an airspeed calculation system, a device, a storage medium and a program product of an aircraft. Background Art

[0002] At present, the air data system on the aircraft calculates parameters such as airspeed mainly based on the traditional sensor measurement method. When the air data probe (or sensor) encounters ice crystals or volcanic ash, its measurement accuracy is affected and it is impossible to obtain a relatively accurate airspeed. Therefore, a solution is proposed to calculate the airspeed based on the dynamic pressure calculated by aerodynamics and the static pressure calculated based on GPS altitude, so as to avoid the above situation.

[0003] However, the inventors of the present application have discovered that the solution of calculating the airspeed based on the dynamic pressure obtained by aerodynamic calculations in combination with the above-mentioned static pressure has a relatively single data source and may have a reduced accuracy under certain flight conditions. Summary of the invention

[0004] Technical problem to be solved by the invention

[0005] The present invention is formed in consideration of the above situation, and its purpose is to provide an airspeed calculation method for an aircraft, which can improve the accuracy of airspeed calculation in the full flight phase. On this basis, the present invention also provides an airspeed calculation system, a computer device, a computer-readable storage medium and a computer program product corresponding to the airspeed calculation method.

[0006] Technical solutions adopted to solve technical problems

[0007] The first technical solution of the present invention provides a method for calculating the airspeed of an aircraft, characterized in that it includes:

[0008] Calculating a first dynamic pressure based on at least a wing area, a lift coefficient, and a lift of the aircraft;

[0009] calculating a second dynamic pressure based on engine operating parameters of the aircraft;

[0010] Calculating a first static pressure according to the flight altitude of the aircraft and an atmospheric statics equation;

[0011] calculating a second static pressure according to a pressure sensor provided in the engine;

[0012] fusing the first dynamic pressure and the second dynamic pressure to form a fused dynamic pressure, and fusing the first static pressure and the second static pressure to form a fused static pressure; and

[0013] The airspeed of the aircraft is calculated based on the fusion dynamic pressure and the fusion static pressure.

[0014] Based on the first technical solution, in the second technical solution of the present invention, the lift coefficient is determined according to the angle of attack of the aircraft, the lift coefficient increment and the initial lift coefficient, the lift coefficient increment is the lift coefficient increment caused by the angle change of the high lift device of the aircraft, and the initial lift coefficient is the lift coefficient when the angle of attack is zero.

[0015] On the basis of the second technical solution, in the third technical solution of the present invention, the lift coefficient is determined by the following mathematical formula:

[0016] C L =C L0 +ΔC L +C Lα ×α

[0017] Among them, C L represents the lift coefficient, α represents the angle of attack, C L0 represents the initial lift coefficient, ΔC L represents the lift coefficient increment, C Lα is the slope when the angle of attack is the independent variable and the lift coefficient is the dependent variable.

[0018] Based on the first technical solution, in the fourth technical solution of the present invention, the engine operating parameters include the fan inlet static pressure, the intake duct inlet static pressure and the low-pressure rotor speed of the engine.

[0019] On the basis of the first technical solution, in the fifth technical solution of the present invention, the atmospheric statics equation is expressed by the following mathematical formula:

[0020]

[0021] Wherein, P represents the first static pressure, dP represents the differential change of the first static pressure, g represents the angular velocity of gravity, R represents the universal gas constant, T represents the absolute temperature, and dh is the differential change of the flight altitude.

[0022] Based on the first technical solution, in a sixth technical solution of the present invention, the lift is determined by the angle of attack, lift-to-drag ratio, mass and normal load coefficient of the aircraft.

[0023] On the basis of the second technical solution, in a seventh technical solution of the present invention, the lift coefficient increment is determined according to at least one of a change in a flap angle and a change in a slat angle of the aircraft.

[0024] Effects of the Invention

[0025] According to the airspeed calculation method described in the present application, the first dynamic pressure calculated based on aerodynamics and the second dynamic pressure calculated based on engine operating parameters are fused to form a fused dynamic pressure, and at the same time, the first static pressure calculated based on the flight altitude and atmospheric statics principles and the second static pressure calculated based on the sensing value of the engine pressure sensor are fused to form a fused static pressure. That is to say, in the airspeed calculation method described in the first technical solution, the dynamic pressure (static pressure) calculated in two different ways are fused, and the airspeed is calculated based on the fused dynamic pressure and the fused static pressure. In this way, the problem of a single data source as the basis for calculating the airspeed can be avoided, thereby improving the calculation accuracy of the airspeed during the entire flight phase. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 It is a main flow chart showing the method for calculating the airspeed of an aircraft according to an embodiment of the present application.

[0027] Figure 2 It is a functional structure diagram of an airspeed calculation system according to one embodiment of the present application.

[0028] Figure 3 The present invention is a schematic diagram of the structure of a computer system for realizing an airspeed calculation method of an aircraft. DETAILED DESCRIPTION

[0029] Below, refer to Figure 1 , the principle and specific process of the aircraft airspeed calculation method of one embodiment of the present application are explained.

[0030] like Figure 1 As shown, the airspeed calculation method of an aircraft according to one embodiment of the present application is mainly realized by the cooperation of four parts, which are an aerodynamic model part S1, an engine model part S2, a GNSS altitude back-calculation static pressure part S3 and an airspeed synthesis part S4.

[0031] The aerodynamic model part S1 is used to calculate the first dynamic pressure Pt1 based on the parameters such as the wing area of ​​the aircraft (or the wing reference area), the lift coefficient and the lift. In the aerodynamic model part S1, the sensing data representing the angle of attack information is first obtained from the AOA sensor, i.e., the angle of attack sensor. The angle of attack (Angle of Attack), also known as the angle of attack, is an important concept in aeronautics, which refers to the angle between a flying object (such as the wing of an aircraft) and the relative airflow direction. The size of the angle of attack directly affects the lift, drag and stability of the aircraft. Taking into account that the above sensing data may contain noise, the sensing data is filtered to remove the noise therein, so as to calculate the specific angle of attack value. After obtaining the specific angle of attack value, the processing is bifurcated into one path for calculating the lift and one path for calculating the lift coefficient.

[0032] In the process of calculating lift, the angle of attack value is multiplied by the lift-to-drag ratio of the aircraft. The lift-to-drag ratio of an aircraft is an important parameter to measure the aerodynamic efficiency of an aircraft. It represents the ratio of lift to drag generated by the aircraft during flight. The higher the lift-to-drag ratio, the less thrust the aircraft needs to overcome drag at the same speed, thus saving more fuel. Generally speaking, the lift-to-drag ratio of an aircraft can be calculated through wind tunnel experiments and based on the aircraft's Mach number. Generally speaking, the basic steps for obtaining the lift-to-drag ratio through wind tunnel experiments are:

[0033]

Wind Tunnel Preparation

[0034] First, a wind tunnel is needed that can generate stable and controllable airflow to simulate the flight conditions of an aircraft at different speeds.

[0035]

Model making

[0036] Make one or more aircraft models that are scaled down to the actual aircraft and replicate the aircraft's aerodynamic characteristics as accurately as possible.

[0037]

Mach number setting

[0038] According to the purpose of the experiment, different Mach numbers are set, that is, the ratio of the speed of the aircraft relative to the surrounding air to the speed of sound. The Mach number is an important parameter to describe the flight speed of the aircraft, which affects the aerodynamic characteristics of the aircraft.

[0039] [Experimental conduct]

[0040] In the wind tunnel, the aircraft model is placed in the test section and adjusted to a specific angle of attack. Then, the wind speed is gradually increased and the Mach number is changed to simulate the conditions of the aircraft at different flight speeds.

[0041]

Data collection

[0042] Various sensors and measuring devices are used to collect data including lift, drag, pressure distribution, etc. This data will be used to calculate the lift-to-drag ratio.

[0043] [Lift-to-Drag Ratio Calculation] The lift-to-drag ratio L / D is calculated based on the collected lift and drag data. This calculation is usually performed at multiple angles of attack and Mach numbers to obtain a series of lift-to-drag ratio data.

[0044]

Data Analysis

[0045] Analyze the experimental data to find out how the lift-to-drag ratio changes with angle of attack and Mach number. This helps determine the best flight conditions for the aircraft and the areas that need to be optimized in the design.

[0046]

Result Verification

[0047] The accuracy of the wind tunnel experiment is verified by comparing it with other experimental results or computational fluid dynamics (CFD) simulation results.

[0048] The weight of the airplane is multiplied by the airplane's normal load factor before, after, or simultaneously with the multiplication of the angle of attack by the lift-to-drag ratio. The lift is then calculated by dividing the product of the weight and the normal load factor by the product of the angle of attack and the lift-to-drag ratio.

[0049] On the other hand, in one method of calculating the lift coefficient, the lift coefficient can be determined by, for example, the following mathematical formula:

[0050] C L =C L0 +ΔC L +C Lα ×α

[0051] Among them, C L represents the lift coefficient, α represents the angle of attack, C L0 represents the initial lift coefficient, ΔC L represents the lift coefficient increment, C Lα It is the slope when the angle of attack is the independent variable and the lift coefficient is the dependent variable, and is obtained through experimental testing.

[0052] After the lift coefficient is determined, the lift coefficient is multiplied by the wing reference area of ​​the aircraft. After the multiplication is completed, when the lift of the aircraft has been calculated in the above manner, the lift is divided by the product of the lift coefficient and the wing reference area, thereby calculating the first dynamic pressure Pt1 of the aircraft.

[0053] It should be noted that, in this embodiment, the initial lift coefficient and the lift coefficient increment are set by using the flap angle and the slat angle, but the invention is not limited thereto. The angles and angle changes of other high lift devices other than flaps and slats may also be used for setting. That is, the lift coefficient model may also be established by other methods.

[0054] The engine model part S2 is used to calculate the second dynamic pressure Pt2 according to the engine operating parameters of the aircraft and obtain the second static pressure Ps0 according to the pressure sensing result of the pressure sensor arranged in the engine. Specifically, the engine operating parameters used to calculate the second dynamic pressure Pt2 include the engine operating parameters PS12, P0SEL and the low-pressure rotor speed N1 or the corrected low-pressure rotor speed N1K12. Among them, the operating parameter PS12 represents the fan inlet static pressure of the engine, and the operating parameter P0SEL represents the inlet static pressure of the engine, which is calculated by the engine using the inlet static pressure data source. Specifically, the second dynamic pressure Pt2 can be calculated by the following formula:

[0055] Pt2=P0SEL*f(PS12 / P0SEL, N1K12).

[0056] in:

[0057] N1K12 is based on the N1 corrected speed standardized by TATSEL, where TATSEL is the corrected total temperature data;

[0058] Function f is a function with PS12 / P0SEL and N1K12 as independent variables.

[0059] The GNSS altitude back-calculation static pressure part S3 is the part that calculates the first static pressure Ps1 based on the flight altitude of the aircraft and the atmospheric static equation. Specifically, the GNSS device is used to obtain the current flight altitude signal of the aircraft, that is, the GNSS altitude signal, and the noise is removed by filtering. Then, the first static pressure Ps1 is calculated based on the following mathematical formula:

[0060]

[0061] Wherein, P represents the first static pressure, dP represents the differential change of the first static pressure, g represents the angular velocity of gravity, R represents the universal gas constant, T represents the absolute temperature, and dh is the differential change of the flight altitude.

[0062] The airspeed synthesis part S4 is a part for fusing the first dynamic pressure and the second dynamic pressure into a fused dynamic pressure and fusing the first static pressure and the second static pressure into a fused static pressure, and calculating the airspeed of the aircraft based on the fused dynamic pressure and the fused static pressure. Regarding the fusion of the first dynamic pressure and the second dynamic pressure and the fusion of the first static pressure and the second static pressure, for example, a pressure mean fusion method is adopted. After fusing the first dynamic pressure and the second dynamic pressure and fusing the first static pressure and the second static pressure to form a fused dynamic pressure and a fused static pressure respectively, both are filtered to remove harmful noise. On this basis, the airspeed is calculated based on the above-mentioned fused dynamic pressure and fused static pressure, wherein the synthetic Mach number can be determined by the lift and fused static pressure calculated in the above-mentioned manner.

[0063] Next, refer to Figure 2 , the functional structure of the airspeed calculation system T according to one embodiment of the present application is described.

[0064] like Figure 2As shown, the airspeed calculation system T includes a first dynamic pressure calculation module T1, a second dynamic pressure calculation module T2, a first static pressure calculation module T3, a second static pressure calculation module T4, a pressure fusion module T5 and an airspeed calculation module T6. The first dynamic pressure calculation module T1 calculates the first dynamic pressure Pt1 based on at least the wing area, lift coefficient and lift of the aircraft. The second dynamic pressure calculation module T2 calculates the second dynamic pressure Pt2 based on the above-mentioned engine operating parameters of the aircraft. The first static pressure calculation module T3 calculates the first static pressure Ps1 based on the flight altitude of the aircraft and the atmospheric statics equation. The second static pressure calculation module T4 calculates the second static pressure Ps0 based on the sensing result of the pressure sensor set in the engine. The pressure fusion module T5 fuses the first dynamic pressure Pt1 and the second dynamic pressure Pt2 to form a fused dynamic pressure, and fuses the first static pressure Ps1 and the second static pressure Ps0 to form a fused static pressure. The airspeed calculation module T6 calculates the airspeed of the aircraft based on the fused dynamic pressure and the fused static pressure.

[0065] Figure 3 is a schematic diagram of the structure of a computer system 100 for implementing an aircraft airspeed calculation method. It should be noted that: Figure 3 The computer system shown is only an example of an electronic device and should not impose any limitation on the functions and scope of use of the embodiments of the present invention. Figure 3 The computer system shown.

[0066] like Figure 3 As shown, the computer system 100 includes a central processing unit (CPU) 101, which can perform various appropriate actions and processes according to a program stored in a read-only memory (ROM) 102 or a program loaded from a storage unit 108 to a random access memory (RAM) 103. In the RAM 103, various programs and data required for the operation of the computer system 100 are also stored. The CPU 101, the ROM 102, and the RAM 103 are connected to each other through a bus 104. An input / output (I / O) interface 105 is also connected to the bus 104.

[0067] The I / O interface 105 is connected to an I / O device, which may include an input unit 106 such as a keyboard and a mouse, an output unit 107 such as a liquid crystal display (LCD) and a speaker, a storage unit 108 such as a hard disk, and a communication unit 109 of a network interface card such as a modem. The communication unit 109 performs communication processing via a network such as the Internet. A drive 110 may also be connected to the I / O interface 105 as needed. In addition, a removable medium 111 may also be installed on the drive 110 as needed, so that a computer program read from the removable medium 111 can be installed in the storage unit 108 as needed.

[0068] In particular, refer to Figure 1 The described process, i.e., the method for calculating the airspeed of an aircraft according to one embodiment of the present invention, can be implemented as a computer software program. For example, one embodiment of the present invention includes a computer program product, which includes a computer program carried on a computer readable medium, the computer program including a method for executing Figure 1 In such an embodiment, the computer program may be downloaded and installed from a network via the communication unit 109, and / or installed from the removable medium 111. When the computer program is executed by the central processing unit (CPU) 101, the above functions defined in the system of the present invention are performed.

[0069] It should be noted that the computer-readable medium shown in the present invention may be a computer-readable signal medium or a computer-readable storage medium or any combination of the above two. The computer-readable storage medium may be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device or device, or any combination of the above. More specific examples of computer-readable storage media may include, but are not limited to: an electrical connection with one or more wires, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the above. In the present invention, a computer-readable storage medium may be any tangible medium containing or storing a program that can be used by or in combination with an instruction execution system, device or device. In the present invention, a computer-readable signal medium may include a data signal propagated in a baseband or as part of a carrier wave, which carries a computer-readable program code. This propagated data signal may take a variety of forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination of the above. Computer-readable signal media may also be any computer-readable medium other than computer-readable storage media, which may send, propagate or transmit a program for use by or in conjunction with an instruction execution system, apparatus or device. The program code contained on the computer-readable medium may be transmitted using any appropriate medium, including but not limited to: wireless, wire, optical cable, RF, etc., or any suitable combination of the above.

[0070] As another aspect, the present invention further provides a computer-readable storage medium, which may be included in the computer system described in the above embodiment, or may exist independently without being installed in the computer system. The above computer-readable storage medium carries one or more programs, and when the above one or more programs are executed by a computer system, the computer system implements the method of the above embodiment and its variants. For example, the above computer system may implement Figure 1 The steps shown.

[0071] According to one aspect of the present invention, a computer program product is provided, the computer program product comprising computer instructions, the computer instructions being stored in a computer-readable storage medium. A processor of a computer device reads the computer instructions from the computer-readable storage medium, and the processor executes the computer instructions, so that the computer device executes the method provided in various optional implementations of the above-mentioned embodiments and variants.

[0072] It should be understood that the present invention is not limited to the exact construction that has been described above and shown in the drawings and that various modifications and changes may be made without departing from the scope thereof. The scope of the present invention is limited only by the appended claims.

Claims

1. A method for calculating the airspeed of an aircraft, characterized in that: include: Calculating a first dynamic pressure based on at least a wing area, a lift coefficient, and a lift of the aircraft; calculating a second dynamic pressure based on engine operating parameters of the aircraft; Calculating a first static pressure according to the flight altitude of the aircraft and an atmospheric statics equation; calculating a second static pressure according to a pressure sensor provided in the engine; fusing the first dynamic pressure and the second dynamic pressure to form a fused dynamic pressure, and fusing the first static pressure and the second static pressure to form a fused static pressure; as well as The airspeed of the aircraft is calculated based on the fusion dynamic pressure and the fusion static pressure.

2. The airspeed calculation method according to claim 1, characterized in that: The lift coefficient is determined based on the angle of attack of the aircraft, a lift coefficient increment and an initial lift coefficient, wherein the lift coefficient increment is a lift coefficient increment caused by at least an angle change of a high lift device of the aircraft, and the initial lift coefficient is a lift coefficient when the angle of attack is zero.

3. The airspeed calculation method according to claim 2, characterized in that: The lift coefficient is determined by the following mathematical formula: C L =C L0 +ΔC L +C Lα ×α Among them, C L represents the lift coefficient, α represents the angle of attack, C L0 represents the initial lift coefficient, ΔC L represents the lift coefficient increment, C Lα is the slope when the angle of attack is the independent variable and the lift coefficient is the dependent variable.

4. The airspeed calculation method according to claim 1, characterized in that: The engine operating parameters include the fan inlet static pressure, the intake duct inlet static pressure and the low-pressure rotor speed of the engine.

5. The airspeed calculation method according to claim 1, characterized in that: The atmospheric statics equation is expressed by the following mathematical formula: Wherein, P represents the first static pressure, dP represents the differential change of the first static pressure, g represents the angular velocity of gravity, R represents the universal gas constant, T represents the absolute temperature, and dh is the differential change of the flight altitude.

6. The airspeed calculation method according to claim 1, characterized in that: The lift is determined by the aircraft's angle of attack, lift-to-drag ratio, mass, and normal load coefficient.

7. The airspeed calculation method according to claim 2, characterized in that: The lift coefficient increment is determined based on at least one of a change in a flap angle and a change in a slat angle of the aircraft.

8. An airspeed calculation system, characterized in that: include: a first dynamic pressure calculation module, wherein the first dynamic pressure calculation module calculates a first dynamic pressure according to at least a wing area, a lift coefficient, and a lift of the aircraft; a second dynamic pressure calculation module, wherein the second dynamic pressure calculation module calculates a second dynamic pressure according to an engine operating parameter of the aircraft; A first static pressure calculation module, wherein the first static pressure calculation module calculates a first static pressure according to a flight altitude of the aircraft and an atmospheric statics equation; a second static pressure calculation module, wherein the second static pressure calculation module calculates a second static pressure according to a pressure sensor provided in the engine; a pressure fusion module, wherein the pressure fusion module fuses the first dynamic pressure and the second dynamic pressure to form a fused dynamic pressure, and fuses the first static pressure and the second static pressure to form a fused static pressure; as well as An airspeed calculation module is used to calculate the airspeed of the aircraft according to the fusion dynamic pressure and the fusion static pressure.

9. A computer device comprising a memory, a processor and a computer program stored in the memory, characterized in that: The processor executes the computer program to implement the method according to any one of claims 1 to 7.

10. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the method according to any one of claims 1 to 7 is implemented.

11. A computer program product, comprising a computer program, characterized in that When the computer program is executed by a processor, the method according to any one of claims 1 to 7 is implemented.

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