Method and experimental device for determining aerodynamic characteristics and parameters of aircraft

Abstract

The present invention discloses a method for determining aerodynamic characteristics and parameters of an aircraft and an experimental device. The experimental device comprises an onboard computer, an onboard controller and an onboard information network, wherein the onboard computer and the onboard controller establish a data connection via the onboard information network. The method comprises: step one, preparation work; step two, data acquisition; step three, calculating the actual power performance and aerodynamic parameters of the engine; and step four, calculating the flight characteristics. The present invention determines the actual available capacity of the aircraft engine by identifying the actual power performance and aerodynamic parameters of the aircraft's actual engine under actual flight conditions, and evaluating the flight characteristics of the aircraft in combination with changes in the external environment, thereby reducing the risk of emergencies occurring when flight conditions change. In addition, the present invention does not require testing the engine on a professional test bench or performing aerodynamic testing on the aircraft in a wind tunnel, thereby reducing experimental costs and improving experimental efficiency.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft, and in particular to a method and an experimental device for determining aerodynamic characteristics and parameters of an aircraft. Background Art

[0002] Aircraft are instruments that fly in the atmosphere or in outer space, including aircraft and spacecraft. At present, manned / unmanned aircraft are able to achieve automatic flight, but improving the flight safety of manned / unmanned aircraft is still an urgent problem to be solved. If the actual maximum available power or thrust of the aircraft engine and reliable information on aerodynamic characteristics can be obtained, the aircraft can be made to fly at a given speed and path. The existing methods for determining the actual characteristics of engines and aircraft are mostly completed in wind tunnels or other special test benches. The aerodynamic research of aircraft is carried out through wind tunnel tests and flight tests, which has the following defects: First, a wind tunnel or other special test bench is required, and it needs to be equipped with complex hardware equipment. The experimental process is also relatively cumbersome, and there are defects of long time and high cost; second, the changes in flight characteristic parameters caused by changes in the external environment are not considered, thereby reducing the safety of the aircraft in a changing environment. Summary of the invention

[0003] The purpose of the present invention is to provide a method and an experimental device for determining the aerodynamic characteristics and parameters of an aircraft, so as to solve the problems raised in the above-mentioned background technology.

[0004] To achieve the above-mentioned purpose, the present invention provides the following technical solution: an experimental device for aircraft aerodynamic characteristics and parameters, including an airborne computer, an airborne controller and an airborne information network, wherein the airborne computer and the airborne controller establish a data connection through the airborne information network.

[0005] Preferably, the airborne controller includes a sensor module, a computing module, a microprocessor and a Beidou / GPS satellite navigation receiver.

[0006] Preferably, the sensor module includes a gyroscope, an accelerometer, a magnetometer, a barometer, a thermometer and an airspeed sensor.

[0007] Preferably, the computing module includes a microcomputer mainboard and a radio altimeter, and the processor used in the microcomputer mainboard is a roadcom BCM2837B0 processor.

[0008] The method for determining the aerodynamic characteristics and parameters of an aircraft comprises the following steps: step 1, preparation; step 2, data collection; step 3, calculation of the actual power performance and aerodynamic parameters of the engine; step 4, calculation of the flight characteristics;

[0009] In the above step 1, an onboard computer and an onboard controller are installed on the aircraft, and a data connection between the onboard computer and the onboard controller is established through an onboard information network;

[0010] In the above step 2, the aircraft performs short-term horizontal acceleration flight within 15-30 seconds when the engine is operating at the maximum speed, and the onboard controller records the data of the flight speed change. When the Beidou / GPS satellite navigation receiver receives the independent data of the current coordinates and flight speed, the onboard controller calculates the coordinates and flight speed and corrects them, and sends the corrected flight speed data to the onboard computer at a frequency of not less than 2 Hz;

[0011] In the above step 3, after receiving the flight speed data of the aircraft, the onboard computer calculates the actual power performance and aerodynamic parameters of the aircraft engine;

[0012] In the above step 4, under the given aerodynamic shape, current external environment and actual engine power performance, the onboard computer calculates the flight characteristics and sends them to the onboard controller navigation calculation unit, wherein the flight characteristics include maximum endurance flight speed, maximum range flight speed, climb rate, maximum level flight altitude, actual cruising speed and flight distance.

[0013] Preferably, in step 2, the time interval for the onboard controller to record the data of flight speed changes is equal to the interval for the correction unit to correct the aircraft coordinates and speed using the Kalman filter, and the time interval is determined by the correction interval associated with the current coordinate and speed information update frequency of the Beidou / GPS satellite navigation system.

[0014] Preferably, in step three, the actual power performance of the engine of the jet aircraft is the actual maximum available thrust of the engine, and the actual power performance of the engine of the rotorcraft is the actual maximum available power of the engine, and the aerodynamic parameters of the jet aircraft and the rotorcraft include the zero-lift drag coefficient and the induced drag coefficient.

[0015] Preferably, in step 3, the program of the onboard computer forms a regression equation in the process of acquiring the flight speed data. The left side of the equation is the drag and maximum thrust of the aircraft, and the right side is the inertial force of the jet aircraft in horizontal accelerated flight. The equation is a balance equation for the available residual thrust, and the specific form is as follows:

[0016]

[0017] Among them, T a(max) is the maximum available thrust, C x0 is the zero-lift drag coefficient, k Dis the induced drag coefficient, W is the gravity on the aircraft (N), m is the mass of the aircraft (kg), and ρ is the air density (kg / m 3 ), S is the wing area (m 2 ), V i t i The flight speed at the moment, V a(i) t i The air velocity at the moment, n is the amount of velocity data obtained during the short-term acceleration flight;

[0018] For a rotary-wing aircraft, the regression equation is the following balance equation for the available excess thrust:

[0019]

[0020] Among them, P a(max) is the maximum available power (W).

[0021] Preferably, in step 3, the balance equation of available residual thrust is solved by the least square method to obtain the actual power performance and aerodynamic parameters of the aircraft engine, specifically:

[0022] For equation (1), a(max) , C x0 and k D Three variables and equation (2) about P a(max) , C x0 and k D The three quantities are solved by the least square method. The number of regression equations is equal to the number n of recorded speed values ​​and should be much larger than the number of required parameters determined by this method, i.e., n>>3;

[0023] For the least squares method, the thrust balance equation (1) or the power balance equation (2) is expressed as follows:

[0024] a i ·Xb i ·C x0 -d i ·k D =e i , i=1,n (3)

[0025] Among them, for equation (1), we have:

[0026] X=T a(max) , a i =1,

[0027] For equation (2), we have:

[0028] X=P a(max) , a i=1,

[0029] According to the least squares algorithm, the regression equation (1) or (2) will be simplified to the following matrix equation about three unknown parameters:

[0030]

[0031] in,

[0032] a 12 =a 21 =-∑ i a i b i , a 13 =a 31 =-∑ i a i d i , s1=∑ i a i e i ;

[0033] a 23 =a 32 =∑ i b i d i , s2=∑ i b i e i ;

[0034] s3=∑ i d i e i .

[0035] Preferably, in step 4, the aircraft onboard computer (1) is based on the maximum thrust T max or power P max , zero lift drag coefficient C x0 , induced drag coefficient k D , calculate the flight characteristics of the aircraft under the front aerodynamic shape, specifically:

[0036] Maximum range flight speed:

[0037]

[0038] Cruising speed:

[0039] V cruise =1.3V bg

[0040] Maximum endurance flight speed:

[0041]

[0042] Maximum speed:

[0043]

[0044] Maximum climb rate speed:

[0045] V ROCmax =V md (7)

[0046] Maximum horizontal flight altitude: First, calculate the ratio of the air density corresponding to the maximum altitude to the air density at sea level according to formula (8), and then calculate the corresponding altitude by the formula;

[0047]

[0048]

[0049] Rotorcraft flight distance:

[0050]

[0051] Among them, σ abs_ceiling is the ratio of the air density at the maximum flight altitude to the air density at sea level, c t is the fuel consumption rate, C y is the lift coefficient of cruising flight, W1 is the weight of the aircraft at the current moment, W2 is the weight of the aircraft when it reaches the destination, C L is the lift coefficient, C x is the drag coefficient of the aircraft, and:

[0052]

[0053] Compared with the prior art, the beneficial effects of the present invention are as follows: the present invention determines the actual available capacity of the aircraft engine by identifying the actual engine power performance and aerodynamic parameters of the aircraft under actual flight conditions, and evaluating the flight characteristics of the aircraft in combination with changes in the external environment, thereby reducing the risk of emergencies when flight conditions change. In addition, the present invention does not require testing the engine on a professional test bench or performing aerodynamic testing on the aircraft in a wind tunnel, thereby reducing experimental costs and improving experimental efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0054] Figure 1 It is a structural block diagram of the experimental device of the present invention;

[0055] Figure 2 It is a structural block diagram of the onboard controller of the present invention;

[0056] Figure 3It is a data flow diagram of the hardware module of the experimental device of the present invention;

[0057] Figure 4 It is a speed variation curve diagram of a rotorcraft in horizontal accelerated flight when the piston engine is running at the maximum speed;

[0058] Figure 5 It is a speed variation curve diagram of the aircraft during horizontal acceleration flight when the jet engine is running at maximum thrust;

[0059] Figure 6 The following is a typical graph showing the variation of the residual power of a rotorcraft with the flight speed when the piston engine is running at the maximum speed;

[0060] Figure 7 The present invention is a flow chart of the method.

[0061] In the figure: 1. Onboard computer; 2. Onboard controller; 21. Sensor module; 211. Gyroscope; 212. Accelerometer; 213. Magnetometer; 214. Barometer; 215. Thermometer; 216. Airspeed sensor; 22. Computing module; 221. Microcomputer motherboard; 222. Radio altimeter; 23. Microprocessor; 24. Beidou / GPS satellite navigation receiver; 3. Onboard information network. DETAILED DESCRIPTION

[0062] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0063] See also Figure 1-3The present invention provides an embodiment: an experimental device for aerodynamic characteristics and parameters of an aircraft, comprising an onboard computer 1, an onboard controller 2 and an onboard information network 3. The onboard computer 1 and the onboard controller 2 establish a data connection through the onboard information network 3. The onboard computer 1 is used to determine the actual maximum available thrust or power and main aerodynamic parameters of the aircraft engine, and calculate the flight characteristics according to the speed change data during short-time horizontal acceleration flight corrected by the Kalman filter from the onboard controller 2. The onboard controller 2 is used to calculate and correct the coordinates and attitude, flight speed and navigation parameters of the aircraft in the inertial system in real time. The onboard information network 3 is used to realize data transmission between the onboard computer 1 and the onboard controller 2; the onboard controller 2 includes a sensor module 21, a computing module 22, a microprocessor 23 and a Beidou / GPS satellite navigation receiver 24, the sensor module 21, the microprocessor 23 and the Beidou / GPS satellite navigation receiver 24 constitute an airborne automatic control system, the calculation module 22 is used to calculate parameters, and the model of the microprocessor 23 is STM32F427ZG; the sensor module 21 includes a gyroscope 211, an accelerometer 212, a magnetometer 213, a barometer 214, a thermometer 215 and an airspeed sensor 216, the sensor module 21 is used to collect aircraft attitude data and environmental data, the gyroscope 211 and the accelerometer 212 constitute an inertial sensor SINS; the calculation module 22 includes a microcomputer motherboard 221 and a radio altimeter 222, the processor used by the microcomputer motherboard 221 is a roadcom BCM2837B0Cortex-A53 processor, the microcomputer motherboard 221 is used to calculate the speed change data of the aircraft during short-term horizontal acceleration flight, and the radio altimeter 222 is used to obtain aircraft altitude data.

[0064] See also Figure 4-7 , an embodiment provided by the present invention: a method for determining aerodynamic characteristics and parameters of an aircraft, comprising step one, preparation work; step two, data collection; step three, calculating the actual power performance and aerodynamic parameters of the engine; step four, calculating the flight characteristics;

[0065] In the above step 1, the onboard computer 1 and the onboard controller 2 are installed on the aircraft, and a data connection is established between the onboard computer 1 and the onboard controller 2 via the onboard information network 3;

[0066] In the above step 2, the aircraft performs short-term horizontal accelerated flight within 15-30 seconds when the engine is operating at the maximum speed, and the airborne controller 2 records the data of flight speed change. When the Beidou / GPS satellite navigation receiver 24 receives the independent data of the current coordinates and flight speed, the airborne controller 2 calculates the coordinates and flight speed and corrects them, and sends the corrected flight speed data to the airborne computer 1 at a frequency of not less than 2Hz; wherein the time interval for the airborne controller 2 to record the data of flight speed change is equal to the interval for the correction unit to correct the coordinates and speed of the aircraft using the Kalman filter, and the time interval is determined by the correction interval associated with the update frequency of the current coordinate and speed information of the Beidou / GPS satellite navigation system;

[0067] In the above step 3, after receiving the flight speed data of the aircraft, the onboard computer 1 calculates the actual engine power performance and aerodynamic parameters of the aircraft; wherein, the actual engine power performance of the jet aircraft is the actual maximum available thrust of the engine, and the actual engine power performance of the rotary-wing aircraft is the actual maximum available power of the engine, and the aerodynamic parameters of the jet aircraft and the rotary-wing aircraft include the zero-lift drag coefficient and the induced drag coefficient; the program of the onboard computer 1 forms a regression equation in the process of obtaining the flight speed data, the left side of the equation is the drag and maximum thrust of the aircraft, and the right side is the inertia force of the jet aircraft in horizontal accelerated flight, and the equation is a balance equation for the available residual thrust, and the specific form is as follows:

[0068]

[0069] Among them, T a(max) is the maximum available thrust, C x0 is the zero-lift drag coefficient, k D is the induced drag coefficient, W is the gravity on the aircraft (N), m is the mass of the aircraft (kg), and ρ is the air density (kg / m 3 ), S is the wing area (m 2 ), V i t i The flight speed at the moment, V a(i) t i The air velocity at the moment, n is the amount of velocity data obtained during the short-term acceleration flight;

[0070] For a rotary-wing aircraft, the regression equation is the following balance equation for the available excess thrust:

[0071]

[0072] Among them, P a(max) is the maximum available power (W);

[0073] The balance equation of available residual thrust is solved by the least square method to obtain the actual engine power performance and aerodynamic parameters of the aircraft, specifically:

[0074] For equation (1), a(max) , C x0 and k D Three variables and equation (2) about P a(max) , C x0 and k D The three quantities are solved by the least square method. The number of regression equations is equal to the number n of recorded speed values ​​and should be much larger than the number of required parameters determined by this method, i.e., n>>3;

[0075] For the least squares method, the thrust balance equation (1) or the power balance equation (2) is expressed as follows:

[0076] a i ·Xb i ·C x0 -d i ·k D =e i , i=1,n (3)

[0077] Among them, for equation (1), we have:

[0078] X=T a(max) , a i =1,

[0079] For equation (2), we have:

[0080] X=P a(max) , a i =1,

[0081] According to the least squares algorithm, the regression equation (1) or (2) will be simplified to the following matrix equation about three unknown parameters:

[0082]

[0083] in,

[0084] a 12 =a 21 =-∑ i a i b i , a 13 =a 31 =-∑ i a i d i , s1=∑ i ai e i ;

[0085] a 23 =a 32 =∑ i b i d i , s2=∑ i b i e i ;

[0086] s3=∑ i d i e i ;

[0087] In the above step 4, under the given aerodynamic shape, current external environment and actual engine power performance, the onboard computer 1 calculates the flight characteristics and sends them to the navigation calculation unit of the onboard controller 2, wherein the flight characteristics include the maximum endurance flight speed, the maximum range flight speed, the climb rate, the maximum altitude of horizontal flight, the actual cruising speed and the flight distance of the aircraft; the onboard computer 1 calculates the flight characteristics based on the maximum thrust T max or power P max , zero lift drag coefficient C x0 , induced drag coefficient k D , calculate the flight characteristics of the aircraft under the front aerodynamic shape, specifically:

[0088] Maximum range flight speed:

[0089]

[0090] Cruising speed:

[0091] V cruise =1.3V bg

[0092] Maximum endurance flight speed:

[0093]

[0094] Maximum speed:

[0095]

[0096] Maximum climb rate speed:

[0097] V ROCmax =V md (7)

[0098] Maximum horizontal flight altitude: First, calculate the ratio of the air density corresponding to the maximum altitude to the air density at sea level according to formula (8), and then calculate the corresponding altitude by the formula;

[0099]

[0100]

[0101] Rotorcraft flight distance:

[0102]

[0103] Among them, σ abs_ceiling is the ratio of the air density at the maximum flight altitude to the air density at sea level, c t is the fuel consumption rate, C y is the lift coefficient of cruising flight, W1 is the weight of the aircraft at the current moment, W2 is the weight of the aircraft when it reaches the destination, C L is the lift coefficient, C x is the drag coefficient of the aircraft, and:

[0104]

[0105] Verify the accuracy of the above method:

[0106] A typical curve of the residual power of a rotorcraft with a piston engine running at maximum speed mode versus flight speed is as follows: Figure 6 As shown, by using a fixed speed interval Δt=t i -t i-1 =const(Δt=1sor2s) The right side of equation (2) is obtained by calculating during horizontal accelerated flight. The dotted line is the reference residual power change, which is calculated based on the wind tunnel test data of the aircraft model and the thrust characteristic data of the brushless motor propeller. Table 1 is a numerical table of reference values ​​and calculated values. The maximum power, zero-lift drag coefficient, and induced drag coefficient of the aircraft are obtained by least squares calculation and experiment during the flight of the experimental aircraft. Table 2 gives the numerical table of parameters (reference values ​​and calculated values). The maximum power, zero-lift drag coefficient, and induced drag coefficient of the experimental aircraft are obtained by modeling the accelerated flight of the aircraft, so as to select the optimal time interval Δt for a fixed flight speed. The accuracy of the above method is verified by comparing the wind tunnel test data with the data calculated by the proposed method.

[0107] Table 1 Reference values ​​and calculated values

[0108]

[0109] Table 2 Numerical table of parameter variables

[0110]

[0111] Based on the above, the advantage of the present invention is that when the present invention is used, the speed change data of the aircraft performing horizontal accelerated flight at the maximum engine speed in a short period of time is used to identify the actual maximum thrust or power of the aircraft engine, as well as the zero-lift drag coefficient and induced drag coefficient of the aircraft, and determine the flight characteristics of the aircraft under a given aerodynamic shape, the current external environment and the actual power performance of the engine. This method does not require wind tunnel testing of the aircraft model, nor does it require specifying the aerodynamic characteristics of the aircraft and the thrust or power parameters of the main thrust engine in the form of a table or function. Instead, it determines the actual zero-lift drag coefficient, the maximum actual available thrust or the maximum actual available power, and the actual induced drag coefficient of the current aerodynamic shape of the aircraft. These parameters are sufficient to calculate the flight characteristics of the aircraft under the current external environmental conditions, thereby reducing the risk of emergency situations occurring in the aircraft due to changes in flight conditions while reducing experimental costs and improving experimental efficiency.

[0112] It will be apparent to those skilled in the art that the invention is not limited to the details of the exemplary embodiments described above and that the invention can be implemented in other specific forms without departing from the spirit or essential features of the invention. Therefore, the embodiments should be considered exemplary and non-limiting in all respects, and the scope of the invention is defined by the appended claims rather than the foregoing description, and it is intended that all variations falling within the meaning and scope of the equivalent elements of the claims be included in the invention. Any reference numeral in a claim should not be considered as limiting the claim to which it relates.

Claims

1. A method for determining aerodynamic characteristics and parameters of an aircraft, the method using an experimental device to determine the aerodynamic characteristics and parameters of an aircraft, the experimental device comprising an onboard computer (1), an onboard controller (2) and an onboard information network (3), the onboard computer (1) and the onboard controller (2) establishing a data connection via the onboard information network (3); It is characterized in that The steps of this method are as follows: Step 1: Preparation; Step 2: Data collection; Step 3: Calculation of actual engine power performance and aerodynamic parameters; Step 4: Calculation of flight characteristics; In the above step 1, the onboard computer (1) and the onboard controller (2) are installed on the aircraft, and a data connection is established between the onboard computer (1) and the onboard controller (2) via the onboard information network (3); In the above step 2, the aircraft performs short-term horizontal acceleration flight within 15-30 seconds when the engine is operating at the maximum speed, and the onboard controller (2) records the data of the flight speed change. When the Beidou / GPS satellite navigation receiver (24) receives the independent data of the current coordinates and the flight speed, the onboard controller (2) calculates the coordinates and the flight speed and corrects them, and sends the corrected flight speed data to the onboard computer (1) at a frequency of not less than 2 Hz; In the above step 3, after receiving the flight speed data of the aircraft, the onboard computer (1) calculates the actual power performance and aerodynamic parameters of the aircraft engine; In the above step 4, under the given aerodynamic shape, current external environment and actual engine power performance, the onboard computer (1) calculates the flight characteristics and sends them to the navigation calculation unit of the onboard controller (2), wherein the flight characteristics include the maximum endurance flight speed, the maximum range flight speed, the climb rate, the maximum altitude of horizontal flight, the actual cruising speed and the flight distance.

2. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 1, characterized in that: The onboard controller (2) comprises a sensor module (21), a computing module (22), a microprocessor (23) and a Beidou / GPS satellite navigation receiver (24).

3. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 2, characterized in that: The sensor module (21) comprises a gyroscope (211), an accelerometer (212), a magnetometer (213), a barometer (214), a thermometer (215) and an airspeed sensor (216).

4. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 2, characterized in that: The computing module (22) comprises a microcomputer mainboard (221) and a radio altimeter (222); the processor used by the microcomputer mainboard (221) is a Roadcom BCM2837B0 processor.

5. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 1, characterized in that: In the step 2, the time interval at which the onboard controller (2) records the data of the flight speed change is equal to the interval at which the correction unit uses the Kalman filter to correct the coordinates and speed of the aircraft, and the time interval is determined by the correction interval associated with the update frequency of the current coordinate and speed information of the Beidou / GPS satellite navigation system.

6. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 1, characterized in that: In the step three, the actual power performance of the engine of the jet aircraft is the actual maximum available thrust of the engine, and the actual power performance of the engine of the rotary-wing aircraft is the actual maximum available power of the engine. The aerodynamic parameters of the jet aircraft and the rotary-wing aircraft include the zero-lift drag coefficient and the induced drag coefficient.

7. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 1, characterized in that: In step 3, the program of the onboard computer (1) forms a regression equation in the process of acquiring the flight speed data. The left side of the equation is the drag and maximum thrust of the aircraft, and the right side is the inertia force of the jet aircraft in horizontal accelerated flight. The equation is a balance equation for the available residual thrust, and the specific form is as follows: Among them, T a(max) is the maximum available thrust, C x0 is the zero-lift drag coefficient, k D is the induced drag coefficient, W is the gravity of the aircraft, m is the mass of the aircraft, ρ is the air density, S is the wing area, V i t i The flight speed at the moment, V a(i) t i The air velocity at the moment, n is the amount of velocity data obtained during the short-term acceleration flight; For a rotary-wing aircraft, the regression equation is the following balance equation for the available excess thrust: Among them, P a(max) is the maximum available power (W).

8. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 1, characterized in that: In step 3, the balance equation of available residual thrust is solved by the least square method to obtain the actual power performance and aerodynamic parameters of the aircraft engine, specifically: For equation (1), a(max) , C x0 and k D Three variables and equation (2) about P a(max) , C x0 and k D The three quantities are solved by the least square method. The number of regression equations is equal to the number n of recorded speed values ​​and should be much larger than the number of required parameters determined by this method, i.e. n>>3; For the least squares method, the thrust balance equation (1) or the power balance equation (2) is expressed as follows: a i ·X-b i ·C x0 -d i ·k D =e i ,i=1,n (3) Among them, for equation (1): For equation (2), we have: X=P a(max) ,a i =1, According to the least squares algorithm, the regression equation (1) or (2) will be simplified to the following matrix equation about three unknown parameters: in, a 12 =a 21 =-∑ i a i b i ,a 13 =a 31 =-∑ i a i d i ,s1=∑ i a i e i ; a 23 =a 32 =∑ i b i d i ,s2=Σ i b i e i ; s3=∑ i d i And i 。 9. The method for determining aerodynamic characteristics and parameters of an aircraft according to claim 1, characterized in that: In step 4, the aircraft onboard computer (1) calculates the maximum thrust T max or power P max , zero lift drag coefficient C x0 , induced drag coefficient k D , calculate the flight characteristics of the aircraft under the front aerodynamic shape, specifically: Maximum range flight speed: Cruising speed: V cruise =1.3V bg Maximum endurance flight speed: Maximum speed: Maximum climb rate speed: V ROCmax =V md (7) Maximum horizontal flight altitude: First, calculate the ratio of the air density corresponding to the maximum altitude to the air density at sea level according to formula (8), and then calculate the corresponding altitude by the formula; Rotorcraft flight distance: Among them, σ abs_ceiling is the ratio of the air density at the maximum flight altitude to the air density at sea level, c t is the fuel consumption rate, C y is the lift coefficient of cruising flight, W1 is the weight of the aircraft at the current moment, W2 is the weight of the aircraft when it reaches the destination, C L is the lift coefficient, C x is the drag coefficient of the aircraft, and:

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