Parameter design method for hybrid composite wing aircraft

Abstract

The present invention provides a parameter design method for a hybrid composite wing aircraft, first determining the takeoff mass target requirement and design input parameters of a fixed rotor aircraft; then determining the aircraft design boundary according to the aircraft cruise constraint, climb constraint, ceiling constraint, stall constraint and vertical climb constraint in a fixed wing mode, drawing the relationship curve between each constraint and the power-to-weight ratio and the wing load, and determining the design boundary of the engine, generator, fixed wing mode motor and rotor mode motor according to the relationship curve; selecting a design point on the design boundary, considering the additional mass of the distributed propulsion system, and calculating the propulsion system mass; traversing all design points of the aircraft design boundary, iteratively solving the takeoff mass corresponding to each design point, selecting the minimum value as the design required aircraft takeoff mass, and calculating the mass of the aircraft propulsion system, fuel and battery and the wing load. The estimation result of the method of the present invention is accurate and conforms to the actual aircraft.

Description

Technical Field

[0001] The invention belongs to the technical field of hybrid power composite wing aircraft, and in particular relates to a parameter design method for a hybrid power composite wing aircraft. Background Art

[0002] A compound wing aircraft is an aircraft that combines traditional fixed wings and rotors. It has both the high-speed cruising capability of traditional fixed wings during horizontal flight and the vertical take-off and landing function of traditional rotors, making it suitable for more flight scenarios.

[0003] At present, composite wing aircraft mainly use pure electric propulsion, and pure electric propulsion aircraft are widely used in various short-distance and light missions. However, in the case of long-distance flight requirements, due to the relatively low energy density of batteries, the current electric propulsion system may find it difficult to achieve flight distance and endurance comparable to traditional internal combustion engine power systems. The thermal efficiency of a general internal combustion engine is around 40%. In order to further improve the overall efficiency of the propulsion system, a hybrid propulsion system has become a major solution. Hybrid propulsion systems include three types: series, parallel, and hybrid. The use of a series hybrid propulsion system allows the propulsion system of a composite wing aircraft to adopt a distributed layout, thereby providing greater flexibility and freedom in the design of the aircraft.

[0004] Most of the research on the design method of hybrid composite wing aircraft focuses on pure electric propulsion. However, since these methods only consider the propulsion system of a single source of energy, and the propulsion system of a hybrid aircraft includes an engine and an electric motor, such design methods are not suitable for hybrid aircraft. In addition, in the existing parameter design method of hybrid composite wing aircraft, the vertical climb performance limitation of the aircraft and the additional weight added by the distributed propulsion system are not fully considered, which will lead to the estimated take-off weight of the aircraft being too small, and also lead to the estimated power of the motor in the rotor mode being too low. Therefore, a more accurate method is needed to complete the preliminary parameter design of hybrid composite wing aircraft. Summary of the invention

[0005] The purpose of the present invention is to solve the problems of low motor power of rotor mode and small total weight of aircraft in the conceptual design stage in the traditional aircraft design method, and to provide a parameter design method for hybrid composite wing aircraft.

[0006] To achieve the above purpose, the technical solution provided by the present invention is:

[0007] A parameter design method for a hybrid composite wing aircraft, wherein the parameters include aircraft take-off mass, propulsion system mass, fuel mass, battery mass and wing loading; the special feature of the method is that the design method includes the following steps:

[0008] Step 1: Determine the takeoff mass target requirement M of the fixed rotor aircraft payload and design input parameters;

[0009] The design input parameters include cruising speed, rotor mode climb rate, fixed wing mode climb rate, ceiling, stall, lift-to-drag ratio, take-off mass initial value MTOM 0 and payload;

[0010] Step 2: Establish a hybrid composite wing aircraft model. In fixed-wing mode, determine the design boundaries of the aircraft through the aircraft's cruise constraints, climb constraints, ceiling constraints, stall constraints, and vertical climb constraints, and plot each constraint against the power-to-weight ratio. and wing loading and determining the design boundaries of the engine, the generator, the fixed-wing mode motor, and the rotary-wing mode motor according to the relationship curve;

[0011] Step 3: On the design boundary determined in step 2, select the design points of the engine, generator, fixed-wing mode motor, and rotor mode motor under the same wing load, and calculate the power-to-weight ratio of the selected design point according to the initial value of the takeoff mass and the power-to-weight ratio of the selected design point. and wing loading The mass M of the aircraft's propulsion system is calculated ps , the mass of the propulsion system does not include the mass of the engine;

[0012] In the process of calculating the mass of the propulsion system, the additional mass brought by the distributed propulsion system is taken into account, and the additional mass includes the mass of the accessories of the motor and the speed regulator, and the mass of the power supply line;

[0013] Step 4: Calculate the total energy transmitted by the aircraft at each stage of flight, and convert and calculate the total energy mass of the aircraft according to the energy utilization efficiency. The total energy mass is equal to the fuel mass M fuel and battery mass M ba sum;

[0014] Step 5: Select the design point on the aircraft design boundary and calculate the initial takeoff mass MTOM of the aircraft according to step 1. 0 And the process of step 3 and step 4 calculates the total energy mass M of the aircraft energy and propulsion system mass M ps , with total energy mass M energy , propulsion system mass M ps and take-off mass target requirement M payload As the input parameter for the first iteration calculation of the aircraft takeoff mass; set the iteration stop threshold, and iterate to obtain the current iteration output value MTOM NEW, and according to the output value MTOM NEW Update the total energy mass and propulsion system mass of the aircraft using the formulas in steps 3 and 4, and convert the takeoff mass target requirement M payload The updated total energy mass and propulsion system mass are used as input parameters for the next iterative calculation; all design points on the aircraft design boundary are traversed to obtain the aircraft takeoff mass corresponding to each design point when the iteration stops, and the minimum takeoff mass is selected as the aircraft takeoff mass required for the design;

[0015] Step 6: Calculate the mass of the aircraft propulsion system M according to the parameter information of the design point corresponding to the minimum takeoff mass selected in step 5. ps 、Fuel quality M fuel 、Battery quality M ba and wing loading

[0016] Furthermore, the step 2 specifically includes the following steps:

[0017] Step 2.1: Plot the power-to-weight ratio based on the cruise constraint formula for composite wing aircraft and wing loading Cruise constraint curve in coordinates;

[0018] The cruise constraint formula of the composite wing aircraft is:

[0019] Where T is thrust, W is the takeoff weight of the aircraft, S is the wing area, and C is D,min is the minimum drag coefficient of the aircraft, C D,min The estimated value is 0.028~0.035; k is the induced drag constant, V is the aircraft speed, η prop is the propeller efficiency, q is the dynamic pressure, ρ is the air density, the standard value is 1.225;

[0020] Step 2.2: Plot the power-to-weight ratio based on the climb constraint formula in fixed-wing mode and wing loading Climb constraint curve in coordinates;

[0021] The climb constraint formula in the fixed-wing mode is:

[0022] Where V 1 is the aircraft's climb rate, is the best climb rate;

[0023] Step 2.3: Plot the power-to-weight ratio based on the stall constraint formula for composite wing aircraft and wing loading Stall curve in coordinates;

[0024] The stall constraint formula of the composite wing aircraft is:

[0025] Where V S is the stall speed, C Lmax is the maximum lift coefficient;

[0026] Step 2.4, according to the ceiling constraint formula Plotting the power-to-weight ratio and wing loading The ceiling constraint curve in coordinates, where V L is the maximum climb rate, and its value is V L =0.508m / s;

[0027] Step 2.5, according to the vertical climb constraint formula Plotting the power-to-weight ratio and wing loading The vertical climb constraint curve of the aircraft under the coordinates;

[0028] Among them, S w is the ratio of the aircraft area to the wing area, which is 1.35; FM is the rotor factor, which ranges from 0.6 to 0.7; V I is the axial climbing speed;

[0029] Step 2.6, according to the cruise constraint curve, climb constraint curve, stall constraint curve, ceiling constraint curve and vertical climb constraint curve drawn in steps 2.1 to 2.5, determine the design boundaries of the engine, generator, fixed-wing mode motor and rotor mode motor; the rules for determining each design boundary are:

[0030] Since the aircraft adopts a series hybrid system, the power of the engine and generator only needs to meet the cruising power, so the design boundaries of the engine and generator are taken on the cruise constraint curve; since the fixed-wing mode motor needs to provide climb power, the design boundaries of the fixed-wing mode motor are taken on the climb constraint curve; based on the fact that the rotor mode motor needs to provide vertical climb power, the design boundaries of the rotor mode motor are taken on the vertical climb constraint curve; and the fixed-wing mode motor power is greater than the power of the engine or generator.

[0031] Furthermore, the step 3 specifically includes the following steps:

[0032] Step 3.1, selecting the design points of the engine, generator, fixed-wing mode motor and rotor mode motor in each design at the same wing load position, calculating the power of the engine, generator, rotor mode motor and fixed-rotor mode motor at the design point, and calculating the total power of the motor;

[0033] Engine power P ICE =k 1 *W; generator power P gen =k 1 *W;

[0034] Motor power in rotor mode P Rmot =k 2 *W; fixed rotor mode motor power P fmot =k 3 *W;

[0035] Total motor power P mot =P Rmot +P fmot ;

[0036] Among them, k 1 , k 2 , k 3 are the values ​​of power-to-weight ratio corresponding to the selected design points respectively;

[0037] Step 3.2, calculate the propeller mass, motor mass, generator mass and electric speed governor mass,

[0038] The propeller mass calculation formula is:

[0039]

[0040] The motor mass calculation formula is:

[0041] The generator mass calculation formula is: M GE =0.385*(P gen +0.44);

[0042] The mass calculation formula of the electric speed regulator is:

[0043] Among them, K m is the correction factor, the value is 0.6; n prop is the number of propellers, n b is the number of leaves; is the motor power coefficient, which takes a value of 0.2; is the power coefficient of the electronic speed regulator, and its value is 0.05;

[0044] Step 3.3, calculate the additional mass brought by the distributed propulsion system, including the mass of the motor and speed regulator accessories and the mass of the power supply line; the mass of the motor and speed regulator accessories is taken as 0.3*M motor ;

[0045] Additional mass M brought by distributed distributed propulsion system EX =0.3*Mmotor +ρ pl *s pl *l pl , where ρ pl , l pl 、s pl They are the material density of the power line, the length of the power line and the cross-sectional area of ​​the power line;

[0046] Step 3.4, calculate the mass of the propulsion system excluding the engine M ps ,

[0047] M ps =M prop +M motor +M GE +M esc +M EX .

[0048] Furthermore, the step 4 specifically includes the following steps:

[0049] Step 4.1, calculating the total energy transferred at each stage of the flight according to the energy requirement of the aircraft to perform the mission; the energy requirement comes from aerodynamic drag, acceleration and altitude change potential energy;

[0050] The total transmission energy of the aircraft is calculated as: Where L / D is the lift-to-drag ratio of the aircraft, t is the cruising time, and m is the takeoff mass of the aircraft;

[0051] Step 4.2, calculate the fuel mass;

[0052] Since the designed engine power is based on the cruising power, it is assumed that the engine power remains unchanged during the cruising process, so the fuel quality where η GE , η mot , η prop , η esc are generator efficiency, motor efficiency, rotor efficiency, and ESC efficiency, and their values ​​are 0.9, 0.925, 0.80, and 0.95, respectively. SFC is the fuel consumption rate, and t is the cruising time;

[0053] Step 4.3, according to the formula Calculate the battery mass M ba , where E * is the battery density, which is 200wh / kg; η ba is the battery efficiency, P ICE is the engine power; ddq is the depth of discharge inclusion coefficient, ranging from 0.2 to 0.25;

[0054] Step 4.4, calculate the total energy mass M of the aircraftenergy =M fuel +M ba .

[0055] Furthermore, the step 5 specifically includes the following steps:

[0056] Step 5.1, select design points on the design boundaries of the engine, generator, fixed-wing mode motor and rotor mode motor determined in step 2, first according to the initial takeoff mass MTOM in step 1 0 And the calculation formulas of step 3 and step 4, we can get the total energy mass M energy and propulsion system mass M ps ; Based on the total energy mass M energy , propulsion system mass M ps and take-off mass target requirement M payload As input parameter for the first iteration calculation of the aircraft takeoff mass;

[0057] Step 5.2, based on the first iteration of the aircraft takeoff mass calculation input parameters and the aircraft takeoff mass iteration formula Iterate the takeoff mass to get the output value MTOM NEW , the current iteration output value MTOM NEW Substitute into steps 3 and 4 to update the total energy mass M energy and propulsion system mass M ps , and the take-off quality target requirement M payload And the updated total energy mass M energy and propulsion system mass M ps As the input parameter of the next iteration, until the output value of this iteration is MTOM NEW(n) The output value of the last iteration is MTOM NEW(n-1) Meet | MTOM NEW(n) -MTOM NEW(n-1) |<ε, the iteration is stopped, where ε is the set stop threshold, n is the number of iterations, and n>1;

[0058] The aircraft takeoff mass iteration formula MF PS is the aircraft mass coefficient, ranging from 0.45 to 0.55;

[0059] Step 5.3, according to steps 5.1-5.2, traverse all design points on the aircraft design boundary, obtain the aircraft takeoff mass output value corresponding to each design point when the iteration stops, and select the minimum takeoff mass as the aircraft takeoff mass required for the design.

[0060] The advantages of the present invention are:

[0061] The present invention provides a parameter design method for a hybrid composite wing aircraft. According to the design input parameters of the aircraft in the conceptual design stage and the aircraft constraint calculation formula, the design boundaries of an engine, a generator, a fixed-wing mode motor and a rotor mode motor are determined. In the process of determining the design boundaries, the vertical climb constraint of the composite wing aircraft is considered; in the process of calculating the mass of a propulsion system by selecting a design point on the design boundary, the additional mass brought by the distributed propulsion system is considered, and the additional mass includes the mass of the accessories of the motor and the speed regulator, and the mass of the power supply line, so that the designed aircraft mass parameter result is accurate and conforms to the actual flight of the aircraft, and avoids the problem that the vertical climb performance limitation of the aircraft and the additional weight added by the distributed propulsion system are not fully considered, resulting in a small calculation result of the aircraft take-off weight and a low estimated power of the motor in the rotor mode. BRIEF DESCRIPTION OF THE DRAWINGS

[0062] The above and / or additional aspects and advantages of the present invention will become apparent and easily understood from the description of the embodiments in conjunction with the following drawings, in which:

[0063] Figure 1 It is the overall flow chart of the parameter design method of the present invention;

[0064] Figure 2 It is a schematic diagram of the design range of traditional aircraft;

[0065] Figure 3 It is the design boundary of the engine, generator, fixed-wing mode motor and rotary-wing mode motor of the hybrid composite wing aircraft drawn by the present invention through five constraints;

[0066] Figure 4 It is a diagram of iterative simulation results of the present invention on the take-off mass of the aircraft;

[0067] Figure 5 It is a simulation result diagram of the take-off mass, propulsion system mass, and energy weight of the aircraft of the present invention. DETAILED DESCRIPTION

[0068] Embodiments of the present invention are described in detail below. The embodiments are exemplary and intended to be used to explain the present invention, but should not be construed as limiting the present invention.

[0069] The parameter design method of a hybrid composite wing aircraft comprises the following steps:

[0070] Step 1: Obtain the target demand M of the fixed rotor aircraft payload and design input parameters, the design input parameters including cruising speed, rotor mode climb rate, fixed wing mode climb rate, ceiling, stall, lift-to-drag ratio, take-off mass initial value MTOM 0 and payload.

[0071] Step 2: In fixed-wing mode, determine the design boundaries of the aircraft through cruise constraints, climb constraints, ceiling constraints, stall constraints, and vertical climb constraints, and draw the power-to-weight ratio respectively. and wing loading The relationship curves under various constraints are used to determine the design boundaries of the engine, generator, fixed-wing mode motor, and rotary-wing mode motor. Specifically, the following steps are included:

[0072] Step 2.1: Plot the power-to-weight ratio based on the cruise constraint formula for the composite wing aircraft and wing loading Cruise constraint curve in coordinates;

[0073] The cruise constraint formula of the composite wing aircraft is:

[0074] Where T is thrust, W is aircraft weight, S is wing area, C is D,min is the minimum drag coefficient of the aircraft, C D,min The estimated value is 0.028~0.035; k is the induced drag constant, V is the aircraft speed, η prop is the propeller efficiency, q is the dynamic pressure, ρ is the air density, and the standard value is 1.225.

[0075] Step 2.2: Plot the power-to-weight ratio based on the climb constraint formula in fixed-wing mode and wing loading Climb constraint curve in coordinates.

[0076] The climb constraint formula in the fixed-wing mode is:

[0077] Where V 1 is the aircraft's climb rate, For the best climb rate.

[0078] Step 2.3: Plot the power-to-weight ratio based on the stall constraint formula for composite wing aircraft and wing loading Stall curve in coordinates;

[0079] The stall constraint formula of the composite wing aircraft is:

[0080] Where V S is the stall speed, C Lmax is the maximum lift coefficient.

[0081] Step 2.4: Calculate the formula using the ceiling constraint Plotting the power-to-weight ratio and wing loading The ceiling constraint curve in coordinates, where V L is the maximum climb rate, and its value is V L =0.508m / s.

[0082] Step 2.5: Use the vertical climb constraint formula Plotting the power-to-weight ratio and wing loading The vertical climb constraint curve of the aircraft in coordinates.

[0083] Among them, S w is the ratio of the aircraft area to the wing area, generally estimated to be 1.35, V I is the axial climbing speed, FM is the rotor factor, and its value is between 0.6 and 0.7.

[0084] Step 2.6: Determine the design boundaries of the aircraft through the five constraint curves drawn in steps 2.1 to 2.5, and determine the design boundaries of the engine, generator, fixed-wing mode motor, and rotor mode motor.

[0085] Since a series hybrid system is used, the power of the engine and generator only needs to meet the cruising power, and the design boundaries of the engine and generator should be on the cruise constraint curve; the fixed-wing mode motor needs to provide climbing power, that is, the design boundary of the fixed-wing mode motor is on the climb constraint curve; the rotor mode motor needs to provide vertical climb power, that is, the design boundary of the rotor mode motor is on the vertical climb constraint curve; in addition, the power of the fixed-wing mode motor must be greater than the power provided by the engine and generator. Based on the above rules, the design boundaries of the engine, generator, fixed-wing mode motor and rotor mode motor are obtained, such as Figure 3 shown.

[0086] Step 3: Select the design points of the engine, generator, fixed-wing mode motor and rotor mode motor on the design boundary determined in step 2, and use the initial takeoff mass MTOM given in step 1 0 And the power-to-weight ratio corresponding to the selected point and wing loading Estimate the mass of the aircraft's propulsion system. Specifically, the following steps are included:

[0087] Step 3.1, select the design points of the engine, generator, fixed-wing mode motor and rotor mode motor in each design at the same wing load position, calculate the power of the engine, generator, rotor mode motor and fixed-rotor mode motor at the design point, and calculate the total motor power. The calculation formulas are:

[0088] Engine power P ICE =k 1*W, generator power P gen =k 1 *W, motor power P in rotor mode Rmot =k 2 *W, fixed rotor mode motor power P fmot =k 3 *W. The total motor power is: P mot =P Rmot +P fmot , where k 1 , k 2 , k 3 are the power-to-weight ratio values ​​of the selected design points respectively.

[0089] Step 3.2, calculate the propeller mass, motor mass, generator mass, and electric speed governor mass;

[0090] The propeller mass calculation formula is:

[0091]

[0092] The formula for calculating motor mass is:

[0093] The formula for calculating the mass of the electric speed regulator is:

[0094] The formula for calculating the generator mass is: M GE =0.385*(P gen +0.44).

[0095] Among them, K m is the correction factor, generally estimated to be 0.6, n prop is the number of propellers, n b is the number of leaves. is the motor power coefficient, which is 0.2; is the power coefficient of the electronic speed regulator, and its value is 0.05.

[0096] Step 3.3, calculate the additional mass brought by the distributed propulsion system. The additional mass brought by the distributed propulsion system includes the mass of the accessories of the motor and the speed regulator, and the mass of the power supply line. The mass of the accessories of the motor and the speed regulator is taken as 0.3*M motor .

[0097] The additional mass M brought by the distributed propulsion system is EX =0.3*M motor +ρ pl *s pl *l pl , where ρ pl , l pl 、s plThey are the material density of the power line, the length of the power line and the cross-sectional area of ​​the power line, respectively, and can be obtained by referring to the composite wing aircraft of the same level.

[0098] Step 3.4: Calculate the mass of the propulsion system excluding the engine, M ps =M prop +M motor +M GE +M esc +M EX .

[0099] Step 4: Calculate the total energy transmitted by the aircraft at each stage of the flight, convert and calculate the fuel mass and battery mass through the energy efficiency formula, and calculate the total energy mass. Step 4 is implemented in the following steps:

[0100] Step 4.1. Calculate the energy required for the aircraft to perform the flight mission. The energy required in the flight mission comes from aerodynamic resistance, acceleration (kinetic energy), and altitude change (potential energy). The total transmission energy calculation formula of the aircraft is: L / D is the lift-to-drag ratio of the aircraft, t is the cruising time, and m is the takeoff mass of the aircraft.

[0101] Step 4.2, calculate the fuel mass: Since the designed engine power is based on the cruising power, it can be assumed that the engine power remains unchanged during the cruising process, so the fuel mass can be expressed as: where η GE , η mot , η prop , η esc They are generator efficiency, motor efficiency, rotor efficiency, and electronic control efficiency. At the current stage, the corresponding estimates are 0.9, 0.925, 0.80, and 0.95, SFC is the fuel consumption rate, and t is the cruising time.

[0102] Step 4.3, calculate the battery mass: The power of the battery is to supplement the power required by the aircraft during vertical take-off and landing. The formula for calculating the battery mass is: Where E * is the battery density, and the battery density at the current stage is 200wh / kg; η ba is the battery efficiency, P ICE is the engine power, ddq is the depth of discharge inclusion coefficient, and the general value range is 0.2 to 0.25.

[0103] Step 4.4: Calculate the total energy mass M of the aircraft using the fuel mass and battery mass obtained in steps 4.2 and 4.3. energy =M fuel +M ba.

[0104] Step 5: Select a design point on the aircraft design boundary and calculate the initial takeoff mass MTOM of the aircraft according to step 1. 0 And steps 3 and 4 calculate the total energy mass and propulsion system mass, with the total energy mass M energy , propulsion system mass M ps and take-off mass target requirement M payload As the input parameter for the first iteration calculation of the aircraft takeoff mass. Based on the input parameters and the iterative formula, the new takeoff mass MTOM is calculated. NEW , MTOM NEW Substitute steps 3 and 4 into the updated calculation to obtain the total energy mass and propulsion system mass, and convert the takeoff mass target requirement M payload The updated total energy mass and propulsion system mass are used as input parameters for the next iteration until the output value of this iteration is MTOM NEW(n) The output value of the last iteration is MTOM NEW(n-1) Meet | MTOM NEW(n) -MTOM NEW(n-1) |<ε, the iteration is stopped; wherein ε is the set stop threshold, ε is the minimum value, and the value in this embodiment is 0.0001. n is the number of iterations, n>1. Step 5 is specifically implemented according to the following steps:

[0105] Step 5.1. Calculate the parameters required for the first iteration of the aircraft takeoff mass.

[0106] Since the input parameters of the iterative calculation are unknown at the first iteration, that is, the total energy mass and the propulsion system mass are unknown, we first select the design points on the design boundaries of the engine, generator, fixed-wing mode motor and rotor mode motor determined in step 2, and then calculate the initial takeoff mass MTOM according to the initial value of the takeoff mass MTOM in step 1. 0 And the calculation formulas of steps 3 and 4 are used to calculate the total energy mass M energy and propulsion system mass M ps , based on the take-off mass target requirement M payload And the total energy mass M energy and propulsion system mass M ps As input parameter for the first iteration calculation of the aircraft takeoff mass;

[0107] Step 5.2: Calculate the input parameters and the takeoff mass iteration formula based on the first iteration of the takeoff mass

[0108]

[0109] Calculate the takeoff mass output value MTOM for this iteration NEW , put MTOMNEW Substitute into the formulas of step 3 and step 4 to update the propulsion system mass M ps and total energy mass M energy , the takeoff quality target requirement M payload and the updated propulsion system mass M ps and total energy mass M energy It is used as the input parameter for the next iteration until the output value of this iteration is MTOM. NEW(n) The output value of the last iteration is MTOM NEW(n-1) Meet | MTOM NEW(n) -MTOM NEW(n-1) |<ε, the iteration is stopped; ε is the set stop threshold, ε is the minimum value, and ε is taken as 0.0001 in this embodiment; n is the number of iterations, n>1; MF in the aircraft mass iteration formula PS is the vehicle mass coefficient, ranging from 0.45 to 0.55.

[0110] Step 5.3, according to the method of steps 5.1-5.2, traverse all design points on all design boundaries of the aircraft, obtain the output value of the aircraft take-off mass at the iteration cutoff corresponding to each design point, and select the minimum take-off mass as the aircraft take-off mass required for the design.

[0111] Step 6: Calculate the propulsion system mass M according to the parameter information of the design point corresponding to the minimum take-off mass of the aircraft selected in step 5. ps 、Fuel quality M fuel 、Battery quality M ba and wing loading

[0112] The working process of the present invention is further described below with a specific example:

[0113] A vertical take-off and landing hybrid fixed-wing aircraft takes off vertically and climbs to an altitude of 700m to complete the cruise mission. The cruise time is 3h. The aircraft's aspect ratio is 20, the cruise speed is 30m / s, the rotor mode climb rate is 3m / s, the fixed-wing mode climb rate is 3m / s, the ceiling is 1000m, the stall speed is 11m / s, the payload is 15kg, and the fuel consumption rate is 0.3kg*KW -1 *h -1 According to the above design input parameters, it is required to determine the minimum take-off mass, propulsion system mass, fuel mass, battery mass, power-to-weight ratio, wing loading and other parameters of the vertical take-off and landing hybrid fixed-rotor aircraft.

[0114] According to the design input parameters given above, the initial take-off mass value MTOM of the aircraft is set. 0=200kg; draw the design boundaries of the engine, generator, fixed-wing mode motor, and rotor mode motor through step 2; select the design point on the design boundary to obtain the power-to-weight ratio and wing load of the design point, and bring the obtained parameters and the initial value of the take-off mass into the formula of step 3 to calculate the mass of the aircraft propulsion system; calculate the fuel mass and battery mass of the aircraft according to the formula of step 4 through the above selected design points and the obtained propulsion system mass; bring the propulsion system mass, total energy mass, and take-off mass target requirements into the aircraft iteration formula to obtain the first iteration take-off mass output value MTOM NEW , and then perform iterative calculation according to the takeoff mass iteration calculation formula, traverse each design point on the design boundary, and obtain the aircraft takeoff mass output value when the iteration stops corresponding to each design point, such as Figure 4 The figure shows the iterative simulation result of the aircraft takeoff mass in this embodiment. Then the minimum takeoff mass of the aircraft is selected as the aircraft takeoff mass required for the design, and the propulsion system mass, fuel mass, battery mass and wing load are calculated based on the obtained aircraft takeoff mass. The parameters of the aircraft required for the design obtained by iteration in this embodiment are: aircraft takeoff mass MTOM = 116.527kg, wing load Battery mass M ba =11.870kg, fuel mass M fuel =5.239kg, propulsion system mass M ps =47.650kg, such as Figure 5 shown.

[0115] Using the traditional method, without considering the vertical climb constraint of the aircraft and the additional mass added by the distributed propulsion system, the aircraft take-off mass is determined to be MTOM = 108.968 kg, which is smaller than the aircraft mass result estimated by the method of the present invention.

[0116] The above description is only a specific implementation mode of the present invention, but the protection scope of the present invention is not limited thereto. Any technician familiar with the technical field can easily think of various equivalent modifications or substitutions within the technical scope disclosed in the present invention, and these modifications or substitutions should be included in the protection scope of the present invention.

Claims

1. A parameter design method for a hybrid composite wing aircraft, wherein the parameters include aircraft take-off mass, propulsion system mass, fuel mass, battery mass and wing loading; characterized in that: The design method comprises the following steps: Step 1: Determine the takeoff mass target requirement M of the fixed rotor aircraft payload and design input parameters; The design input parameters include cruising speed, rotor mode climb rate, fixed wing mode climb rate, ceiling, stall, lift-to-drag ratio, take-off mass initial value MTOM0 and payload; Step 2: Establish a hybrid composite wing aircraft model. In fixed-wing mode, determine the design boundaries of the aircraft through the aircraft's cruise constraints, climb constraints, ceiling constraints, stall constraints, and vertical climb constraints, and plot each constraint against the power-to-weight ratio. and wing loading and determining the design boundaries of the engine, the generator, the fixed-wing mode motor, and the rotary-wing mode motor according to the relationship curve; Step 3: On the design boundary determined in step 2, select the design points of the engine, generator, fixed-wing mode motor, and rotor mode motor under the same wing load, and calculate the power-to-weight ratio of the selected design point according to the initial value of the takeoff mass and the power-to-weight ratio of the selected design point. and wing loading The mass M of the aircraft's propulsion system is calculated ps , the mass of the propulsion system does not include the mass of the engine; In the process of calculating the mass of the propulsion system, the additional mass brought by the distributed propulsion system is taken into account, and the additional mass includes the mass of the accessories of the motor and the speed regulator, and the mass of the power supply line; Step 4: Calculate the total energy transmitted by the aircraft at each stage of flight, and convert and calculate the total energy mass of the aircraft according to the energy utilization efficiency. The total energy mass is equal to the fuel mass M fuel and battery mass M ba sum; Step 5: Select a design point on the aircraft design boundary and calculate the total energy mass M of the aircraft based on the initial value of the aircraft takeoff mass MTOM0 in step 1 and the process of steps 3 and 4. energy and propulsion system mass M ps , with total energy mass M energy , propulsion system mass M ps and take-off mass target requirement M payload As the input parameter for the first iteration calculation of the aircraft takeoff mass; set the iteration stop threshold, and iterate to obtain the current iteration output value MTOM NEW , and according to the output value MTOM NEW Update the total energy mass and propulsion system mass of the aircraft using the formulas in steps 3 and 4, and convert the takeoff mass target requirement M payload The updated total energy mass and propulsion system mass are used as input parameters for the next iterative calculation; all design points on the aircraft design boundary are traversed to obtain the aircraft takeoff mass corresponding to each design point when the iteration stops, and the minimum takeoff mass is selected as the aircraft takeoff mass required for the design; Step 6: Calculate the mass of the aircraft propulsion system M according to the parameter information of the design point corresponding to the minimum take-off mass selected in step 5. ps 、Fuel quality M fuel 、Battery quality M ba and wing loading 2. The parameter design method according to claim 1, characterized in that: The step 2 comprises the following steps: Step 2.1: Plot the power-to-weight ratio based on the cruise constraint formula for composite wing aircraft and wing loading Cruise constraint curve in coordinates; The cruise constraint formula of the composite wing aircraft is: Where T is thrust, W is the takeoff weight of the aircraft, S is the wing area, and C is D,min is the minimum drag coefficient of the aircraft, C D,min The estimated value is 0.028~0.035; k is the induced drag constant, V is the aircraft speed, η prop is the propeller efficiency, q is the dynamic pressure, ρ is the air density, the standard value is 1.225; Step 2.2, plot the power-to-weight ratio based on the climb constraint formula in fixed-wing mode and wing loading Climb constraint curve in coordinates; The climb constraint formula in the fixed-wing mode is: Where V1 is the climb rate of the aircraft, is the best climb rate; Step 2.3: Plot the power-to-weight ratio based on the stall constraint formula for composite wing aircraft and wing loading Stall curve in coordinates; The stall constraint formula of the composite wing aircraft is: Where V S is the stall speed, C Lmax is the maximum lift coefficient; Step 2.4, according to the ceiling constraint formula Plotting the power-to-weight ratio and wing loading The ceiling constraint curve in coordinates, where V L is the maximum climb rate, and its value is V L =0.508m / s; Step 2.5, according to the vertical climb constraint formula Plotting the power-to-weight ratio and wing loading The vertical climb constraint curve of the aircraft under the coordinates; Among them, S w is the ratio of the aircraft area to the wing area, which is 1.35; FM is the rotor factor, which ranges from 0.6 to 0.7; V I is the axial climbing speed; Step 2.6, according to the cruise constraint curve, climb constraint curve, stall constraint curve, ceiling constraint curve and vertical climb constraint curve drawn in steps 2.1 to 2.5, determine the design boundaries of the engine, generator, fixed-wing mode motor and rotor mode motor; the rules for determining each design boundary are: Since the aircraft adopts a series hybrid system, the power of the engine and generator only needs to meet the cruising power, so the design boundaries of the engine and generator are taken on the cruise constraint curve; since the fixed-wing mode motor needs to provide climb power, the design boundaries of the fixed-wing mode motor are taken on the climb constraint curve; based on the fact that the rotor mode motor needs to provide vertical climb power, the design boundaries of the rotor mode motor are taken on the vertical climb constraint curve; and the fixed-wing mode motor power is greater than the power of the engine or generator.

3. The parameter design method according to claim 1 or 2, characterized in that: The step 3 specifically comprises the following steps: Step 3.1, selecting the design points of the engine, generator, fixed-wing mode motor and rotor mode motor in each design at the same wing load position, calculating the power of the engine, generator, rotor mode motor and fixed-rotor mode motor at the design point, and calculating the total power of the motor; Engine power P ICE =k1*W; generator power P gen =k1*W; Rotor mode motor power P Rmot =k2*W; Motor power P in fixed rotor mode fmot =k3*W; Total motor power P mot =P Rmot +P fmot ; Among them, k1, k2, k3 are the values ​​of power-to-weight ratio corresponding to the selected design points; W is the take-off weight of the aircraft; Step 3.2, calculate the propeller mass, motor mass, generator mass and electric speed governor mass, The propeller mass calculation formula is: The motor mass calculation formula is: The generator mass calculation formula is: M GE =0.385*(P gen +0.44); The mass calculation formula of the electric speed regulator is: Among them, K m is the correction factor, the value is 0.6; n prop is the number of propellers, n b is the number of leaves; is the motor power coefficient, which takes a value of 0.2; is the power coefficient of the electronic speed regulator, and its value is 0.05; Step 3.3, calculate the additional mass brought by the distributed propulsion system, including the mass of the motor and speed regulator accessories and the mass of the power supply line; the mass of the motor and speed regulator accessories is taken as 0.3*M motor ; The additional mass M brought by the distributed propulsion system is EX =0.3*M motor +ρ pl *s pl *l pl , where ρ pl , l pl 、s pl They are the material density of the power line, the length of the power line and the cross-sectional area of ​​the power line; Step 3.4, calculate the mass of the propulsion system excluding the engine M ps ; M ps =M prop +M motor +M GE +M esc +M EX 。 4. The parameter design method according to claim 3, characterized in that: The step 4 specifically comprises the following steps: Step 4.1, calculating the total energy transferred at each stage of the flight according to the energy requirement of the aircraft to perform the mission; the energy requirement comes from aerodynamic drag, acceleration and altitude change potential energy; The total transmission energy of the aircraft is calculated as: Where L / D is the lift-to-drag ratio of the aircraft, t is the cruising time, and m is the takeoff mass of the aircraft; Step 4.2, calculate the fuel mass; Since the designed engine power is based on the cruising power, it is assumed that the engine power remains unchanged during the cruising process, so the fuel quality where η GE , η mot , η prop , η esc are generator efficiency, motor efficiency, rotor efficiency, and ESC efficiency, and their values ​​are 0.9, 0.925, 0.80, and 0.95, respectively. SFC is the fuel consumption rate, and t is the cruising time; Step 4.3, according to the formula Calculate the battery mass M ba , where E * is the battery density, which is 200wh / kg; η ba is the battery efficiency, P ICE is the engine power; ddq is the depth of discharge inclusion coefficient, ranging from 0.2 to 0.25; Step 4.4, calculate the total energy mass M of the aircraft energy =M fuel +M ba .

5. The parameter design method according to claim 4, characterized in that: The step 5 specifically comprises the following steps: Step 5.1, select design points on the design boundaries of the engine, generator, fixed-wing mode motor and rotary-wing mode motor determined in step 2, first obtain the total energy mass M according to the initial takeoff mass value MTOM0 in step 1 and the calculation formulas in steps 3 and 4 energy and propulsion system mass M ps ; Based on the total energy mass M energy , propulsion system mass M ps and take-off mass target requirement M payload As input parameter for the first iteration calculation of the aircraft takeoff mass; Step 5.2, based on the first iteration of the aircraft takeoff mass calculation input parameters and the aircraft takeoff mass iteration formula Iterate the takeoff mass to get the output value MTOM NEW , the current iteration output value MTOM NEW Substitute into steps 3 and 4 to update the total energy mass M energy and propulsion system mass M ps , and the take-off quality target requirement M payload And the updated total energy mass M energy and propulsion system mass M ps As the input parameter of the next iteration, until the output value of this iteration is MTOM NEW(n) The output value of the last iteration is MTOM NEW(n-1) Meet | MTOM NEW(n) -MTOM NEW(n-1) |<ε, the iteration is stopped, where ε is the set stop threshold, n is the number of iterations, and n>1; The aircraft takeoff mass iteration formula MF PS is the aircraft mass coefficient, ranging from 0.45 to 0.55; Step 5.3, according to steps 5.1-5.2, traverse all design points on the aircraft design boundary, obtain the aircraft takeoff mass output value corresponding to each design point when the iteration stops, and select the minimum takeoff mass as the aircraft takeoff mass required for the design.

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