Folding type tilting four-rotor aircraft concept scheme design method based on task requirements

Abstract

The invention relates to a foldable tilting four-rotor aircraft concept scheme design method based on task requirements. According to the method, task demand indexes such as voyage, cruising speed, task load, cruising height and hovering height are used as input, and the total design weight, power device parameters, rotor wing geometric parameters and fuselage and wing parameters of the aircraft are determined in sequence; and performance indexes such as hovering lift limit, level flight speed, climbing performance, practical lift limit, voyage and endurance in a helicopter mode and a fixed wing mode are calculated and analyzed respectively. And performing matching judgment on a performance calculation result and a task requirement, and performing iterative adjustment on key parameters such as a paddle disc load when the task requirement is not met, so as to form a closed-loop concept design process driven by the task requirement. According to the method, the vertical take-off and landing capacity and the fixed wing cruising performance can be systematically considered in the conceptual design stage, and systematicness, repeatability and engineering feasibility of scheme design of the folding type tilting four-rotor aircraft are improved.

Description

Technical Field

[0001] This invention relates to the field of aircraft design methodology, and more specifically, to a design method for a conceptual scheme of a folding tilting quadcopter based on mission requirements. Background Technology

[0002] Rotary-wing aircraft, with their vertical takeoff and landing, hovering, and low-altitude maneuverability, have wide applications in logistics, emergency rescue, and inspection and monitoring. However, traditional helicopters are less efficient in high-speed cruising and long-range missions, while fixed-wing aircraft, although having significant advantages in cruising speed and range, rely on runways for takeoff and landing, making them difficult to adapt to complex terrain and restricted takeoff and landing environments. To balance vertical takeoff and landing capabilities with the cruising efficiency of fixed-wing aircraft, tiltrotor aircraft are gradually becoming an important development direction.

[0003] Existing tiltrotor aircraft mainly include dual-tilt rotor and multi-rotor tiltrotor configurations. Among them, the multi-rotor tiltrotor configuration has certain advantages in terms of structural redundancy, flight safety, and mission adaptability, making it particularly suitable for small and medium-sized unmanned aerial vehicle (UAV) applications. With increasing mission complexity, aircraft not only need sufficient lift and power margins during takeoff, landing, and hovering, but also require high aerodynamic efficiency and range during cruise, which places higher demands on the overall aircraft design.

[0004] In existing technologies, the conceptual design of tiltrotor aircraft typically relies on empirical parameters or local optimizations for a single flight mode. For example, some design methods focus on hovering performance and rotor disk load selection in helicopter mode, while insufficiently considering cruise lift balance, service ceiling, and range and flight time analysis in fixed-wing mode. Other methods are mainly based on fixed-wing aircraft design concepts, making it difficult to simultaneously consider the aerodynamic and power constraints of the rotor during hovering, forward flight, and tilt transitions. Furthermore, in existing design processes, the coupling relationship between rotor parameters, fuselage dimensions, and wing geometry parameters often lacks a systematic quantitative description, leading to repeated manual corrections during the design process and resulting in low design efficiency.

[0005] On the other hand, in mission-driven applications, different missions have significantly different requirements for aircraft, such as hovering altitude, mission payload, cruise speed, range, and service ceiling. However, existing conceptual design methods often fail to establish a closed-loop relationship between "mission requirements—overall parameters—performance indicators—scheme iteration," lacking a systematic design process that starts with mission requirements and uses performance matching as a criterion. This means that while the design may meet one performance indicator, it may not be able to accommodate other key mission requirements, and may even require significant modifications in the later stages of the design process.

[0006] Furthermore, for the configuration of folding tilting multirotor aircraft, there is relatively little research on the conceptual design method of rotor folding, tilting and fixed-wing mode working together in the existing public technology. In particular, there is a lack of a method that can comprehensively consider the total weight of the aircraft, power plant, rotor parameters, fuselage and wing geometry parameters in the conceptual design stage, and systematically calculate and match the performance of helicopter mode and fixed-wing mode.

[0007] Therefore, there is an urgent need for a conceptual design method for a foldable tilting quadcopter based on mission requirements to solve these problems. Summary of the Invention

[0008] The purpose of this invention is to solve the technical problems mentioned in the background section and to provide a design method for a foldable tilting quadcopter based on mission requirements, comprising the following steps: (1) Obtain mission requirement indicators, which include at least range, cruise speed, mission payload, cruise altitude and hovering altitude; (2) Determine the total design weight of the aircraft based on the mission requirements indicators. The total design weight consists of the empty weight, fuel weight and payload. Combine the empty weight ratio range of similar tiltrotor aircraft to determine the empty weight ratio parameter in order to meet the mission payload constraints. (3) Select the power unit, determine the total power requirement of the engine based on the correspondence between the power load parameters and the design total weight, and select the engine model accordingly; (4) Perform rotor parameter design, select disk load as design variable, and determine rotor radius based on total design weight and disk load, while determining the number of blades, rotor solidity and blade tip speed constraints. (5) Design the fuselage and wing parameters. Determine the fuselage size parameters under the premise of satisfying the constraints of no flapping of the front and rear rotor blades and interference. Under the assumption of a rectangular wing, determine the wing span, chord length and front and rear wing area. (6) Calculate the hovering ceiling, level flight speed, climb performance, service ceiling, range and flight time in helicopter mode and fixed-wing mode respectively, and perform a matching analysis between the performance indicators and the mission requirements. When the matching analysis does not meet the mission requirements, return to step (4) to adjust the rotor disk load and iterate the calculation until the mission requirements are met, and then output the conceptual design result.

[0009] As a preferred technical solution of the present invention: in step (2), the total design weight is determined by the weight combination relationship between the empty weight, fuel weight and payload; the empty weight is obtained by the ratio relationship between the empty weight ratio parameter and the total design weight, wherein the empty weight ratio parameter is selected from the statistical range of similar tiltrotor aircraft.

[0010] As a preferred technical solution of the present invention: In step (2), the maximum task load in the task requirements is used as a constraint condition, and the total design weight is determined under the premise of satisfying the maximum task load constraint, so as a unified input for subsequent power plant selection, rotor parameter design and wing parameter design.

[0011] As a preferred technical solution of the present invention: In step (3), the power load parameter characterizes the weight capacity that a unit engine power can bear under standard atmospheric conditions. Based on the power load parameter, the design total weight, and the performance requirements for vertical take-off and landing, hovering and forward flight, the total engine power requirement is determined, and the engine model is selected accordingly.

[0012] As a preferred technical solution of the present invention: In step (4), the rotor disk load is selected according to the hovering ceiling performance requirements and is used as an adjustable design variable; after determining the rotor disk load, the rotor radius is determined according to the correspondence between the total design weight and the rotor disk load; when the performance matching analysis does not meet the mission requirements, the rotor disk load is iteratively adjusted.

[0013] As a preferred technical solution of the present invention: In step (4), blade tip speed constraints are set for helicopter mode and forward flight mode respectively. In helicopter mode, blade tip speed is determined by combining noise limit, stall limit and compressibility limit. In forward flight mode, blade tip resultant speed is obtained by combining rotor rotation speed component and forward flight speed component, and blade tip shock wave is avoided by upper limit constraint related to the speed of sound corresponding to cruise altitude.

[0014] As a preferred technical solution of the present invention: in step (4), the number of blades of a single rotor is selected as three, and the rotor solidity is determined according to the rotor disk load and hovering performance requirements; the rotor solidity is used to constrain the aerodynamic capability of the rotor to ensure sufficient lift margin at the designed hovering height.

[0015] As a preferred technical solution of the present invention: In step (5), the fuselage length is set to be greater than the rotor diameter to meet the structural constraint that the front and rear rotor blades do not flap, and the fuselage length, width and height are determined in combination with the mission loading requirements. The determination of the fuselage size is also used to reduce the aerodynamic interference effect between the front and rear rotors.

[0016] As a preferred technical solution of the present invention: In step (5), the wing adopts a rectangular wing configuration, and the wing geometric parameters are determined according to the preset aspect ratio; wherein the wing span is determined by the rotor radius, fuselage width and dimensional margin, the wing chord length is determined by the wing span and aspect ratio, and the front and rear wing areas are determined by their respective spans and chord lengths; at the cruising speed, the lift parameters are calculated according to the balance relationship between the wing area and the aircraft weight.

[0017] As a preferred technical solution of the present invention: In step (6), in helicopter mode, by comparing the difference between the available power of the engine and the power required by the rotor, the hovering ceiling, level flight speed and climb performance are determined respectively, and the corresponding altitude or speed limit is determined when the remaining power approaches zero. In fixed-wing mode, the service ceiling of the aircraft is determined by calculating the climb performance under given weight and engine conditions, and when the rate of climb decreases to a preset threshold close to zero. During the cruise phase, the remaining time is determined based on the available fuel and the average fuel consumption during the cruise phase, and the cruise range is determined based on the remaining time and the cruise speed, which is used to match and determine the range and flight time indicators in the mission requirements.

[0018] Compared with the prior art, the present invention has the following beneficial effects: This invention proposes a conceptual design method for a folding tiltrotor quadcopter based on mission requirements. Starting with mission requirements, it directly incorporates mission constraints such as range, cruise speed, payload, and hovering altitude into the overall parameter determination process, establishing a systematic design path from mission requirements to the overall design weight, power plant, rotor parameters, and fuselage and wing parameters. Compared to existing conceptual design methods that rely on experience or single flight mode optimization, this method can simultaneously consider multiple flight conditions such as vertical takeoff and landing, hovering, and fixed-wing cruise during the conceptual design phase. This makes the design results closer to actual mission requirements and avoids the problem of repeated modifications due to unclear mission assumptions or arbitrary parameter selection.

[0019] This invention introduces a closed-loop design mechanism of "performance calculation—mission matching—parameter iteration" by treating propeller disk load as a key design variable. This mechanism unifies and coordinates the coupling relationships between rotor parameters, power requirements, and fuselage and wing geometry. When a design scheme fails to meet mission requirements in terms of hovering ceiling, level flight speed, climb performance, or fixed-wing service ceiling, range, and flight time, the scheme can be quickly corrected and optimized by adjusting the propeller disk load and updating the rotor radius and related geometric parameters. This improves the repeatability and engineering feasibility of the conceptual design and reduces reliance on the designer's experience.

[0020] Furthermore, this invention systematically calculates and analyzes key performance indicators for both helicopter and fixed-wing modes simultaneously during the conceptual design phase, making the performance boundaries of the folding tiltrotor aircraft clearly identifiable in both flight modes. This method helps identify potential performance bottlenecks early in the design process, rationally balances the relationship between hovering capability and cruise efficiency, improves the aircraft's overall adaptability in complex mission scenarios, and provides a reliable overall parameter basis for subsequent detailed design and engineering implementation. Attached Figure Description

[0021] Figure 1 Flowchart for the design of a concept scheme for a foldable tilting quadcopter based on mission requirements; Figure 2 Flowchart for calculating the hovering ceiling of a folding tilting quadcopter. Detailed Implementation

[0022] To make the objectives, technical solutions, and advantages of this invention clearer, the following description is provided in conjunction with embodiments and appendices. Figures 1-2 The present invention will be further described in detail below. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention.

[0023] This invention proposes a conceptual design method for a folding tiltrotor quadcopter based on mission requirements. Taking mission requirements as input, it sequentially completes the following steps: determining the total weight of the aircraft, selecting the power plant, designing rotor parameters, designing fuselage / wing parameters, calculating multi-mode performance indicators, and analyzing the compatibility of the design with mission requirements. If the requirements are not met, key design parameters are iteratively updated until a conceptual design that satisfies the mission is output. The overall logic is shown in the disclosure materials. Figure 1 The process is consistent: first, determine the mission requirements; then, determine the total weight and power of the aircraft; next, design the geometric parameters of the rotor, fuselage, and wings; then, calculate the key performance indicators for helicopter and fixed-wing modes; finally, determine the mission matching; if the requirements are not met, return to the rotor design stage to adjust the rotor disk load and perform iterative calculations.

[0024] Step 1: Determine the total weight of the aircraft based on usage requirements; In this implementation, mission requirement indicators are first given or obtained for the target mission scenario, including: range, cruise speed, mission payload, cruise altitude, and hovering altitude. Mission payload refers to the "sum of payload and fuel weight," used to constrain the aircraft's transport / endurance capability under a given mission profile.

[0025] Based on the above mission requirements, a mass decomposition model of the aircraft's total design weight is established. The total design weight of the aircraft consists of the empty weight, fuel weight, and payload, satisfying the following relationship: (1) in, Design the overall weight of the aircraft; This is the empty weight of the machine; This refers to the weight of the fuel. This refers to the effective payload weight (including mission equipment, cargo / personnel, etc.). During the conceptual design phase, the above quantities can be consistently referred to as "weight" (in engineering, they can be treated as equivalent to weight / mass, maintaining the same standard).

[0026] To enable the design total weight to be derived from the "mission payload" during the conceptual design phase, this implementation further introduces the empty weight ratio (the proportion of empty weight to total weight) and denoted as . By statistically analyzing the range of empty weight ratios of similar tiltrotor / rotor aircraft, a suitable one for this mission was selected. And taking the maximum payload requirement (i.e., maximum mission load) from the mission requirements as the design constraint, we obtain the inverse expression for the total design weight of the aircraft: (2) in, Design the overall weight of the aircraft; This refers to the weight of the fuel. The effective payload weight; The empty weight ratio (dimensionless) represents the proportion of the empty weight to the total design weight.

[0027] By combining equations (1) and (2), this implementation method forms a closed loop of "task load - empty weight ratio - design gross weight" in the conceptual stage: first, the task load and experience are combined. get Then, the weight composition of the empty engine, fuel and load is decomposed and verified by equation (1) to provide a unified total weight input for subsequent power and rotor design.

[0028] Step Two: Power Unit Selection; Obtain the total design weight Next, the power unit is selected. This implementation method uses "power load" for rapid estimation of the engine's total power. Power load is defined as: the weight that the engine's rated power can lift under standard atmospheric conditions at sea level, satisfying: (3) in, Engine power load (weight / power ratio parameter); Design the overall weight of the aircraft; This refers to the total engine power (the total usable power under standard atmospheric conditions at sea level).

[0029] During implementation, the target power load is selected based on the empirical range of power loads for similar rotorcraft / tilt rotorcraft, and taking into account the mission's requirements for vertical takeoff and landing, level flight, and hovering performance. Then, the total engine power requirement is calculated by back-calculating using equation (3). Based on this, engine models and power system schemes that meet power level, weight and installation constraints are selected (including conceptual matching such as transmission, tilting mechanism drive / power supply interface, etc.), thereby ensuring that the available power supply boundary is clear in subsequent performance calculations.

[0030] Step 3: Rotor parameter design; Folding tiltrotor quadcopters need to meet hovering and vertical performance requirements in helicopter mode, and forward propulsion efficiency in airplane mode. Therefore, rotor parameters must be comprehensively balanced between hovering and forward flight. This implementation first selects the rotor disk load based on the mission's performance requirements for hovering ceiling and other parameters, referring to the empirical range of rotor disk loads for conventional helicopters and tiltrotor aircraft. In subsequent performance calculations, if the proposed solutions do not meet the task requirements, then... Make iterative adjustments until the task requirements are met.

[0031] Sure and Then, the radius of a single rotor is derived from the following formula: (4) in, For propeller disk load; Design the overall weight of the aircraft; Pi is a constant. The rotor radius of the four tiltrotor aircraft; "4" indicates that the total disk area of ​​the four rotors is [missing information]. .

[0032] Subsequently, the blade tip speed constraints were determined. In helicopter mode, the blade tip speed is mainly affected by noise, stall, and compressibility limits (empirically, a range of 210-240 m / s is used as the basis for selection). In aircraft mode, during forward flight, to reduce drag power and improve propulsion efficiency, the blade tip speed should be appropriately reduced. When the tiltrotor aircraft is flying forward at speed V0, the resultant blade tip speed is approximately: (5) in, The tip velocity; The tip velocity component related to rotor rotation during forward flight (which can be understood as "tip velocity component caused by rotation"). This is the forward velocity of the aircraft. To avoid shock waves generated at the propeller tips, Mach number constraints are generally satisfied. ,in The speed of sound at cruising altitude. Furthermore, this can be achieved at the maximum speed required by the forward flight design. Determine Upper limit: (6) in, This represents the upper limit of the propeller tip velocity component in forward flight mode; The speed of sound at cruising altitude; The maximum forward flight speed is given for mission requirements. Based on this, this implementation selects the rotor solidity according to the empirical relationship between rotor disk load and rotor solidity, combined with hovering performance requirements. Simultaneously select the number of rotor blades. A single rotor typically uses a configuration of three blades, which is commonly used in conceptual designs. Based on the selected... , and Calculate the mean chord length of the rotor: (7) in, The average chord length of the rotor blades; For rotor real number (dimensionless); Pi is a constant. The rotor radius; This refers to the number of blades on a single rotor.

[0033] At this point, step three outputs the set of key concept parameters for the rotor. , , upper limit , , (etc.), and "propeller disk load" "As the primary parameter for subsequent iterations: when performance is not up to standard, adjust..." And thus change Based on the relevant power requirements, perform step five calculations and step six matching judgments again.

[0034] Step 4: Design of fuselage and wing parameters; After obtaining the aircraft's overall design weight and rotor geometry parameters, conceptual-level geometric design of the fuselage and wings is carried out. The fuselage section focuses on ensuring: no rotor blade flapping, minimizing front and rear interference effects, and meeting mission loading space requirements. During implementation, the fuselage length, width, and height are determined (corresponding to the requirements for fuselage length, width, and height in the disclosure materials). The wing section adopts a rectangular wing in the conceptual stage to quickly establish lift-drag estimation and structural layout constraints. The wing aspect ratio λ is initially determined based on existing tiltrotor aircraft parameters, and the wingspan L is determined based on rotor radius and fuselage width constraints. This implementation provides the geometric constraint expression for the wingspan: (8) Where L is the wingspan; R is the rotor radius; Width of the fuselage; This is a dimensional margin reserved to prevent rotor blade flapping. Given the wing aspect ratio λ and span L, the wing chord length is calculated using the following formula: (9) Where c is the wing chord length; L is the wing span; and λ is the wing aspect ratio (dimensionless). Furthermore, in this embodiment, the areas of the forewing and aft wing are calculated separately, satisfying: (10) in, This refers to the area of ​​the forewing; For the forward wing span; This refers to the chord length of the forewing.

[0035] (11)

[0036] in, This refers to the area of ​​the rear wing; For the rear wing span; This refers to the rear wing chord length.

[0037] To ensure lift balance at cruise speed, the lift coefficient is determined from the lift balance relationship under cruise speed conditions: (12) in, Air density (related to cruising altitude); Cruise speed; This refers to the area of ​​the forewing; This refers to the area of ​​the rear wing; The lift coefficient during cruise. It is gravity.

[0038] Step 4: Output the conceptual parameters for the fuselage and wings. (etc.), together with the rotor parameters in step three, constitute the geometric basis for the two types of aerodynamic / power estimation: "helicopter mode" and "fixed-wing mode".

[0039] Step 5: Calculate the performance indicators of the design scheme; This step calculates key performance indicators for the same set of conceptual parameters in both helicopter and fixed-wing modes, and compares and matches them with mission requirements.

[0040] (a) Calculation of performance indicators in helicopter mode: The folding tiltrotor aircraft adopts helicopter mode during takeoff and vertical flight, with the four rotors providing lift. At this time, the key indicators affecting mission capability include hovering ceiling, level flight speed, and climb performance.

[0041] 1) Hovering Ceiling: The climb rate is approximately 0 during hovering. During implementation, a hovering weight (consistent with the mission load state) is given, and the following steps are followed: Figure 2 The process uses hovering altitude as an iterative variable to calculate step by step: the required power of the rotor system and the available power of the engine at that altitude, to obtain the remaining power. When the remaining power is approximately equal to 0, the corresponding altitude is the hovering ceiling for the current gross weight state. This process ensures that the power margin is determined under different altitude density conditions, and obtains a feasible conclusion that meets the mission hovering altitude requirements.

[0042] 2) Level Flight Speed: Based on three types of boundaries—available power, airflow separation limitation, and shock wave limitation—the maximum level flight speed boundaries limited by power, separation, and shock waves are obtained respectively. The minimum value of the three is taken as the comprehensive boundary for the maximum level flight speed in helicopter mode, and the maximum level flight speed at any altitude is obtained accordingly. This process ensures that the three types of constraints—power, aerodynamics, and compressibility—are satisfied simultaneously.

[0043] 3) Climbing Performance: This mainly analyzes the diagonal climbing capability while maintaining forward speed. During implementation, a calculation altitude and an economical speed point (usually a speed point with lower power requirements) are selected. Power requirements and available power are calculated from altitude and speed, and the remaining power is calculated and converted into the climb rate. As altitude increases, when the climb rate falls below a set threshold close to 0, the climb limit is considered reached. This method utilizes the power margin advantage of the economical speed point to obtain an assessment of maximum climbing capability.

[0044] (II) Calculation of performance indicators in fixed-wing mode: After the aircraft reaches a certain speed and climbs to a certain altitude, the rotor folds and tilts 90°, converting it into a fixed-wing aircraft; the lift and drag characteristics are determined with reference to the fixed-wing aircraft method. At this time, the key indicators include service ceiling, range, and flight time.

[0045] 1) Service Ceiling: The service ceiling is defined as the altitude at which, given the aircraft weight and engine conditions, the maximum rate of climb in subsonic flight decreases to a relatively small value close to zero. In practice, the remaining power is calculated based on the required and available power in fixed-wing mode, and then converted into a fixed-wing rate of climb. As altitude increases, the service ceiling is reached when the rate of climb falls below a set value. This process is consistent with the description in the handover materials, ensuring an objective assessment of power margin in fixed-wing mode.

[0046] 2) Range and flight time: First, calculate the available fuel quantity during the cruise phase: (13) in, This refers to the amount of fuel available during the cruise phase. This refers to the total fuel consumption of the aircraft. This represents the fuel consumption during the ascent and descent phases, as well as the proportion of dead fuel and remaining fuel (dimensionless proportionality coefficient).

[0047] Next, calculate the average hourly fuel consumption during the cruising phase: (14) in, This represents the average hourly fuel consumption. For cruise drag; Cruise speed (unit: m / s); 0.001 is the unit conversion factor; Engine fuel consumption rate (unit: g / kWh).

[0048] The driving range is calculated based on the amount of fuel available during cruising and the average hourly fuel consumption: (15) in, Battery life (h); This refers to the amount of fuel available during the cruise phase. This represents the average hourly fuel consumption. Further, the cruising distance is obtained: (16) in, The cruising distance (m); Battery life (h); The cruise speed is 3600 (m / s); 3600 is the conversion factor between hours and seconds (1h = 3600s).

[0049] Through equations (13) to (16), this embodiment realizes a calculable chain of "fuel consumption rate - flight time - range", which directly corresponds to the range / cruising speed requirement in step one, thereby providing a quantitative basis for task matching in step six.

[0050] Step Six: Analysis and Iteration of the Matching between the Solution and Task Requirements; The helicopter mode indicators (hovering ceiling, maximum level flight speed, climb ceiling) and fixed-wing mode indicators (service ceiling, flight time, range) obtained in step five are compared item by item with the mission requirements given in step one to form a matching conclusion: if all are satisfied, the conceptual scheme parameter set of the current folding tiltrotor aircraft (including...) is output. Power (e.g., rotor geometry and tip speed constraints, fuselage / wing geometry and cruise lift balance). If any of these parameters are not met, return to step three, using the rotor disk load... Reselect and update the rotor radius as the main iteration variable. (Updated in conjunction with Equation (4), which in turn affects the geometric constraints of the fuselage / wing (such as the span constraint in Equation (8)) and the calculation results of performance indicators. Steps three to six are executed in a loop until the mission requirements are met. This strategy of "using propeller disk load as the iterative lever" is consistent with the description in the briefing materials, which can ensure that the iteration path is clear, reproducible and strongly correlated with the mission requirements.

[0051] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A design method for a conceptual scheme of a folding tilting quadrotor aircraft based on mission requirements, characterized in that, Includes the following steps: (1) Obtain mission requirement indicators, which include at least range, cruise speed, mission payload, cruise altitude and hovering altitude; (2) Determine the total design weight of the aircraft based on the mission requirements indicators. The total design weight consists of the empty weight, fuel weight and payload. Combine the empty weight ratio range of similar tiltrotor aircraft to determine the empty weight ratio parameter in order to meet the mission payload constraints. (3) Select the power unit, determine the total power requirement of the engine based on the correspondence between the power load parameters and the design total weight, and select the engine model accordingly; (4) Perform rotor parameter design, select disk load as design variable, and determine rotor radius based on total design weight and disk load, while determining the number of blades, rotor solidity and blade tip speed constraints. (5) Design the fuselage and wing parameters. Determine the fuselage size parameters under the premise of satisfying the constraints of no flapping of the front and rear rotor blades and interference. Under the assumption of a rectangular wing, determine the wing span, chord length and front and rear wing area. (6) Calculate the hovering ceiling, level flight speed, climb performance, service ceiling, range and flight time in helicopter mode and fixed-wing mode respectively, and perform a matching analysis between the performance indicators and the mission requirements. When the matching analysis does not meet the mission requirements, return to step (4) to adjust the rotor disk load and iterate the calculation until the mission requirements are met, and then output the conceptual design result.

2. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (2), the total design weight is determined by the weight combination relationship between the empty weight, fuel weight and payload; the empty weight is obtained by the ratio relationship between the empty weight ratio parameter and the total design weight, wherein the empty weight ratio parameter is selected from the statistical range of similar tiltrotor aircraft.

3. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (2), the maximum task load in the task requirements is used as a constraint, and the total design weight is determined under the premise of satisfying the maximum task load constraint, so as a unified input for subsequent power plant selection, rotor parameter design and wing parameter design.

4. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (3), the power load parameter represents the weight capacity that a unit engine power can bear under standard atmospheric conditions. Based on the power load parameter, the design total weight, and the performance requirements for vertical takeoff and landing, hovering and forward flight, the total engine power requirement is determined, and the engine model is selected accordingly.

5. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (4), the rotor disk load is selected based on the hovering ceiling performance requirements and is used as an adjustable design variable. After determining the rotor disk load, the rotor radius is determined based on the correspondence between the design total weight and the rotor disk load. When the performance matching analysis does not meet the mission requirements, the rotor disk load is iteratively adjusted.

6. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (4), blade tip speed constraints are set for helicopter mode and forward flight mode respectively. In helicopter mode, blade tip speed is determined by combining noise limit, stall limit and compressibility limit. In forward flight mode, blade tip resultant speed is obtained by combining rotor rotation speed component and forward flight speed component, and blade tip shock wave is avoided by upper limit constraint related to the speed of sound corresponding to cruise altitude.

7. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (4), the number of blades of a single rotor is selected as three, and the rotor solidity is determined according to the rotor disk load and hovering performance requirements; the rotor solidity is used to constrain the aerodynamic capability of the rotor to ensure sufficient lift margin at the designed hovering height.

8. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (5), the fuselage length is set to be greater than the rotor diameter to meet the structural constraint of the front and rear rotor blades not flapping. The fuselage length, width and height are determined in combination with the mission loading requirements. The determination of the fuselage size is also used to reduce the aerodynamic interference effect between the front and rear rotors.

9. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (5), the wing adopts a rectangular wing configuration, and the wing geometric parameters are determined according to the preset aspect ratio; the wing span is determined by the rotor radius, fuselage width and dimensional margin, the wing chord length is determined by the wing span and aspect ratio, and the front and rear wing areas are determined by their respective spans and chord lengths; at the cruise speed, the lift parameters are calculated based on the balance relationship between the wing area and the aircraft weight.

10. The design method for a folding tilting quadcopter concept based on mission requirements according to claim 1, characterized in that: In step (6), In helicopter mode, the hovering ceiling, level flight speed and climb performance are determined by comparing the difference between the engine's available power and the rotor's required power, and the corresponding altitude or speed limit is determined when the remaining power approaches zero. In fixed-wing mode, the service ceiling of the aircraft is determined by calculating the climb performance under given weight and engine conditions, and when the rate of climb decreases to a preset threshold close to zero. During the cruise phase, the remaining time is determined based on the available fuel and the average fuel consumption during the cruise phase, and the cruise range is determined based on the remaining time and the cruise speed, which is used to match and determine the range and flight time indicators in the mission requirements.

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