A method of distributed tail volume design for a propulsion aircraft and a computing device

Abstract

A tail capacity design method of a distributed propulsion aircraft and a computing device, according to the overall design parameters of the aircraft and the target take-off rolling distance, the take-off speed and the thrust demand corresponding to the take-off limit trim state are determined; according to the thrust demand and the take-off speed, the induced speed of the aircraft rotor position is determined; according to the take-off speed, the first dynamic pressure of the wing non-slip flow area is determined, and according to the induced speed, the second dynamic pressure of the wing slip flow area is determined; according to the first dynamic pressure and the second dynamic pressure, the wing aerodynamic moment is determined; according to the wing aerodynamic moment, the horizontal tail trim moment corresponding to the take-off limit trim state is determined; according to the horizontal tail trim moment, the tail capacity design parameters of the aircraft are determined, without the help of high-fidelity CFD simulation model to solve the flow field, only the overall layout parameters can quickly obtain more accurate wing aerodynamic moment in the conceptual design stage by using analytical calculation, and the tail capacity design parameters of the aircraft are determined.

Description

Technical Field

[0001] The embodiments in this specification belong to the field of data processing technology, and in particular relate to a tail capacity design method and computing device for a distributed propulsion aircraft. Background Technology

[0002] Tail capacity is a core parameter in the longitudinal static stability design of an aircraft, defined as the dimensionless ratio of the product of the horizontal stabilizer area and the tail arm length to the product of the wing area and the mean aerodynamic chord length. Tail capacity measures the weight of the pitching moment generated by the horizontal stabilizer behind the aircraft's center of gravity relative to the aerodynamic loads on the wing: if the tail capacity is too small, the horizontal stabilizer will not be able to provide sufficient longitudinal restoring and control moments; if the value is too large, the weight cost of the horizontal stabilizer and tail section structure will exceed the necessary level. In the conceptual design phase of conventional fixed-wing aircraft, tail capacity is usually determined based on statistical experience ranges of similar aircraft types (e.g., 0.5 to 1.0 for general aviation aircraft). This empirical method assumes that all aerodynamic surfaces of the aircraft are in a free flow with basically uniform dynamic pressure—the dynamic pressure felt by the wing and horizontal stabilizer is similar, and there is no structural deviation of order of magnitude between trim requirements and trim capabilities.

[0003] Distributed electric propulsion (DEP) ultra-short take-off and landing (USTOL) aircraft are a new type of configuration that has attracted widespread attention in recent years. They utilize multiple small propellers arranged along the wing span to improve low-speed aerodynamic characteristics by leveraging the influence of propeller slipstreams on the airflow around the wing. Combined with high-lift devices, they can achieve ultra-short take-off and landing (USTOL). In the design practice of such aircraft, determining the tail capacity presents a particular challenge: the effective range of the propeller slipstream is concentrated near the wing, and there is a significant difference in dynamic pressure between the flow field at the wing and the flow field at the horizontal stabilizer. Furthermore, the degree of this difference varies drastically with changes in thrust level, flight speed, and propeller configuration parameters.

[0004] For this type of configuration, the current engineering practice is to first use the empirical tail capacity coefficient under the uniform flow assumption as the initial design value, and then use high-fidelity computational fluid dynamics (CFD) numerical simulation to check and correct the trim capability. The limitation of this approach is that CFD simulation consumes a lot of computational resources and has a long single-run solution cycle. During the conceptual design phase, the overall parameters of the aircraft—including takeoff weight, wing area, number and layout of propellers, center of gravity position, etc.—are frequently adjusted. Each adjustment requires re-gridning, resetting boundary conditions, and iterating again until convergence. The simulation time of several hours or even longer per run presents a significant contradiction with the design rhythm required for rapid response to changes in the conceptual phase. In the absence of quantitative and rapid evaluation methods, the tail capacity design approach relying on intensive CFD iterations is difficult to effectively support the efficient screening and comparison of a large number of alternative layout schemes during the conceptual design phase.

[0005] Therefore, this specification provides a tail capacity design scheme for a distributed propulsion aircraft to at least partially solve the above-mentioned problems. Summary of the Invention

[0006] This specification provides an embodiment of a tail capacity design method and computing device for a distributed propulsion aircraft, including:

[0007] The first aspect of this specification provides a tail capacity design method for a distributed propulsion aircraft, including:

[0008] Based on the overall design parameters of the aircraft and the target takeoff run distance, determine the takeoff speed and thrust requirements corresponding to the takeoff limit trim state;

[0009] Based on the thrust requirement and the takeoff speed, determine the induced velocity for the aircraft rotor disk position;

[0010] The first dynamic pressure in the non-slipstream region of the wing is determined based on the takeoff speed, and the second dynamic pressure in the slipstream region of the wing is determined based on the induced velocity.

[0011] The wing aerodynamic torque is determined based on the first dynamic pressure and the second dynamic pressure;

[0012] Based on the wing aerodynamic moment, determine the horizontal stabilizer trim moment corresponding to the takeoff limit trim state;

[0013] The tail capacity design parameters of the aircraft are determined based on the horizontal stabilizer trim moment.

[0014] A second aspect of this specification provides a computer-readable storage medium having a computer program stored thereon that, when executed in a computer, causes the computer to perform the method described in the first aspect.

[0015] A third aspect of this specification provides a computing device including a memory and a processor, wherein the memory stores executable code, and the processor, when executing the executable code, implements the method described in the first aspect.

[0016] The tail capacity design technology for a distributed propulsion aircraft provided in this embodiment does not require the use of a high-fidelity CFD simulation model to solve the flow field. Only the overall layout parameters are needed. During the conceptual design stage, analytical calculations can be used to quickly obtain a relatively accurate wing aerodynamic torque and determine the tail capacity design parameters of the aircraft accordingly. Attached Figure Description

[0017] To more clearly illustrate the technical solutions of the embodiments in this specification, the drawings used in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in this specification. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0018] Figure 1 This is a schematic diagram of a distributed propulsion ultra-short takeoff and landing aircraft. Figure 2 This is a flowchart of a tail capacity design method for a distributed propulsion aircraft in one embodiment of this specification; Figure 3 This is a schematic diagram showing the positional relationship between the propeller disk and the leading edge of the wing; Figure 4 This is a schematic diagram showing the distribution of slipstream and non-slipstream areas; Figure 5 This is a schematic diagram of the structure of a distributed propulsion aircraft tail capacity design device in one embodiment of this specification. Detailed Implementation

[0019] To enable those skilled in the art to better understand the technical solutions in this specification, the technical solutions in the embodiments of this specification will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this specification, and not all embodiments. Based on the embodiments in this specification, all other embodiments obtained by those skilled in the art without creative effort should fall within the scope of protection of this specification.

[0020] In this embodiment, the tail capacity design of the distributed electric propulsion ultra-short takeoff and landing aircraft is a sub-stage in the overall conceptual design phase, which is usually carried out after the design parameters of other parts of the aircraft (fuselage, wings, propeller layout, etc.) are initially determined.

[0021] Figure 1 A schematic diagram of a distributed propulsion ultra-short takeoff and landing (UST) aircraft is shown. Figure 1 As shown, multiple small electric propellers are arranged along the spanwise direction on or above the leading edge of the aircraft's wing. Figure 1 (Illustrative diagram of the circular area above the wing). During takeoff, each propeller operates at full power, generating a high-speed slipstream that washes over the wing surface. The flaps on the trailing edge of the wing deflect at a large angle to utilize the lift-enhancing effect of the slipstream to generate high lift. The horizontal stabilizer is mounted at the rear of the fuselage, outside the area covered by the propeller slipstream. Thanks to the aforementioned distributed electric propulsion system (multiple small electric propellers arranged along the wing span), the takeoff speed and takeoff distance of the distributed propulsion ultra-short takeoff and landing aircraft are significantly lower than those of aircraft using conventional propulsion systems.

[0022] The primary purpose of tail capacity design is to balance the overall aircraft torque and prevent the aircraft from losing control under various conditions. This embodiment mainly considers the takeoff limit trim state (or takeoff critical state). On one hand, under the takeoff limit trim state, the lift generated by the wing section produces a large nose-down moment. On the other hand, compared to flight, the aircraft speed is lower under the takeoff limit trim state, and the effectiveness of traditional control surfaces (such as elevators) decreases sharply; moreover, the horizontal stabilizer is far from the propeller at the wing and is in a free flow without acceleration, making it difficult for the dynamic pressure at the horizontal stabilizer to provide sufficient trim torque. Therefore, it is necessary to design the horizontal stabilizer (i.e., tail capacity design) to provide the corresponding trim torque to maintain the aircraft's attitude and prevent attitude instability and loss of control.

[0023] In the longitudinal plane, the pitching moment acting on an aircraft can be decomposed into several main sources. One is the wing aerodynamic moment, which originates from the moment of the resultant force of the pressure distribution on the wing surface relative to the center of gravity. When the flaps deflect at a large angle, the center of lift shifts rearward, and this moment typically manifests as a nose-down direction. The aerodynamic moment generated by the horizontal stabilizer is the main means of counteracting the aforementioned nose-down moment and maintaining takeoff attitude balance—the deflection of the elevators on the horizontal stabilizer changes the magnitude and direction of the lift, thereby generating a nose-up moment behind the center of gravity. The thrust line of the distributed propeller is generally located below or above the center of gravity. The thrust vector and the lever arm from the thrust point of application to the center of gravity generate a thrust offset moment, which also needs to be included in the overall aircraft moment balance.

[0024] In one embodiment of this specification, the tail volume coefficient is... Defined as flat tail area With tail arm length The product divided by the wing reference area With mean aerodynamic chord length The product of, i.e.

[0025]

[0026] Since both the numerator and denominator have volume dimensions, this ratio is a dimensionless coefficient. Tail capacity reflects the weight of the pitching moment contributed by the horizontal stabilizer relative to the aerodynamic load on the wing: the larger the tail capacity, the stronger the trim capability of the horizontal stabilizer, but at the same time it means a larger horizontal stabilizer size or a longer tail section structure, which in turn brings greater structural weight and wetted area drag.

[0027] In this embodiment, the wing reference area is used for tail capacity design. With mean aerodynamic chord length All of these have been determined in the preliminary design steps. Specifically, the design parameter to be optimized in this embodiment is the horizontal tail area. With tail arm length .

[0028] Before introducing the various embodiments of this specification, the design concept of this embodiment will be briefly explained first.

[0029] For distributed electric propulsion ultra-short takeoff and landing (UST) aircraft, the incoming airflow experiences a velocity jump as it passes over the propeller disk during operation. According to momentum theory, this velocity increment at the propeller disk (i.e., induced velocity) is... ) and thrust T, incoming flow velocity and propeller disk area There is a deterministic relationship between them. For a given thrust and incoming flow velocity, the larger the disk area, the smaller the thrust load per unit area, and the smaller the induced velocity. The induced velocity is superimposed on the incoming flow velocity, making the airflow velocity downstream of the disk greater than the free incoming flow velocity. The increase in airflow velocity leads to an increase in dynamic pressure, which is determined by the following formula: ,in Here, V represents atmospheric density, and V represents the local point (a specific local location on the aircraft surface) airflow velocity, which is affected by induced velocity. Therefore, the dynamic pressure experienced by the wing surface region within the propeller wake (slipstream region) is significantly higher than that experienced only by the free flow (non-slipstream region).

[0030] However, existing tail capacity design methods largely follow the parametric design approach of traditional propulsion systems for aircraft. In traditional propulsion systems, the engine (propeller) is usually located in a pod under the wing, and the resulting slipstream can only cover a small portion of the inner side of the wing, or even not at all. Therefore, traditional tail capacity design methods almost do not consider the influence of slipstream on pitching moment, and only perform aerodynamic analysis of the wing and horizontal stabilizer based on the incoming flow velocity.

[0031] In the takeoff state, the incoming flow velocity It is very small (close to the stall speed of the takeoff configuration), while the thrust is... Due to the requirements of ultra-short takeoff, the induced velocity is very large, in order to generate the corresponding thrust-induced velocity. The values ​​are extremely high. This means that the dynamic pressure in the slipstream region of a distributed electric propulsion (EPP) VSTOL aircraft can be several times that of the non-slipstream region. This magnitude of dynamic pressure difference also makes traditional tail capacity design methods ineffective for distributed electric propulsion configurations. Therefore, this embodiment proposes a tail capacity design method suitable for distributed electric propulsion configurations.

[0032] Figure 2 A flowchart illustrating a tail capacity design method for a distributed propulsion aircraft according to one embodiment of this specification is shown below. Figure 1 The description will be presented step by step.

[0033] Step S200: Based on the overall design parameters of the aircraft and the target takeoff run distance, determine the takeoff speed and thrust requirements corresponding to the takeoff limit trim state.

[0034] First, obtain the overall design parameters of the aircraft and the target takeoff distance, and then use this information to solve for the takeoff speed and thrust requirements under the worst operating conditions—the takeoff limit trim state.

[0035] The overall design parameters include takeoff weight (m) and wing reference area. and the maximum lift coefficient of the takeoff configuration Target takeoff distance It is the top-level indicator given for ultra-short takeoff missions, namely the maximum ground takeoff distance that an aircraft is allowed to accelerate from a standstill to takeoff.

[0036] It should be noted that, generally speaking, tail capacity design is carried out after the design parameters of other parts of the fuselage are initially determined. Therefore, in this embodiment, the takeoff weight m and wing reference area are not specified. and the maximum lift coefficient of the takeoff configuration In execution such Figure 2 The method shown has already been determined accordingly.

[0037] As mentioned earlier, under low-speed, high-thrust conditions, the dynamic pressure difference between the slipstream and non-slipstream regions is most pronounced, and the trim requirements for the horizontal stabilizer are the most demanding. The purpose of step S200 is to lock in the design input parameters for this extreme condition, providing a benchmark for subsequent torque calculations and tail capacity design that does not rely on empirical point selection.

[0038] Specifically, it can be based on the takeoff weight (m) and the wing reference area. Maximum lift coefficient of takeoff configuration and the preset safety margin multiple Determine takeoff speed :

[0039]

[0040] in, It is the acceleration due to gravity. For atmospheric density, the safety margin k is usually taken as 1.1 to 1.2 to ensure that the takeoff speed is above the stall boundary and has a certain control margin. Other parameters can be referred to the previous introduction.

[0041] Furthermore, in determining the takeoff speed Subsequently, based on takeoff and landing dynamics, the ground taxiing phase approximates uniform acceleration, and the average acceleration can be obtained. :

[0042]

[0043] Takeoff thrust needs to provide the inertial force required for acceleration and balance aerodynamic drag and ground rolling friction, thus yielding the thrust requirement for takeoff. :

[0044]

[0045] in, For average aerodynamic drag, For average aerodynamic lift, denoted as the runway friction coefficient, all of which are known parameters.

[0046] It should be noted that in some implementations, due to the high thrust-to-weight ratio and extremely low takeoff speed and short acceleration time of ultra-short takeoff and landing (UST) aircraft, aerodynamic forces and friction can be considered secondary factors compared to thrust. That is, under low-precision design conditions, the thrust requirement... It can be determined using the following formula:

[0047]

[0048] Step S202: Determine the induced velocity of the aircraft rotor disk position based on the thrust requirement and the takeoff speed.

[0049] Based on the thrust requirement obtained in step S200 and takeoff speed By combining momentum theory, the induced velocity at the propeller disk of the aircraft can be determined. Induction speed The velocity increment of the airflow at the propeller disk relative to the distant incoming flow is also a major factor affecting the dynamic pressure in the slipstream region.

[0050] Specifically,

[0051]

[0052] in, This represents the propeller disk area; other parameters can be found in the previous text.

[0053] Step S204: Determine the first dynamic pressure of the non-slipstream region of the wing based on the takeoff speed, and determine the second dynamic pressure of the slipstream region of the wing based on the induced velocity.

[0054] The dynamic pressure in the non-slip flow region is referred to as the first dynamic pressure. Determined solely by takeoff speed and free flow pressure:

[0055] Determining the dynamic pressure (referred to as the second dynamic pressure) in the slipstream region requires considering the contribution of the induced velocity and the degree of axial development of the wake. (Refer to...) Figure 3 There is an axial distance Z between the propeller disk and the leading edge of the wing. The larger this axial distance, the more fully the propeller disk wake develops upon reaching the wing, and the greater the slipstream velocity. Conversely, the larger the propeller disk radius R, the greater the radial spread of the wake, and the lower the relative development degree of the wake at the same axial distance. The actual airflow velocity in the slipstream region... From takeoff speed Superimposed with wake development coefficient Weighted induced velocity composition:

[0056]

[0057] Correspondingly, the second dynamic pressure This is represented as:

[0058]

[0059] In this embodiment, considering the slipstream evolution properties in momentum theory, the degree of wake development varies at different axial positions downstream of the rotor disk. Immediately behind the rotor disk, the slipstream velocity has not yet fully developed, and the axial component of the induced velocity has not yet reached its far-field extreme. As the axial distance increases, surrounding air is continuously entrained into the wake, and the slipstream velocity gradually increases, approaching the far-field theoretical limit—approximately twice the induced velocity. This means that the axial installation distance between the wing and the rotor disk will affect the magnitude of the slipstream velocity felt at the wing, and thus affect the actual dynamic pressure in the slipstream region. In this embodiment, the wake development coefficient is... The value range is set to 1 to 2, and the larger the axial distance, the closer the wake development coefficient is to 2; since the propeller disk radius affects the wake diffusion range, the larger the propeller disk radius, the closer the wake development coefficient is to 1. In some implementations, the wake development coefficient can be set as:

[0060]

[0061] Referring to the above formula, when Z=0 (the wing is close to the rear of the rotor disk), =1 indicates that the wake has not yet begun to develop, and the axial component of the induced velocity at the wing is the same as the induced velocity at the rotor disk. As Z increases, As the induced velocity gradually increases and approaches 2, the axial component of the induced velocity at the wing approaches the far-field theoretical extreme value 2. For the same axial distance Z, the larger the propeller disk radius R, the better. The smaller the value, The closer it is to 1, the lower the relative development of the wake of a large-diameter propeller disk at the same axial distance. The value of is strictly limited to [1,2), and is positively correlated with the axial distance and negatively correlated with the propeller disk radius.

[0062] Step S206: Determine the wing aerodynamic torque based on the first dynamic pressure and the second dynamic pressure.

[0063] After determining the values ​​of the first dynamic pressure (non-slipstream region) and the second dynamic pressure (slipstream region), the aerodynamic torque of the wing can be determined by integrating and summing the aerodynamic torques generated by the dynamic pressures of these two regions along the wing span.

[0064] Reference Figure 4 It can be seen that each section of the wing spanwise is situated within a different flow field. The section directly behind the propeller is in the slipstream region, where the local dynamic pressure is taken as the second dynamic pressure value; the sections between adjacent propellers and the sections outside the propeller spanwise coverage area are in the non-slipstream region, where the local dynamic pressure is taken as the first dynamic pressure value. The boundary between these two regions is determined by the spanwise mounting position of each propeller and the equivalent radius of the wake at the wing. To be determined jointly. It can be determined by the mass flow rate conservation relationship—due to the slipstream velocity. The speed is greater than the speed at the propeller disk. The mass flow rate conservation requirement states that the mass flow rate through the propeller disk must equal the mass flow rate through the wake section at the wing. Therefore, the radius of the wake section at the wing is... Smaller than the propeller disk radius R, that is:

[0065]

[0066] Therefore, a location can be identified as a slipstream region if the following conditions are met:

[0067]

[0068] in, Represents the spanwise coordinates of any position on the wing. This represents the spanwise coordinates of the propeller center closest to this location. It should be noted that the slipstream regions of multiple propellers may be adjacent, overlapping, or separate in the spanwise direction, depending on the propeller spacing and... Based on the relative size, the slipstream region shared by multiple propellers is still identified as a simple slipstream region.

[0069] This allows us to determine the dynamic pressure (first or second dynamic pressure) at various locations on the wing. Integrating the local dynamic pressure at each location along the spanwise direction yields the wing's aerodynamic torque. :

[0070]

[0071] Where b is the wingspan, within the slipstream region Non-slip flow area , For the local string length, The local zero-lift pitch moment coefficient, The local lift coefficient, The location of the body's center of mass. This is the location of the local pressure center.

[0072] Step S208: Determine the horizontal stabilizer trim moment corresponding to the takeoff limit trim state based on the wing aerodynamic moment.

[0073] wing aerodynamic moment The main nose-down moment in the takeoff limit trim state can be determined by the wing aerodynamic moment, and the nose-up moment that the aircraft tail needs to provide is the horizontal tail trim moment.

[0074] In some implementations, for low-precision design conditions, the wing aerodynamic torque can be... As the pitching torque of the entire aircraft Then, the trim moment of the horizontal stabilizer is determined accordingly. .

[0075] However, in distributed electric propulsion configurations, the distributed propellers are typically mounted forward of the wings, and the thrust line does not cross the aircraft's center of gravity. This layout feature introduces a torque source that is less of a concern in traditional tail-capacity designs—the thrust eccentricity torque. Consequently, in some other implementations, the nose-down torque... Excluding wing aerodynamic moment It also includes the additional torque generated by the distributed propeller thrust offset (i.e., thrust eccentricity torque). :

[0076]

[0077] Where N is the total number of propellers on the wing. For the first propeller thrust vector Indicates the first The lever arm vector from the point of application of propeller thrust to the aircraft's center of gravity. During the conceptual design phase... Thrust requirement can be taken as the standard. The distribution values ​​among the propellers. If all propellers have the same specifications and rotational speed, . The lever arm vector is determined by the geometric relationship between the propeller installation position and the center of gravity position. In the conceptual design stage, the propeller installation position is also a known quantity. The spanwise, chordwise, and vertical coordinates of the propeller can be directly read, and then the difference between them and the center of gravity coordinates is used to obtain the lever arm vector.

[0078] It should be noted that under different design parameters (the main influencing factor being the relative position of the propeller thrust line and the aircraft center), the direction of this additional torque is different; that is, the additional torque may be a pitching torque or a nose-down torque. This embodiment reflects this characteristic through the direction of the vector (the additional torque). (represented as a vector).

[0079] Therefore, the pitching torque of the entire aircraft for:

[0080]

[0081] To balance To determine the target, the required pitching moment that the horizontal stabilizer needs to provide is determined; this pitching moment is the horizontal stabilizer trimming moment. ,satisfy .

[0082] Step S210: Determine the tail capacity design parameters of the aircraft based on the horizontal tail trim moment.

[0083] After determining the trim moment of the horizontal stabilizer, this step converts it into tail capacity design parameters. The horizontal stabilizer is in free flow, and its maximum usable pitching moment is determined by the free flow pressure (first dynamic pressure). Flat tail area Tail arm length Horizontal tail lift coefficient at elevator deflection limit Horizontal tail dynamic pressure efficiency and lift safety margin Common constraints:

[0084]

[0085] in, Reflecting the obstruction effect of the fuselage wake on the horizontal stabilizer airflow, a value of 0.85 to 0.95 is typically used for T-tail configurations. Lift safety margin. This is used to ensure that there is a margin between the actual operating point of the horizontal stabilizer and the stall boundary.

[0086] Expanding the trim conditions allows us to determine the minimum tail capacity that satisfies the manipulation boundary constraints. :

[0087]

[0088] Determine the minimum tail capacity Then, the area of ​​the flat tail was determined. With tail arm length The product of the two. However, the flat tail area... With tail arm length What values ​​should each take to determine the minimum tail capacity? Not the only decision—different , Both can meet the same tail capacity requirement. Therefore, the specific horizontal tail area can be determined according to design requirements. With tail arm length The value of is used as the tail capacity design parameter of the aircraft.

[0089] like Figure 2 The method for tail capacity design of a distributed propulsion aircraft shown does not require the use of a high-fidelity CFD simulation model to solve the flow field. Only the overall layout parameters are needed. The wing aerodynamic torque can be quickly obtained by analytical calculation during the conceptual design stage, and the tail capacity design parameters of the aircraft can be determined accordingly.

[0090] In some implementations, in step S204, if the wing model is simplified to a rectangular wing, the spanwise length of the slipstream can be used as a reference. spanwise length of non-slip flow region Wing length First dynamic pressure and the second dynamic pressure The mean dynamic pressure of the wing is determined; then, based on the mean dynamic pressure of the wing, the wing area, the mean aerodynamic chord length, and the pitching moment coefficient of the wing relative to the center of gravity, the aerodynamic moment of the wing is determined.

[0091] Specifically, the aerodynamic moment of a rectangular wing The following formula can be used as a reference to determine:

[0092]

[0093] in, The pitching moment coefficient of the wing relative to the center of gravity can be determined by the wing's airfoil data, flap deflection, and angle of attack. It is a known quantity that can be obtained in the conceptual design stage. Other parameters can be found in the previous text.

[0094] Therefore, for cases where the wing plan shape is rectangular or approximately rectangular (i.e., the chord length remains basically unchanged along the spanwise direction), the calculation of the wing aerodynamic moment can be greatly simplified, eliminating the need for spanwise integration at each station.

[0095] Furthermore, when the propeller slipstream completely covers the entire span of the wing, and the wing model is a rectangular wing, the wing aerodynamic moment... The formula for determining can be further simplified to:

[0096]

[0097] In some implementations, such as Figure 2 In step S210, the horizontal tail area and tail arm length can be determined with the goal of minimizing the equivalent weight cost. The equivalent weight cost is the sum of the horizontal tail wetted area increment and the tail strut wetted area increment.

[0098] As mentioned above, a certain It can be made of different and This can be achieved through combination. However, the cost in terms of structural weight is not equivalent to that of "large area with short tail arm" and "small area with long tail arm".

[0099] Generally, a major optimization goal for aircraft is to keep the airframe weight within a low range. Therefore, in this embodiment, the structural weight differences between different combinations are quantified by equivalent weight cost. The equivalent weight cost is defined as the sum of two infiltration area increments: the horizontal stabilizer area... Increase in infiltration area and tail boom length Increase in area .

[0100] The following provides a method for determining the wetted area increment of a T-tail fin. It should be understood that the method for determining the equivalent weight cost in this embodiment is not limited to a fixed airframe configuration; it is only necessary to determine the wetted area increment under the corresponding configuration.

[0101] For the T-shaped tail fin, the horizontal tail area Increase in infiltration area Possible forms:

[0102]

[0103] Tail arm length Increase in area Possible forms:

[0104]

[0105] For the T-tail, this embodiment simplifies the fuselage tail section to an equivalent diameter of [missing information] when determining the area increment. cylinder, This is the baseline value for the tail boom length.

[0106] Therefore, the flat tail area can be obtained. and tail boom length The goal is to minimize the overall increase in infiltration area.

[0107]

[0108] The values ​​for the horizontal stabilizer area and tail arm length are determined and used as design parameters for the aircraft's tail capacity. It should be noted that during the determination of these values, care must be taken to ensure that the values ​​for the horizontal stabilizer area and tail arm length do not exceed a preset parameter space (this preset parameter space is also known during the conceptual design phase).

[0109] Figure 5 A schematic diagram of a tail capacity design device for a distributed propulsion aircraft according to one embodiment of this specification is shown. See also Figure 5 The device can be deployed in a computing platform or server and includes the following modules.

[0110] The takeoff condition determination module 500 is used to determine the takeoff speed and thrust requirements corresponding to the takeoff limit trim state based on the overall design parameters of the aircraft and the target takeoff run distance.

[0111] The induced velocity determination module 502 is used to determine the induced velocity of the aircraft rotor disk position based on thrust requirements and takeoff speed.

[0112] The dynamic pressure zone determination module 504 is used to determine the first dynamic pressure of the non-slipstream region of the wing based on the takeoff speed, and to determine the second dynamic pressure of the slipstream region of the wing based on the induced velocity.

[0113] The wing moment determination module 506 is used to determine the wing aerodynamic moment based on the first dynamic pressure and the second dynamic pressure.

[0114] The trim moment determination module 508 is used to determine the horizontal stabilizer trim moment corresponding to the takeoff limit trim state based on the wing aerodynamic moment.

[0115] The design parameter determination module 510 is used to determine the tail capacity design parameters of the aircraft based on the horizontal tail trim moment.

[0116] The functions of each module correspond to the processing performed in steps S200 to S210 of the aforementioned method embodiments in this specification. The modules are connected via data streams, with the output of the preceding module constituting the input of the following module. Each module can be implemented in software—for example, as computer program code stored in memory and executed by a processor—or synthesized using a hardware description language (such as VHDL or Verilog) into hardware logic circuits in a programmable logic device or application-specific integrated circuit.

[0117] In some implementations, the takeoff condition determination module, induced velocity determination module, dynamic pressure zoning determination module, wing moment determination module, and design parameter determination module can be integrated and run on the same processor, or they can be deployed on multiple processors working collaboratively. The logical division between the modules here is only illustrative—in other implementations, the functions of some modules can be merged (e.g., the induced velocity determination module and the dynamic pressure zoning determination module can be merged into a flow field parameter determination module), or they can be further split (e.g., the tail volume design parameter determination module can be split into a minimum tail volume determination submodule and a parameter optimization submodule).

[0118] In the 1990s, improvements to a technology could be clearly distinguished as either hardware improvements (e.g., improvements to the circuit structure of diodes, transistors, switches, etc.) or software improvements (improvements to the methodology). However, with technological advancements, many methodological improvements today can be considered direct improvements to the hardware circuit structure. Designers almost always obtain the corresponding hardware circuit structure by programming the improved methodology into the hardware circuit. Therefore, it cannot be said that a methodological improvement cannot be implemented using hardware physical modules. For example, a Programmable Logic Device (PLD) (such as a Field Programmable Gate Array (FPGA)) is such an integrated circuit whose logic function is determined by the user programming the device. Designers can program and "integrate" a digital system onto a PLD themselves, without needing chip manufacturers to design and manufacture dedicated integrated circuit chips. Furthermore, nowadays, instead of manually manufacturing integrated circuit chips, this programming is mostly implemented using "logic compiler" software. Similar to the software compiler used in program development, the original code before compilation must also be written in a specific programming language, called a Hardware Description Language (HDL). There are many HDLs, such as ABEL (Advanced Boolean Expression Language), AHDL (Altera Hardware Description Language), Confluence, CUPL (Cornell University Programming Language), HDCal, JHDL (Java Hardware Description Language), Lava, Lola, MyHDL, PALASM, and RHDL (Ruby Hardware Description Language). Currently, the most commonly used are VHDL (Very-High-Speed ​​Integrated Circuit Hardware Description Language) and Verilog. Those skilled in the art should also understand that by simply performing some logic programming on the method flow using one of these hardware description languages ​​and programming it into an integrated circuit, the hardware circuit implementing the logical method flow can be easily obtained.

[0119] The controller can be implemented in any suitable manner. For example, it can take the form of a microprocessor or processor and a computer-readable medium storing computer-readable program code (e.g., software or firmware) executable by the (micro)processor, logic gates, switches, application-specific integrated circuits (ASICs), programmable logic controllers, and embedded microcontrollers. Examples of controllers include, but are not limited to, the following microcontrollers: ARC 625D, Atmel AT91SAM, Microchip PIC18F26K20, and Silicon Labs C8051F320. A memory controller can also be implemented as part of the control logic of the memory. Those skilled in the art will also recognize that, in addition to implementing the controller in purely computer-readable program code form, the same functionality can be achieved by logically programming the method steps to make the controller take the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers. Therefore, such a controller can be considered a hardware component, and the means included therein for implementing various functions can also be considered as structures within the hardware component. Alternatively, the means for implementing various functions can be considered as both software modules implementing the method and structures within the hardware component.

[0120] The systems, devices, modules, or units described in the above embodiments can be implemented by computer chips or physical entities, or by products with certain functions. A typical implementation device is a server system. Of course, this application does not exclude the possibility that, with the future development of computer technology, the computer implementing the functions of the above embodiments can be, for example, a personal computer, a laptop computer, an in-vehicle human-machine interaction device, a cellular phone, a camera phone, a smartphone, a personal digital assistant, a media player, a navigation device, an email device, a game console, a tablet computer, a wearable device, or any combination of these devices.

[0121] While one or more embodiments of this specification provide the operational steps of the methods described in the embodiments or flowcharts, more or fewer operational steps may be included based on conventional or non-inventive means. The order of steps listed in the embodiments is merely one possible order of execution among many steps and does not represent the only possible order. In actual device or end product execution, the methods shown in the embodiments or drawings may be executed sequentially or in parallel (e.g., in a parallel processor or multi-threaded processing environment, or even a distributed data processing environment). The terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, product, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, product, or apparatus. Without further limitations, the presence of other identical or equivalent elements in the process, method, product, or apparatus that includes said elements is not excluded. For example, the use of terms such as "first," "second," etc., is to denote names and does not indicate any particular order.

[0122] For ease of description, the above devices are described in terms of function, divided into various modules. Of course, when implementing one or more of these specifications, the functions of each module can be implemented in one or more software and / or hardware components, or a module that performs the same function can be implemented by a combination of multiple sub-modules or sub-units. The device embodiments described above are merely illustrative. For example, the division of units is only a logical functional division; in actual implementation, there may be other division methods. For example, multiple units or components may be combined or integrated into another system, or some features may be ignored or not executed. Furthermore, the coupling or direct coupling or communication connection shown or discussed may be through some interfaces, indirect coupling or communication connection between devices or units, and may be electrical, mechanical, or other forms.

[0123] This invention is described with reference to flowchart illustrations and / or block diagrams of methods, apparatus (systems), and computer program products according to embodiments of the invention. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, special-purpose computer, embedded processor, or other programmable data processing apparatus to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing apparatus, generate instructions for implementing the flowchart illustrations and / or block diagrams. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0124] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing device to function in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The function specified in one or more boxes.

[0125] These computer program instructions may also be loaded onto a computer or other programmable data processing equipment to cause a series of operational steps to be performed on the computer or other programmable equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0126] In a typical configuration, a computing device includes one or more processors (CPU), input / output interfaces, network interfaces, and memory.

[0127] Memory may include non-persistent storage in computer-readable media, such as random access memory (RAM) and / or non-volatile memory, such as read-only memory (ROM) or flash RAM. Memory is an example of computer-readable media.

[0128] Computer-readable media includes both permanent and non-permanent, removable and non-removable media that can store information by any method or technology. Information can be computer-readable instructions, data structures, modules of programs, or other data. Examples of computer storage media include, but are not limited to, phase-change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other memory technologies, CD-ROM, digital versatile optical disc (DVD) or other optical storage, magnetic tape, magnetic disk storage, graphene storage or other magnetic storage devices, or any other non-transferable medium that can be used to store information accessible by a computing device. As defined herein, computer-readable media does not include transient computer-readable media, such as modulated data signals and carrier waves.

[0129] Those skilled in the art will understand that one or more embodiments of this specification can be provided as a method, system, or computer program product. Therefore, one or more embodiments of this specification may take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, one or more embodiments of this specification may take the form of a computer program product implemented on one or more computer-usable storage media (including, but not limited to, disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0130] One or more embodiments of this specification can be described in the general context of computer-executable instructions, such as program modules, that are executed by a computer. Generally, program modules include routines, programs, objects, components, data structures, etc., that perform a particular task or implement a particular abstract data type. One or more embodiments of this specification can also be practiced in distributed computing environments where tasks are performed by remote processing devices connected via a communication network. In distributed computing environments, program modules can reside in local and remote computer storage media, including storage devices.

[0131] The various embodiments in this specification are described in a progressive manner. Similar or identical parts between embodiments can be referred to mutually. Each embodiment focuses on describing the differences from other embodiments. In particular, system embodiments are basically similar to method embodiments, so the description is relatively simple; relevant parts can be referred to the descriptions in the method embodiments. In the description of this specification, the terms "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., refer to specific features, structures, materials, or characteristics described in connection with that embodiment or example, which are included in at least one embodiment or example of this specification. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described can be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification and the features of different embodiments or examples.

[0132] The above description is merely an embodiment of one or more embodiments of this specification and is not intended to limit the scope of these embodiments. Various modifications and variations can be made to these embodiments by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of this specification should be included within the scope of the claims.

Claims

1. A tail capacity design method for a distributed propulsion aircraft, comprising: Based on the overall design parameters of the aircraft and the target takeoff run distance, determine the takeoff speed and thrust requirements corresponding to the takeoff limit trim state; Based on the thrust requirement and the takeoff speed, determine the induced velocity for the aircraft rotor disk position; The first dynamic pressure in the non-slipstream region of the wing is determined based on the takeoff speed, and the second dynamic pressure in the slipstream region of the wing is determined based on the induced velocity. The wing aerodynamic torque is determined based on the first dynamic pressure and the second dynamic pressure; Based on the wing aerodynamic moment, determine the horizontal stabilizer trim moment corresponding to the takeoff limit trim state; The tail capacity design parameters of the aircraft are determined based on the horizontal stabilizer trim moment. The determination of the second dynamic pressure in the wing slipstream region based on the induced velocity includes: Based on the axial distance between the propeller disk and the wing and the radius of the propeller disk, a wake development coefficient characterizing the degree of axial development of the wake is determined. The wake development coefficient ranges from 1 to 2. The larger the axial distance, the larger the wake development coefficient. The larger the radius of the propeller disk, the smaller the wake development coefficient. The takeoff speed is added to the induced speed weighted by the wake development coefficient to determine the wing slipstream speed; The second dynamic pressure in the wing slipstream region is determined based on the wing slipstream velocity.

2. The method as described in claim 1, characterized in that, The overall design parameters of the aircraft include takeoff weight, wing reference area, and maximum lift coefficient of takeoff configuration.

3. The method as described in claim 2, characterized in that, Determine the takeoff speed and thrust requirements corresponding to the takeoff limit trim state, including: The takeoff speed is determined based on the takeoff weight, the wing reference area, the maximum lift coefficient of the takeoff configuration, and the preset safety margin multiple. The thrust requirement is determined based on the takeoff speed and the target takeoff run distance.

4. The method as described in claim 1, characterized in that, The aerodynamic torque of the wing is determined based on the first dynamic pressure and the second dynamic pressure, specifically including: The aerodynamic torque of the wing is determined by integrating the first dynamic pressure in the slipstream region and the second dynamic pressure in the non-slipstream region along the wing span. The slipstream region and the non-slipstream region are determined based on the equivalent radius of the propeller wake and the propeller position.

5. The method as described in claim 1, characterized in that, For a rectangular wing, the aerodynamic moment of the wing is determined based on the first dynamic pressure and the second dynamic pressure, specifically including: The average dynamic pressure of the wing is determined based on the spanwise length of the slipstream region, the spanwise length of the non-slipstream region, the wing length, the first dynamic pressure, and the second dynamic pressure. The aerodynamic moment of the wing is determined based on the mean dynamic pressure of the wing, the wing area, the mean aerodynamic chord length, and the pitching moment coefficient of the wing relative to the center of gravity.

6. The method as described in claim 1, characterized in that, Based on the wing aerodynamic moment, the horizontal stabilizer trim moment corresponding to the takeoff limit trim state is determined, specifically including: The thrust offset pitching moment is obtained by taking the cross product of the thrust vector and the vector from the point of thrust application to the center of gravity for each propeller and summing them. The wing aerodynamic moment and the thrust offset pitching moment are combined into the overall nose-down moment; The horizontal stabilizer trimming torque is determined with the overall nose-down torque as the target.

7. The method as described in claim 1, characterized in that, The tail capacity design parameters of the aircraft are determined based on the horizontal stabilizer trim moment, specifically including: Based on the preset horizontal stabilizer dynamic pressure hysteresis efficiency, horizontal stabilizer lift coefficient under elevator deflection limit, lift safety margin, and the first dynamic pressure, determine the minimum tail capacity to satisfy the horizontal stabilizer trim torque. Based on the minimum tail capacity, the values ​​of the horizontal stabilizer area and tail arm length are determined as the tail capacity design parameters of the aircraft.

8. The method as described in claim 7, characterized in that, Based on the minimum tail capacity, the horizontal tail area and tail arm length are determined, specifically including: Based on the minimum tail capacity, with the goal of minimizing the equivalent weight cost, the values ​​of the horizontal tail area and tail arm length are determined, wherein the equivalent weight cost is the sum of the increase in the horizontal tail wetted area and the increase in the tail strut wetted area.

9. A computing device comprising a memory and a processor, wherein the memory stores executable code, and the processor, when executing the executable code, implements the method as claimed in any one of claims 1-8.

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