Method for reducing entire aircraft drag

Abstract

The present invention provides a method for greatly reducing entire aircraft drag, in particular to reducing induced drag, comprising: controlling the orientation of an aerodynamic force vector; adjusting a yaw angle of the aerodynamic force vector relative to the vertical axis of a flight path axis system, and / or increasing the aspect ratio of wings or increasing the effective aspect ratio of the wings; and optimizing the pressure distribution over an airfoil surface. The present invention can reduce the induced drag to the maximum extent, or even eliminates the induced drag, and can also convert harmful induced drag into induced thrust which is beneficial to the motion of an aircraft. In the present invention, on the basis of the drag analysis, especially an induced drag generation mechanism and root, a correct approach for drag reduction has been found. On the basis of the present invention, an induced drag calculation method is improved, and an entire aircraft drag mathematical model is perfected. By means of the method in the present invention, the entire aircraft drag can be efficiently reduced, thereby greatly increasing the lift-to-drag ratio, and effectively improving the flight efficiency.

Description

A method to reduce overall aircraft drag Technical Field

[0001] This invention relates to the field of aerospace engineering technology, and in particular, to the field of methods for reducing induced drag of aircraft; specifically, to a method for reducing overall aircraft drag. Background Technology

[0002] Accurate calculation of induced drag for the entire aircraft has long been a challenging problem in aircraft design. When an aircraft flies in the atmosphere, drag and thrust are equally important. Flight performance is primarily determined by residual thrust, which influences performance parameters such as speed, range, maneuverability, fuel consumption, and environmental emissions. Range is directly proportional to lift-to-drag ratio; fuel consumption is inversely proportional to lift-to-drag ratio. Improving the lift-to-drag ratio is a primary objective of aircraft aerodynamic design.

[0003] Although some methods have been developed to reduce zero-lift drag, such as shape optimization and surface treatment techniques, controlling induced drag remains a challenging problem that has not yet been well resolved.

[0004] Aerodynamic drag can be divided into two categories: zero-lift drag and induced drag. Zero-lift drag is an inherent property of objects moving in a fluid; it is absolute drag and can only be reduced, not completely eliminated. Zero-lift drag is always opposite to the direction of motion. The second type of drag is called induced drag, which is related to lift and is sometimes referred to as lift-induced drag. Depending on the aircraft's flight conditions, the percentages of zero-lift drag and induced drag in the total drag vary greatly. During takeoff and low-speed flight, induced drag is dominant, accounting for more than 90% of the total drag. During high-altitude cruise and level flight, due to the high flight speed, only a smaller lift coefficient is needed. The proportion of induced drag decreases, but the proportion of zero-lift drag increases. Even at high altitudes, induced drag remains significant.

[0005] When an aircraft is in motion, in order to maintain steady motion, it must simultaneously satisfy the balance of forces in the horizontal and vertical directions, namely: the balance of horizontal thrust and drag and the balance of vertical lift and weight. Among these, the balance of vertical lift and weight is the most important.

[0006] Currently, our understanding of lift and drag still needs to be refined and improved. Clarifying lift, drag, and their interrelationships is crucial. Aerodynamics, based on Bernoulli's principle, solved the balance between lift and weight, and was a catalyst for the invention of the airplane. Lift is not a Newtonian force, but rather a highly efficient force derived from and induced by Newtonian forces, originating from but superior to Newtonian forces. The direction of lift is parallel to gravity and points in the opposite direction (upward, resisting gravity). When an aircraft is in level flight, aerodynamic drag points horizontally, perpendicular to the direction of lift, and the flight velocity vector and lift vector are orthogonal and perpendicular to each other. Theoretically, two orthogonal vectors are completely independent and do not affect each other; that is, there is no direct relationship between lift and drag. However, lift and drag are not completely independent but have an indirect relationship. This is because both lift and drag originate from the residual pressure difference (actually a pressure difference) generated when air flows over the upper and lower surfaces of the airfoil, which, when integrated along the spanwise direction, forms lift. In an ideal fluid environment, with no viscosity and no friction, the airflow velocity remains constant. Based on the law of conservation of energy, generating lift is effortless, and there is no horizontal resistance. However, in reality, friction and viscosity impede airflow, causing deceleration. To maintain a constant velocity, energy must be continuously supplied to the flow field. An engine generates forward thrust to achieve system equilibrium.

[0007] The expression for the overall aircraft drag mathematical model currently used in classic textbooks, design manuals, and by overall design engineers is: C D =C D0 +C Di

[0008] Among them, C D0 It is zero-lift resistance; C Di It is induced resistance;

[0009] C Di The calculation formula is usually expressed as a function of the lift coefficient C. L The relationship is parabolic, i.e., the induced drag and lift coefficient C L It is proportional to the square of C Di =K2*C L 2 K2 = k / πA; K2 is called the induction factor. A is the effective wingspan of the three-dimensional wing. π is pi; k is the wing planar shape correction factor; k = 1 + δ; δ = 0 for an elliptical wing, and δ for a non-elliptical wing is between 0 and 1.

[0010] Zero-lift resistance C D0 Zero-lift drag refers to the portion of drag experienced during flight that is unrelated to lift and depends only on the flight velocity vector. D0It is an inherent characteristic of an aircraft moving in a flow field, and it always exists; zero lift drag C D0 The direction of the zero-lift drag C is always collinear with but opposite to the flight velocity vector. D0 The causes of zero-lift drag include friction caused by the roughness of the flying object's surface, heat exchange in the boundary layer, the adverse pressure gradient of the boundary layer, and shape drag. Zero-lift drag can be reduced, but cannot be eliminated. For the dimensionless zero-lift drag coefficient C... D0 It is essentially a constant, independent of flight speed. To date, considerable research has been conducted on reducing the zero-lift drag coefficient, but significantly reducing C... D0 It is still very difficult.

[0011] Analyzing the above mathematical model of total aircraft drag (the expression for total aircraft drag) and the formula for calculating induced drag, it is found that total aircraft drag is simply a quadratic parabola function with zero lift drag as the origin. Only when combined with the lift coefficient C... L Quadratic terms that are directly proportional to the square (K2*C) L 2 ), without lift coefficient C L A linear term that is directly proportional (denoted as K1*C) L Therefore, we can conclude that: First, in the case of a two-dimensional airfoil, the overall drag expression only contains the zero-lift drag term C. D0 The second term only appears in the three-dimensional case with wingtip vortices. This indicates that two-dimensional airfoils have no induced drag. Secondly, in the case of a three-dimensional airfoil, in addition to zero lift drag, there is an induced drag term generated by the three-dimensional effect of wingtip vortices. This induced drag term is proportional to the square of the lift coefficient and is called lift-induced drag.

[0012] Both lift and drag originate from the aerodynamic vector R. Lift and drag are the projection components of the aerodynamic vector onto two mutually perpendicular axes in the flight path coordinate system. Lift is the cosine component of the resultant aerodynamic force vector R; drag is the sine component of R. The magnitudes of lift and drag, as well as the value of R, are closely related to the spatial orientation of R.

[0013] All forces acting on the aircraft are expressed in the aircraft's motion coordinate system: that is, the velocity coordinate system or track coordinate system (OX) associated with the flight velocity vector. h Y h Z hSince the study primarily focuses on flight performance parameters, it is assumed that the aircraft has no lateral motion; only the motion of the aircraft in the longitudinal vertical plane (two-dimensional) is studied. To calculate the aircraft's motion performance, the equations of motion are expressed in the track coordinate system. Therefore, all external forces are transformed to the track coordinate system through coordinate transformation. Matrix transformation operations are performed using a conventional linear algebra transfer matrix. The elements in the transfer matrix include the direction cosine or direction sine of the angle between the three axes of the corresponding two coordinate systems. In the case of dealing with two-dimensional plane problems, the angle between two coordinate systems is only one angle. For example, the body coordinate system OX... B Y B and track coordinate system (OX) h Y h Z h The angle between the two points is the angle of attack α. Assuming the flight speed V is on the Earth's horizontal plane, the X-axis of the trajectory coordinate system... h The axis and velocity direction are aligned, and the wing is mounted on the fuselage with no installation angle. Therefore, the semi-body coordinate system, fixed to the airfoil profile, is parallel to the body coordinate system, which is fixed to the aircraft fuselage, and the angle between the semi-body coordinate system and the track system is also α. After establishing the above relationship, the wind tunnel experimental data C expressed in the semi-body coordinate system will be... DB (R X C LB (R Y The aerodynamic forces (T) of the engine mounted on the fuselage and the gravity (W) in the ground coordinate system are expressed in the flight path coordinate system through coordinate transformation. The aerodynamic transformation from the fuselage axis system to the flight path axis system is as follows:

[0014] X h The component on the axis (resistance): D = Rsinθcosα + Rcosθsinα;

[0015] Y h The component on the axis (lift): L = Rcosθcosα - Rsinθsinα;

[0016] The trigonometric identities are: sin(α+b)=sin(α)cos(b)+cos(α)sin(b); cos(α+b)=cos(α)cos(b)-sin(α)cos(b); D=R*sin(α+θ)L=R*cos(α+θ);

[0017] The drag force R in the drag expression is expressed in terms of lift L: R = L / cos(α + θ);

[0018] According to linear theory, for small angles, there is an approximate relationship: D=L*Tan(α+θ)=L*sin(α+θ)=L*(α+θ) (radians);

[0019] Expressed in dimensionless form: C Di1 =K1*C L K1 = Tan(α + θ);

[0020] C Di1 It is related to the lift coefficient C L The drag term is proportional to the first power. The drag factor K1 is the tangent of the angle (α+θ). It depends only on the angle by which the aerodynamic vector R deviates from the vertical axis of the trajectory coordinate system, and is independent of the aspect ratio. C Di1 The first part of the drag related to lift is called the first induced drag.

[0021] There is also drag C, which is proportional to the square of the lift coefficient. Di2 This is called the second induced drag. Any force related to the angle of attack projected onto the velocity vector (the sinusoidal component of the direction) can be defined as induced drag. Data obtained from wind tunnel tests of two-dimensional airfoils does not include the additional induced drag term generated by the downwash from wingtip vortices when applied to three-dimensional wings. When designing aircraft using two-dimensional airfoils to three-dimensional wings, two effects must be considered: first, the three-dimensional effect. This is usually expressed by an efficiency coefficient R related to the wing aspect ratio A, i.e.: C L =C L α (Two-dimensional)*R*α; where R=A / (A+2) is the three-dimensional efficiency coefficient. Second, the additional induced drag caused by the wingtip vortex. The formation mechanism and calculation method of this drag are the same as the first induced drag. The calculation method for the second induced drag is discussed below. The downwash velocity W generated by the wingtip vortex, together with the velocity of the incoming flow far ahead, synthesizes a new velocity vector. Therefore, the original aerodynamic vector Y tilts to the right, deviating from a downwash angle ε. To maintain weight balance, the new aerodynamic vector Y... 1 It must be larger than Y, Y 1 = Y / cos(ε). Previously, the projection component of Y in the X direction of the track coordinate system was zero (no induced drag). Now the new aerodynamic vector V... ∞ Induced drag X1 is generated in the X direction of the flight path coordinate system. This is the physical mechanism of the second negative effect, besides reduced lift, produced when a two-dimensional airfoil is applied to a three-dimensional finite-span wing.

[0022] The formula for calculating the second induced drag is derived using physical concepts (see Figures 2 and 3). At a selected flight speed, the corresponding lift coefficient C is calculated to maintain a constant altitude for the aircraft. L And α. The corresponding aerodynamic vector is R1[R1=C L / cos(α)]. If it is a two-dimensional airfoil with an infinite wingspan, there is no wingtip vortex and no downwash. In this case, the downwash angle ε=0, and the induced drag is OA, that is: CDi1 =R1*sin(α);

[0023] Or use C L Expression, C Di1 =C L *tan(α);

[0024] In the case of a three-dimensional airfoil, a downwash occurs with a downwash angle of ε, which is equivalent to a reduction of ε in the effective angle of attack from the original angle of attack α. The remaining (α-ε) angle is smaller than the original angle α. To maintain constant altitude flight, the angle of attack ε must be increased (linear theory). This results in a larger aerodynamic resultant force R2 than R1. Due to the three-dimensional downwash effect, two negative effects occur: more energy is consumed to generate a larger aerodynamic resultant force, and the new aerodynamic resultant force deviates from the Y-axis of the velocity coordinate system by an angle ε. Therefore, the component of the aerodynamic resultant force R2 projected in the velocity vector direction increases. The increased induced drag is AB. AB is the second induced drag C. Di2 To solve for AB, we can use the triangle in the upper right corner of the mechanical analysis diagram above to perform trigonometric function calculations. This is a right triangle. The base angle (α-ε) and the opposite side ΔC... L They can be calculated separately. Therefore, AB = ΔC can be obtained. L *tan(α-ε). Since α-ε is a small angle (not exceeding 15-20 degrees), due to the characteristics of trigonometric functions for small angles, the following approximate formula holds true with an error of no more than 1%. This is sufficiently accurate for engineering calculations:

[0025] tan(α-ε)=sin(α-ε)=(α-ε) Angles are expressed in radians. Taking 15 degrees as an example: tan(15)=0.2679; sin(15)=0.2588; 15 degrees / 57.3=0.2618.

[0026] Therefore, we have C. Di2 =ΔC L *(α-ε). ΔC L This is due to the reduced lift caused by the downwash.

[0027] ΔC L =C L *(1-R), where R is the three-dimensional efficiency coefficient. This is the conversion coefficient when applying the aerodynamic coefficients of a two-dimensional airfoil to a three-dimensional airfoil. From classical lift line theory, we have C L α (ii) = 2π. R = A / (A+2).

[0028] ε=α(1-R); α-ε=α-α(1-R)=α*R; C L =C L α (ii) *R*α=2πRα;

[0029] We obtain α = C L / 2πR; Substitute into the above C Di2 The expression yields the following result: C Di2 =ΔC L *(α-ε)=C L *(1-R)*R*C L / 2πR=[2 / (A+2) / 2πA] * C L 2 ;

[0030] Finally, there is C. Di2 =kR / (πA)*C L 2 k is the introduced error correction coefficient, which can be taken as 1.0-1.05.

[0031] R = A / (A+2) is the three-dimensional effect coefficient, which is a function of the aspect ratio A. Therefore, we can obtain C. Di2 =kR / (πA)*C L 2 =K2*C L 2 K2 = kR*(1 / πA).

[0032] K2 is the second induced drag factor, a function related to the application of a two-dimensional airfoil to a three-dimensional wing and the wing's planar shape. When A is infinite, i.e., a two-dimensional wing, R = 0, there is no downwash, and the induced drag is 0. As A becomes smaller, R gradually decreases; when A approaches 0, R also approaches 0, which is a singularity. The wing itself ceases to exist, making discussions of induced drag meaningless. R in the second induced drag formula is an objectively existing physical quantity, not an artificial correction coefficient. If R is considered a correction coefficient, it has a more reasonable physical meaning than the various correction methods currently used. The currently popular practice of using only the constant k = 1.05 to correct the second induced drag factor overestimates the second induced drag. The error is even greater in the case of small aspect ratio wings. For example, when A = 2.0, 1 / πA = 0.1592. If the popular constant correction (so-called wing planform correction) is used, k = 1.05 and K2 = 0.1671; if the method of this invention is used, R = 0.5 and K2 = 0.0796. The difference between the two is 2.1 times, which means there is an error of 210%.

[0033] Another correction method is the Oswald efficiency factor method, which uses eA to represent the effective aspect ratio. e is a very complex function that depends only on the aspect ratio A. For example, the expression for e in a straight wing is: e = 1.78(1 - 0.045A). 0.68-0.46. Currently, some people suggest taking the range of e as 0.7-0.85. However, there is no reason to derive this range of e, and the error is large.

[0034] In summary, the aviation industry currently ignores or underestimates the first induced drag when calculating the total drag of an aircraft, while overestimating the second induced drag. The current mathematical model of overall aircraft drag has significant flaws and must be corrected. Establishing a correct and complete mathematical model of the entire system depends on whether the mathematical models of each subsystem reflect the real and objective physical world.

[0035] In the overall mathematical model of an aircraft, the mechanical model includes models for gravity, lift, thrust, and drag. The first two are relatively simple, and modeling them presents no major problems. However, the thrust and drag models are much more complex because they involve numerous factors, with drag being a key focus. In particular, research into the physical mechanisms of induced drag is currently insufficient, leading to incomplete and even flawed drag mathematical models, and consequently, to some extent, a deviation from the overall direction of drag reduction.

[0036] Summary of the Invention

[0037] Therefore, the purpose of this invention is to propose a method to significantly reduce the overall flight drag of an aircraft, not only to minimize drag to the maximum extent, but also to eliminate induced drag, transforming harmful induced drag into induced thrust that is beneficial to aircraft motion; by analyzing the generation mechanism and root causes of drag, especially induced drag, to find the correct direction for drag reduction, and to develop methods to reduce, eliminate, or even change and transform induced drag, turning negative energy drag into positive energy thrust, so as to efficiently reduce the overall drag of the aircraft, improve the lift-to-drag ratio, and effectively improve flight efficiency.

[0038] This invention provides a method for reducing overall aircraft drag, comprising the following steps (as shown in Figure 1):

[0039] S1. Design a method to reduce zero-lift drag and reduce or eliminate induced drag; the method to reduce or eliminate induced drag includes: controlling the orientation of the aerodynamic vector; the method to control the orientation of the aerodynamic vector includes: adjusting the skew angle of the aerodynamic vector from the vertical axis of the track axis system, and / or, increasing the aspect ratio of the wing or increasing the effective aspect ratio of the wing.

[0040] The method for adjusting the skew angle of the aerodynamic vector from the vertical axis of the flight path includes: controlling the azimuth angle θ of the airfoil aerodynamic resultant force vector to be within 0-0.2 degrees when the angle of attack α is zero (design state); maximizing the lift coefficient of the airfoil near the angle of attack α where the maximum lift-to-drag ratio occurs; and then deriving the airfoil surface pressure distribution and corresponding airfoil profile geometry to satisfy the optimization index of limiting the azimuth angle θ of the airfoil aerodynamic resultant force vector to within 0-0.2 degrees when the angle of attack α is zero.

[0041] S2. Optimize the pressure distribution on the airfoil surface to calculate the airfoil geometry. The method for calculating the airfoil geometry includes:

[0042] The geometry of the existing airfoil profile is locally modified, the airfoil midsection is extended to form a flat platform; the airfoil leading edge is bent downwards; the stagnation point is lower; the zero lift angle of attack is close to -5 degrees to -10 degrees; the thickness of the airfoil tail section is increased; the lower wing surface is bent upwards, and the skewness of the R-vector of the entire aircraft is controlled to be less than 1 degree in the high-speed cruise phase.

[0043] Furthermore, the design of step S1 to reduce zero-lift drag and reduce or eliminate induced drag includes: constructing a mathematical model of the overall aircraft's flight drag, which is based on the total drag in flight: the sum of zero-lift drag and induced drag. The expression for the total drag in flight is: C D =C D0 +C Di =C D0 +K1*C L +K2*C L 2 =C D0 +Tan(α+θ)*C L +sign(K1)*kR / πA*C L 2 (1)

[0044] In equation (1), C D0 It is zero-lift resistance; C Di It is induced resistance, K2*C L 2 It is C Di With lift coefficient C L A quadratic term proportional to the square of ; K2 = kR / πA; K2 is called the induction factor; K1*C L It is C Di With lift coefficient C LThe linear term is proportional to the airfoil; K1 is the drag factor; A is the effective aspect ratio of the three-dimensional airfoil; π is pi; k is the error correction coefficient; K = A / (A+2) is the three-dimensional effect coefficient of the airfoil; α is the angle of attack; and θ is the angle between the aerodynamic vector of the two-dimensional airfoil and the vertical axis of the airframe coordinate system.

[0045] Furthermore, the design of step S1 to reduce zero-lift drag and reduce or eliminate induced drag also includes: constructing a mathematical model of total induced drag, which is based on the sum of the first induced drag and the second induced drag in flight. The expression for the total induced drag in flight is: C Di =C Di1 +C Di2 =K1*C L +K2*C L 2 =Tan(α+θ)*C L +sign(K1)*(kR / πA)*C L 2 (2)

[0046] In equation (2), C Di1 It is the first induced resistance, C Di2 It is the second induced resistance.

[0047] Furthermore, the airfoil geometry in step S2 includes: the curvature of the upper and lower airfoil surfaces, the camber of the airfoil, the relative thickness, the geometry of the airfoil leading edge and trailing edge, and optimizing the pressure distribution on the airfoil surface.

[0048] This modification method can achieve excellent lift-to-drag ratio characteristics. For example, the LRT airfoil, the Global Hawk airfoil, and the QYX-5 airfoil designed by the applicant have maximum lift-to-drag ratios of over 155, 130, and 200 respectively in two dimensions.

[0049] Furthermore, the method for designing the airfoil geometry in step S2 also includes:

[0050] By changing the relative orientation of the airfoil and fuselage, and altering the relative orientation of the aerodynamic vector with respect to the track coordinate system, induced drag can be reduced or even eliminated, and induced drag can be converted into induced thrust.

[0051] The methods for changing the relative orientation of the airfoil and the fuselage include one or a combination of two of the following: reversing the orientation of the leading and trailing edges of the airfoil, and adjusting the wing installation angle in the negative direction.

[0052] The relative orientation of the airfoil and fuselage can be adjusted offline or online. Without affecting other flight performance characteristics, over-adjustments can be made to achieve the optimal effect of converting induced drag into induced thrust.

[0053] Furthermore, the method for eliminating induced drag includes: employing a dipole wing or a counter-rotating wing;

[0054] The dipole wing consists of two identical wings mounted on the fuselage, with the leading edge of one wing pointing forward and the trailing edge of the other pointing forward. When the aircraft moves forward, the aerodynamic vectors are opposite in direction to the flight speed, and their projection components on the horizontal axis of the track system cancel each other out, resulting in zero induced drag.

[0055] Furthermore, the counter-damping wing is a single wing mounted on the fuselage, forming a quasi-dipole wing pair with the existing wing layout, thereby counter-damping induced drag.

[0056] In addition, in some cases, the leading and trailing edges of the two wings of a single wing are reversed to offset the induced drag of the left and right wings, thereby basically eliminating the induced drag of the entire aircraft.

[0057] Furthermore, the method for designing the airfoil geometry in step S2 also includes:

[0058] Add a leading-edge slotted suction attachment to the existing airfoil to generate leading-edge suction.

[0059] The main innovative technical points and technical effects of the present invention are as follows:

[0060] I. The mathematical model for overall aircraft drag has been improved:

[0061] By examining existing mathematical models of total aircraft drag, deficiencies and flaws were found, leading to significant errors. Through in-depth analysis of the physical nature of drag and understanding its origins, a correct formula for induced drag was derived, supplementing the existing technology with a first induced drag term proportional to the first power of the lift coefficient. Starting from physical concepts, the additional drag (i.e., second induced drag) resulting from the downwash velocity generated by applying a two-dimensional airfoil to a three-dimensional wing, altering the magnitude and direction of airflow, was derived. While the calculation formulas developed in this invention appear similar to traditional induced drag formulas, they are substantially different and more realistic. The total aircraft drag and induced drag calculation formulas derived from the fundamental nature of induced drag, or the mathematical models of induced drag and total aircraft drag, provide a more realistic drag mathematical model for aircraft design or simulation experiments, ensuring high reliability of the final calculation or simulation results.

[0062] When conducting flight simulation tests, there are two scenarios: First, if a discrete model is used, the wind tunnel test results data are directly read during the preliminary design or flight dynamics calculations.

[0063] The Cartesian coordinate system data C, expressed in body coordinate system or half-body coordinate system, needs to be processed in advance. D and C L Perform preprocessing. L Three-dimensional efficiency conversion is required based on the geometric parameters of the three-dimensional wing; C D A second induced drag coefficient needs to be added (the first induced drag coefficient is already included in the wind tunnel test data, so it does not need to be added again). The wind tunnel test results are two-dimensional airfoil data.

[0064] Second, if an analytical model is applied, the main task is to improve the mathematical expression of the model: supplement the first induced drag coefficient which is proportional to the first power of the lift coefficient; and derive the formula for the second induced drag coefficient which is proportional to the square of the lift using physical concepts.

[0065] When calculating induced drag, it is necessary to know the orientation parameters of the aerodynamic vector related to the lift coefficient: the angle θ between the aerodynamic vector of the two-dimensional airfoil and the vertical axis of the airframe coordinate system, the angle of attack α, and the downwash angle ε generated by the wingtip vortex. To calculate these three angles, θ, α, and ε, it is necessary to calculate the pressure distribution and residual pressure coefficient difference on the airfoil surface based on the airfoil profile shape using Bernoulli's principle or the Navier-Stokes equations, and then calculate the polar coordinate angles of the aerodynamic vector of the entire airfoil profile (the two-dimensional airfoil design corresponds to the maximum lift-to-drag ratio; when calculating the maximum lift-to-drag ratio, the zero lift-to-drag coefficient should be subtracted, which is equivalent to shifting the origin of the airframe axis to the right to C). D0 When calculating the maximum lift-to-drag ratio of the entire aircraft, the total drag of the entire aircraft, including zero-lift drag and induced drag coefficients, is used. A well-designed airfoil has a very small angle of attack in its design state, a very large lift-to-drag ratio, and a very small deflection angle of the aerodynamic vector, with (θ+α) approaching 0 degrees. For example, the bottom of the bulge polar curve of a laminar airfoil is a line segment parallel to the y-axis, with θ being zero degrees. The first induced drag in the airfoil design state is zero, and the total airfoil drag is only zero-lift drag.

[0066] However, the angle of attack in non-design conditions is not zero. During takeoff and low-speed flight, a large angle of attack, close to the stall angle of attack, is needed to obtain sufficient lift, at which point both the first and second induced drag are very high. Therefore, the energy consumption (power) during takeoff is often much greater than that during cruise, sometimes exceeding 10 times. If the first drag term is ignored and only the second induced drag term is retained, the total drag is much smaller, resulting in a large error. Therefore, it is essential to supplement the missing terms in the overall aircraft drag mathematical model and correct the calculation formula for the second induced drag term. The improved overall aircraft drag model of this invention more objectively represents the actual situation of the aircraft, which is also one of the key innovative technical points of this invention.

[0067] II. Improved flight performance:

[0068] By significantly reducing drag and increasing lift, the overall lift-to-drag ratio K (K=C) of the aircraft was improved. L / C D The lift-to-drag ratio is directly proportional to the flight range, while the fuel consumption rate is inversely proportional to K. Increasing K can significantly increase the flight range; or, for the same flight range, fuel consumption can be reduced, thus making the system more energy-efficient and environmentally friendly, reducing pollution emissions. Furthermore, because the lift-to-drag ratio increases, the lift force increases, which in turn allows for a larger payload, thereby increasing the effective payload.

[0069] The core technology of this invention is to highlight the physical nature of induced drag, reduce or even eliminate it, and design specific methods to reduce induced drag. To clearly express the physical nature of induced drag and reveal methods for reducing, eliminating, or even converting it into induced thrust, simple and clear mathematical formulas and accompanying simplified diagrams are used for explanation and demonstration:

[0070] The physical essence of induced drag: the projection (direction sine) of the aircraft's aerodynamic vector onto the horizontal axis Xh of the flight path coordinate system. See Figure 2. 诱导阻力 =R*sin(θ+α+ε)=C L *tan(θ+α+ε) (3)

[0071] Since angles θ, α, and ε are all small angles, especially θ and ε, using the characteristics of trigonometric functions of small angles, equation (3) can be simplified to: D 诱导阻力 =C L *tan(θ+α+ε)=C L *tan(θ+α)+C L *tan(ε) = C L *tan(θ+α)+C L *ε=C L *tan(θ+α)+kR*(1 / πA)*C L 2 (4) = K1*C L +K2*C L 2 = First induced resistance + Second induced resistance

[0072] C L *ε=kR*(1 / πA)*C L 2 The derivation is shown in Figure 3.

[0073] The objective of this invention's method for reducing induced drag is to reduce the angle φ between the overall aircraft's aerodynamic vector and the vertical axis, using the flight path coordinate system as a reference system.

[0074] φ = θ + α + ε. Drag reduction means reducing φ, which means reducing all three included angles θ, α, and ε.

[0075] Analyzing the physical meaning of the three angles θ, α, and ε, θ is mainly determined by the pressure distribution on the airfoil surface, and is also related to the overall aerodynamic layout of the aircraft; α is determined by the orientation of the airfoil's mean aerodynamic chord line relative to the flight velocity vector, and is related to aerodynamic layout and flight conditions; ε is related to the three-dimensional effects of the airfoil, and is also related to α. The first step is to minimize the values ​​of these angles as much as possible, until they are zero. Further, while satisfying other requirements, we will explore ways to make the sum of the three angles θ, α, and ε negative.

[0076] Based on the above ideas, methods to reduce induced drag will be studied from two main directions:

[0077] I. Optimize airfoil design:

[0078] 1. Start by changing the pressure distribution of the airfoil, control and adjust the orientation of the airfoil's aerodynamic vector, so that the θ angle is as small as possible not only in the design state but also in the non-design state.

[0079] 2. Add airfoil auxiliary devices to locally adjust pressure distribution. For example, use slotted leading-edge flaps.

[0080] II. Optimize the overall aerodynamic layout design:

[0081] 1. Without changing the wing airfoil, adjust the orientation of the aerodynamic vector of the entire aircraft relative to the flight velocity vector by using a suitable overall aerodynamic layout to reduce the angle θ.

[0082] 2. Optimize flight conditions, tracking the maximum lift-to-drag ratio during key flight phases such as level cruise. Simultaneously, optimize flight speed and altitude to ensure the aircraft flies at the lowest possible angle of attack α.

[0083] 3. Utilize the synergistic effect of all components to generate an effect that reduces the angle between the aerodynamic vector and the vertical axis of the flight path system. For example, adopting a canard aerodynamic layout with a low aspect ratio, and using the upwash airflow generated by the canard on the outer wing of the main wing to reduce the three-dimensional effect, thereby achieving drag reduction.

[0084] 4. Improve the airfoil design to lower the stagnation point and generate a larger negative zero lift angle (around -10 degrees), and change the wing's mounting angle relative to the fuselage to reduce the θ angle.

[0085] 5. The leading and trailing edges of the airfoil are inverted, changing the θ angle to a negative angle and placing the upper surface of the wing in a positive pressure gradient state that is not easily separated. The leading edge becomes a wedge structure, which is beneficial to increasing the critical Mach number and reducing supersonic wave drag.

[0086] 6. Improve wingtip design to reduce three-dimensional efficiency losses. For example, install wingtip vortex-enhancing and drag-reducing devices to reduce induced drag.

[0087] The present invention also provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the steps of the method for reducing overall aircraft drag as described above.

[0088] The present invention also provides a computer device, the computer device including a memory, a processor and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to implement the steps of the method for reducing overall aircraft drag as described above.

[0089] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0090] The method and system for reducing overall aircraft drag provided by this invention not only minimize drag to the maximum extent, but also eliminate induced drag, transforming harmful induced drag into induced thrust that is beneficial to aircraft motion. By analyzing the generation mechanism and root causes of drag, especially induced drag, the correct direction for drag reduction is found, and methods for reducing, eliminating, or even changing and transforming induced drag are developed, turning negative energy drag into positive energy thrust. This can efficiently reduce overall aircraft drag, thereby significantly improving the lift-to-drag ratio and effectively enhancing flight efficiency. Attached Figure Description

[0091] Various other advantages and benefits will become apparent to those skilled in the art upon reading the following detailed description of preferred embodiments. The accompanying drawings are for illustrative purposes only and are not intended to limit the invention.

[0092] In the attached diagram:

[0093] Figure 1 is a flowchart of a method for reducing overall aircraft drag according to the present invention;

[0094] Figure 2 is a simplified diagram of the formation mechanism of the induced resistance of the present invention;

[0095] Figure 3 is a simplified diagram of the formation and formula derivation of the second induced resistance of the present invention;

[0096] Figure 4 is a schematic diagram of the configuration of a computer device according to an embodiment of the present invention. Detailed Implementation

[0097] Exemplary embodiments will now be described in detail, examples of which are illustrated in the accompanying drawings. When the following description relates to the drawings, unless otherwise indicated, the same numerals in different drawings denote the same or similar elements. The embodiments described in the following exemplary embodiments do not represent all embodiments consistent with this disclosure. Rather, they are merely examples of apparatuses and products consistent with some aspects of this disclosure as detailed in the appended claims.

[0098] The terminology used in this disclosure is for the purpose of describing particular embodiments only and is not intended to be limiting of the disclosure. The singular forms “a,” “the,” and “the” as used in this disclosure and the appended claims are also intended to include the plural forms unless the context clearly indicates otherwise. It should also be understood that the term “and / or” as used herein refers to and includes any and all possible combinations of one or more of the associated listed items.

[0099] It should be understood that although the terms first, second, third, etc., may be used in this disclosure to describe various information, such information should not be limited to these terms. These terms are used only to distinguish information of the same type from one another. For example, without departing from the scope of this disclosure, first information may also be referred to as second information, and similarly, second information may also be referred to as first information. Depending on the context, the word "if" as used herein may be interpreted as "when," "when," or "in response to determination."

[0100] The embodiments of the present invention will be further described in detail below with reference to the accompanying drawings.

[0101] Induced drag is an upward-pointing aerodynamic force R acting on the airfoil surface due to the difference in velocity between the upper and lower surfaces caused by Bernoulli's principle, resulting in different residual pressures relative to the incoming flow. The vector R typically forms an angle with the vertical and horizontal axes of the motion coordinate system (or the flight path coordinate system). This angle consists of two parts: the first part is determined by the shape of the airfoil's pressure distribution, which creates a skew angle θ (based on 90 degrees) between the aerodynamic force vector and the horizontal and vertical axes of the body coordinate system (or semi-body coordinate system). The cosine component RCos(θ) of the magnitude of the aerodynamic force vector R is the vertical force on the airfoil; the sine component Rsin(θ) of the aerodynamic force vector R is the component projected onto the body axis x. In flight mechanics, when applying aerodynamics, a matrix transformation between coordinate systems is ultimately used to project the force onto the flight path coordinate system. Typically, there is an angle α between the body coordinate system and the flight path coordinate system. Relative to the aerodynamic force vector R, the skew angle corresponding to the lift generated is further deflected to the right by an additional angle α. At this point, the total deflection angle is (θ+α). The aerodynamic resultant force R in the trajectory frame X h The directional sine on the axis becomes R*sin(θ+α).

[0102] The nature and magnitude of this component force depend entirely on the angle of deflection of the R vector relative to the vertical axis in the track axis system. The magnitude and sign of R*sin(θ+α) have three possible projections: positive and negative. Physically, this corresponds to no induced drag, induced thrust (negative drag), and induced drag.

[0103] Currently used airfoil geometries are almost 100% (θ+α), always with a negative sign, and the R vector is biased to the right. Therefore, the direction cosine of R is always drag, i.e., induced drag. Both the induction factor of the second induced drag and the correction coefficient δ are inversely proportional to the aspect ratio of the airfoil.

[0104] Therefore, the method for reducing or eliminating the second induced drag provided by this invention is to: increase the geometric aspect ratio of the wing or increase the effective aspect ratio of the wing (while keeping the geometric aspect ratio unchanged); and adjust the magnitude and sign of the skew angle of the aerodynamic resultant force vector R. Figure 1 shows the basic flow of the method for reducing overall aircraft drag according to this invention.

[0105] Having clarified the general direction of reducing induced drag, several specific solutions are proposed below, and the method for significantly reducing drag according to the present invention is further illustrated through several specific embodiments.

[0106] Example

[0107] Example 1

[0108] By adjusting the installation angle, induced drag can be completely eliminated from a regional jet, thus improving its flight performance.

[0109] Design a medium-sized regional jet, HY-024. Its reference model is the C-5XX. The main parameters of the prototype are as follows: length: 28.6 meters; wingspan L: 28.3 meters; height: 11.3 meters; empty weight: 44,100 kg; maximum takeoff weight: 72,500 kg; total seating capacity: 168 people.

[0110] Wing parameters: Aspect ratio A = 8.83, 1 / 4 chord sweep angle 25°, average relative thickness 12.89%.

[0111] For a mid-wing monoplane, using A = L² / S, the wing area is calculated as: S = 28.3² / 8.8³ = 90.7 m² 2

[0112] The average aerodynamic chord C = L / A = 28.3 / 8.83 = 3.205m 2

[0113] Powerplant: Two LEAP-1C turbofan engines mounted under the wings. Each engine has a thrust of approximately 71.2 kN; fuel consumption rate: Ce = 0.53 kg / kg·h; T = 2 × 7260 kg.T / W = 14520 / 72500 = 0.2.

[0114] Cruising altitude = 12,000 meters, M = 0.79, cruising speed V = 233.208 m / s = 839.55 km / h;

[0115] This aircraft employs a supercritical laminar flow airfoil design; the airfoil has a large relative thickness (relative thickness greater than 25% of the aerodynamic chord) greater than 25%C; the stagnation point at the airfoil's leading edge is moved downwards as much as possible; the airfoil has camber under the airfoil design conditions; the optimal deviation angle of the aerodynamic vector does not exceed 2.5 degrees. The zero-lift angle of attack of the airfoil is approximately -8 degrees. Based on the above data, the lift characteristics can be inferred. (The abbreviations in the calculation process below, ZWTA refers to the wingtip vortex lift enhancement device).

[0116] Let α0 = -4 degrees, C La =0.135, R0=8.83 / 10.83=0.8153;

[0117] After installing ZWTA, R1 = 0.9047, an improvement of 11%. L0 =0.135 × 4 = 0.54;

[0118] Formula for calculating lift: C L =0.54 + 0.1101 * α;

[0119] Design lift coefficient (12000 meters): C L =2*72500 / 0.0317*90.7*233.2082=0.9273;

[0120] Generate C L =0.9273 corresponds to an angle of attack of α = 8.422 degrees;

[0121] It needs to be deflected upwards by 8.422 - 4 = 4.422 degrees;

[0122] θ=1.5 degrees, α+θ=-4.422+1.5=5.922 degrees;

[0123] C D0 =0.020; K1=tan(5.922)=0.1037;

[0124] C Di1 =0.09619; K2=0.8153*(1.0 / 8.83 / 3.1416)=0.02939;

[0125] C Di2 =0.02939*0.92732=0.02527; C Di =0.12146;

[0126] CD0 =0.02; C D =0.020 + 0.12146 = 0.14146;

[0127] K = 0.9273 / 0.14146 = 6.556, the resistance is too high and the efficiency is not high.

[0128] The method of improving the original airfoil was adopted to eliminate drag encountered by the entire aircraft. The specific approach was to modify the airfoil locally. The leading edge was lowered, causing the stagnation point to shift downward. As a result, the zero-lift angle of attack became -9.5 degrees.

[0129] The required lift coefficient for cruise is still 0.9273. The airfoil needs to deflect 8.422 degrees relative to the aerodynamic chord. However, the new airfoil has a zero lift coefficient 1.078 degrees higher than required due to its -9.5 degree zero lift angle. To maintain balance, the aerodynamic chord of the airfoil (wing) needs to be deflected counterclockwise by 1.078 degrees. Since the aerodynamic vector of the airfoil is deflected by 1.5 degrees (to the right) in its design state, after the wing rotates around its side axis and pitches down 1.078 degrees, the aerodynamic vector deflects only 0.422 degrees from the vertical axis of the trajectory coordinate system. The tangent of this angle is 0.00737. The first induced drag is C. Di1 =0.00683, the second induced resistance remains unchanged.

[0130] C Di2 =0.02939★0.92732=0.02527; C D =0.0321;

[0131] K = 0.9273 / 0.0321 = 28.89, which is a significant improvement over the previous case using supercritical laminar airfoil design, increasing by 4.4 times.

[0132] Maximum range estimation: The flight conditions are as follows: altitude 12,000 meters, speed 839.55 km / h. To calculate the range, we need to know the flight speed, fuel consumption rate, lift-to-drag ratio, and fuel coefficient. Although there is insufficient data to calculate the fuel coefficient, we can reasonably estimate the fuel weight from the known information. Since the total number of passengers is known to be 168, and each person's weight and luggage are approximately 80 kg, we deduce the fuel coefficient to be 0.23. The final maximum range is: L = KV / Ce*γ = (28.89*839.55 / 0.53)*0.23 = 10525.58 km.

[0133] The maximum range before the optimized design was L = 2388.57 kilometers.

[0134] The prototype's maximum range is L = 5550 kilometers.

[0135] This shows that without careful design, the performance would be very poor, not as good as the prototype.

[0136] After adopting aerodynamic vector control technology that significantly reduces induced drag, flight performance has been greatly improved. The maximum range is 4.4 times that before the optimization design; it is 190% of the prototype, nearly twice as long.

[0137] As seen during the optimization design process, the current results are not optimal. Increasing the cruising altitude further, for example to 12,500 meters, would completely eliminate induced drag. The maximum range would then exceed 24,288.86 kilometers.

[0138] Example 2

[0139] Design an international logistics and express delivery aircraft with multiple wings, large payload, high speed, and long range.

[0140] The design incorporates a four-wing configuration for a logistics and express delivery aircraft. This is a heavy-duty logistics transport aircraft with a total weight exceeding 2024 tons. As is well known, a key technology in designing such massive aircraft is having lifting surfaces that generate sufficient lift. The characteristic value of the lifting surface is C. L S. This is a combined index used to express the lift contribution of each lifting surface. Since the reference area is the wing, each lifting surface is represented by its C. L The S-product expression eliminates the need to calculate the weights of each lifting surface. When using it, the dimensionless lift of all lifting surfaces is summed, and the total is used for calculation. For example, the total dimensionless lift of four lifting surfaces is: C L S = C L1 S1+C L2 S2+C L3 S3+C L4 S4;

[0141] The stall velocity is V = [(2W / (ρ*C)] L S)] 1 / 2 ;

[0142] In this embodiment, W = 2024 x 10 3 The tail guard angle at takeoff is 11 degrees, the zero-lift angle of attack is -7.8 degrees, and the total angle of attack is 18.8 degrees. The lift formula is: C L =0.1101*(α+7.8) =0.1101*18.8 =2.07. Assuming each wing uses the same airfoil, therefore the C of all wings... L =2.07, all are the same. However, the areas of each wing are different. The area ratio of the four wings is known: (based on the area of ​​the largest wing) C LS = 1.592 * S1, and their proportions are: S1:S2:S3:S4 = 0.6281:0.1281:0.1853:0.0584;

[0143] First, using empirical data on wing loading, the total lifting surface area S of the jumbo jet is determined: the wing loading of the super-large aircraft is taken as 586 kg / m. 2 S = 2024000 / 586 = 3454 meters 2 C L S = 3454 * 2.07 = 7149.62

[0144] Stall velocity V = [(2024000*2) / (0.125*7149.62)] 1 / 2 =[(4529.47)] 1 / 2 =67.3 m / s = 242.28 km / h; taking a thrust-to-weight ratio of 0.20, the required total thrust is T = 0.20 * 2024 tons = 404.8 tons. If four turbofan engines are installed, the thrust of each engine is T0 = 101.2 tons = 992.772 kN. The thrust of a single LEAP 1A engine is only 71.2 kN. The engine in this embodiment is 14 times that of the C919 engine. Nuclear power engines may be required.

[0145] Estimate takeoff distance L = μ / (n x *C Lmax = (586 / 0.125 / 9.81) / 0.2 / 2.07 / 0.343 = 3365.3 meters; the track length needs to be at least 3500-4000 meters.

[0146] Estimated range: Assuming a fuel coefficient of 0.20 and a cruising altitude of 20,000 meters, the cruising speed is high subsonic Mach 0.8 (850.18 km / h), and the fuel consumption rate Ce = 0.52 kg / kg·h.

[0147] Lift-to-drag ratio estimation: The lift coefficient under design conditions is 1.0 (20,000 meters);

[0148] Resistance calculation:

[0149] Induced drag elimination technology is adopted. Only a zero-lift drag coefficient is achieved for the entire aircraft. This is typically 0.02 * 1.592 = 0.03184.

[0150] The lift-to-drag ratio K = C L / C D =1.0 / 0.03184=31.41

[0151] Maximum range is L max = (850.18 * 31.41 / 0.52) * 0.20 = 10271 kilometers.

[0152] With additional fuel and a fuel coefficient of 0.5 (in-flight refueling), the maximum range increases to 25,677 kilometers. Based on a 23% percentage of the mission load, the effective payload is calculated as 20,240,000 * 0.23 = 465,520 kg, or 465.5 tons. This translates to 5,000 fully armed soldiers, 19 fighter jets (25 tons each), or 6,000 ordinary passengers.

[0153] Example 3

[0154] An amphibious flying car has been developed. The scaled-down model flight test was successful. However, a drawback is that due to the use of an ultra-short aspect ratio wing as the lifting body, although the installation of wingtip vortex generators significantly increased lift, the induced drag remained quite high, resulting in excessive energy consumption and affecting flight range. A solution is needed. The inventors of this application undertook the demonstration work and proposed a feasible solution: utilizing dipole wings to reduce most of the induced drag, as illustrated below:

[0155] The main parameters of the flying car are as follows:

[0156] Ultra-low aspect ratio - capable of short-distance or vertical takeoff and landing within a single lane on a highway.

[0157] b = 1.8 meters width

[0158] h = 1.8 meters (vehicle height)

[0159] V = 300 km / h cruising speed

[0160] H = 3000 meters cruising altitude

[0161] Flight range L = 1000 kilometers

[0162] Propulsion: ROTAX 914 130HP (Hybrid)

[0163] Fuel weight: 120 liters

[0164] Maximum takeoff weight W = 650 kg

[0165] The horizontal projected area of ​​the vehicle body, Sw, is 9.0 square meters.

[0166] Propulsion: ROTAX 914130HP (Hybrid)

[0167] Fuel weight: 120 liters

[0168] The airframe is designed with an airfoil section with a very low aspect ratio that provides lift: it adopts the laminar flow high airfoil QYX-5, which has a large thickness, high lift, and high lift-to-drag ratio; it has camber; and its stagnation point is low.

[0169] The maximum lift-to-drag ratio exceeds 200.

[0170] The development of this flying car, from its initial concept to its first scaled-down model taking to the skies, took over a decade. This is due to its unique aerodynamic shape; it has virtually no wings, making it a wingless aircraft. Its aspect ratio is only 0.36, and its three-dimensional aerodynamic coefficient is only 15.25%, resulting in a lift loss of 84.75%. Two wind tunnel tests were conducted at an aerospace university, and after two years of repeated theoretical and practical trials, a protective barrier—wingtip vortex-based lift enhancement and drag reduction technology—was finally developed to address the significant drop in lift caused by the ultra-low aspect ratio wings. On March 25, 2023, the dream of a wingless aircraft taking to the skies was realized. The wingless flying car can achieve short takeoff and landing; climb; level flight; maneuvering; and landing. The previously worried lateral stability has not presented any problems. The only drawback is the excessive induced drag due to the relatively small aspect ratio. This is attributed to the inherent limitations of this ultra-short aspect ratio lift-generating fuselage. A short aspect ratio has the fatal drawback of simultaneously reducing lift and increasing induced drag. The latter can be more serious. Therefore, although wingtip vortex-based lift-enhancing and drag-reducing technology has increased lift to meet the basic requirements for takeoff, the reduction in drag is still insufficient, and another solution must be found.

[0171] Mathematical Analysis: Two formulas are used to specifically calculate and compare the changes in lift and drag before and after the installation of the guardrail. The change in lift is compared using the three-dimensional efficiency factor: R = A / (A+2). Before the ZWTA device was installed, A = 0.36, R0 = 0.36 / 2.36 = 0.1525, resulting in a lift loss of 84.75%. After the ZWTA device was installed, the equivalent effective aspect ratio became A1 = 2 / (1 / R0-1) = 0.5627, the lift loss coefficient R1 = 0.5627, and the lift loss was only 43.83%, which is 40.92% of the original 51.7% without the ZWTA. The effective aspect ratio increased from 0.36 to A1 = 2 / (1.0 / R1-1) = 2.5735. The effective aspect ratio is 7.15 times the original 2.5735 / 0.36. The equivalent wingspan becomes 1.8 x 7.15 = 12.87 meters.

[0172] Calculate the change in resistance before and after: C Di2 =R*(1.0 / (3.1416*A)*C L 2 C Di1 =Tan(α+θ); C D0 =0.020; CD=C D0 +C Di +C Di2 ;

[0173] Before the improvement: During takeoff: θ + α = 1.5 + 9.5 = 11 degrees; α = 7.8 + 9.5 = 17.3 degrees; C L =0.1101*α=0.1101*17.3=1.905 C L 2 =3.628; C Di1 =Tan(α+θ)*C L =0.1944 * 1.905 = 0.3703; C D =0.020+0.3703+0.4893=0.8796; K=C L / C D =1.905 / 0.8796 = 2.166;

[0174] After improvement: During takeoff, θ + α = 1.5 + 9.5 = 11 degrees; α = 7.8 + 9.5 = 17.3 degrees; R1 = 0.5627; A1 = 2.5735; C L =1.905 C L 2 =3.628; C Di1 =Tan(α+θ)*C L =0.1944 * 1.905 = 0.3703; C Di2 =R*(1.0 / (3.1416*A)*C L 2 =0.5627*0.1237*3.628=0.2525; C D =0.020+0.3703+0.2525=0.6428; K=1.905 / 0.6428=2.9636 τ=2.9636 / 2.166=136.82%;

[0175] K increased by only 36.82%.

[0176] This shows that the lift of flying cars has increased significantly, from 15.25% to 56.27%, which is 369% of the original, or 3.69 times.

[0177] This embodiment adds a counter-rotating wing to the existing configuration. Specifically, the trailing edge of the airfoil faces forward, and the leading edge faces backward. The aerodynamic vector of this counter-rotating wing is tilted forward, counteracting and eliminating most of the induced drag generated by the lifting fuselage. Since the area of ​​the auxiliary wing is smaller than the horizontal projected area of ​​the lifting fuselage, it can only counteract a large portion of the induced drag, leaving some drag remaining. To achieve more complete drag reduction, other methods can be used in conjunction. For example, the leading and trailing edges of the airfoil-equipped lifting body can be inverted by 180 degrees, changing the orientation of the aerodynamic vector to point directly forward of the flight velocity vector. In this case, the counter-rotating wing may not even be necessary, thus reducing costs.

[0178] This invention also provides a computer device. Figure 4 is a schematic diagram of the structure of a computer device provided in this invention. Referring to Figure 4, the computer device includes: an input device 23, an output device 24, a memory 22, and a processor 21. The memory 22 is used to store one or more programs. When the one or more programs are executed by the one or more processors 21, the one or more processors 21 implement the method for reducing the overall flight drag of the aircraft as provided in the above embodiments. The input device 23, the output device 24, the memory 22, and the processor 21 can be connected by a bus or other means. Figure 4 shows an example of connection via a bus.

[0179] The memory 22, as a read / write storage medium for a computing device, can be used to store software programs and computer-executable programs, such as the program instructions corresponding to the method for reducing overall aircraft drag described in this embodiment of the invention. The memory 22 may primarily include a program storage area and a data storage area. The program storage area may store the operating system and at least one application program required for a function; the data storage area may store data created based on the use of the device. Furthermore, the memory 22 may include high-speed random access memory and non-volatile memory, such as at least one disk storage device, flash memory device, or other non-volatile solid-state storage device. In some instances, the memory 22 may further include memory remotely located relative to the processor 21, and these remote memories can be connected to the device via a network. Examples of such networks include, but are not limited to, the Internet, corporate intranets, local area networks, mobile communication networks, and combinations thereof.

[0180] Input device 23 can be used to receive input digital or character information, and generate key signal inputs related to user settings and function control of the device; output device 24 may include display devices such as a display screen.

[0181] The processor 21 executes various functional applications and data processing of the device by running software programs, instructions and modules stored in the memory 22, thereby realizing the above-mentioned method of reducing the overall flight drag of the aircraft.

[0182] The computer equipment provided above can be used to execute the method for reducing overall aircraft drag provided in the above embodiments, and has corresponding functions and beneficial effects.

[0183] This invention also provides a storage medium containing computer-executable instructions, which, when executed by a computer processor, are used to perform the method for reducing overall aircraft drag as provided in the above embodiments. The storage medium can be any type of memory device or storage device, including: mounting media such as CD-ROM, floppy disk, or magnetic tape; computer system memory or random access memory such as DRAM, DDRRAM, SRAM, EDORAM, Rambus RAM, etc.; non-volatile memory such as flash memory, magnetic media (e.g., hard disk or optical storage); registers or other similar types of memory components; the storage medium may also include other types of memory or combinations thereof; furthermore, the storage medium may reside in a first computer system in which the program is executed, or it may reside in a different second computer system connected to the first computer system via a network (such as the Internet); the second computer system can provide program instructions to the first computer for execution. The storage medium includes two or more storage media that may reside in different locations (e.g., in different computer systems connected via a network). The storage medium may store program instructions (e.g., specifically implemented as a computer program) executable by one or more processors.

[0184] Of course, the storage medium containing computer-executable instructions provided in the embodiments of the present invention is not limited to the method for reducing the overall flight drag of the aircraft as described in the above embodiments, but can also perform related operations in the method for reducing the overall flight drag of the aircraft provided in any embodiment of the present invention.

[0185] The technical solution of the present invention has been described above with reference to the preferred embodiments shown in the accompanying drawings. However, it will be readily understood by those skilled in the art that the scope of protection of the present invention is obviously not limited to these specific embodiments. Without departing from the principles of the present invention, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the scope of protection of the present invention.

[0186] The above description is merely a preferred embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A method for reducing overall aircraft drag, characterized in that, Includes the following steps: S1. Design to reduce zero-lift resistance and reduce or eliminate induced resistance; The methods for reducing or eliminating induced drag include: controlling the orientation of the aerodynamic vector; the methods for controlling the orientation of the aerodynamic vector include: adjusting the skew angle of the aerodynamic vector from the vertical axis of the track axis system, and / or, increasing the aspect ratio of the wing or increasing the effective aspect ratio of the wing. The method for adjusting the skew angle of the aerodynamic vector from the vertical axis of the track axis system includes: controlling the azimuth angle θ of the airfoil aerodynamic resultant force vector to be within 0-0.2 degrees when the angle of attack α is zero, so that the lift coefficient of the airfoil is as large as possible near the angle of attack α where the maximum lift-to-drag ratio coefficient occurs, and then deriving the airfoil surface pressure distribution and the corresponding airfoil profile geometry that satisfy the optimization index of limiting the azimuth angle θ of the airfoil aerodynamic resultant force vector to be within 0-0.2 degrees when the angle of attack α is zero; S2. Optimize the pressure distribution on the airfoil surface and calculate the airfoil geometry. The method for calculating the airfoil geometry includes: The geometry of the existing airfoil profile is locally modified, the airfoil midsection is extended to form a flat platform; the airfoil leading edge is bent downwards; the stagnation point is lower; the zero lift angle of attack is close to -5 degrees to -10 degrees; the thickness of the airfoil tail section is increased; the lower wing surface is bent upwards, and the skewness of the R-vector of the entire aircraft is controlled to be less than 1 degree in the high-speed cruise phase.

2. The method for reducing overall aircraft drag according to claim 1, characterized in that, The design of step S1 to reduce zero-lift drag and reduce or eliminate induced drag includes: constructing a mathematical model of the overall aircraft's flight drag. This mathematical model is based on the total drag in flight: the sum of zero-lift drag and induced drag. The expression for the total drag in flight is: C D =C D0 +C Di =C D0 +K1*C L +K2*C L 2 =C D0 +Tan(α+θ)*C L +sign(K1)*kR / πA*C L 2 (1) In equation (1), C D0 It is zero-lift resistance; C Di It is induced resistance, K2*C L 2 It is C Di With lift coefficient C L A quadratic term proportional to the square of ; K2 = kR / πA; K2 is called the induction factor; K1*C L It is C Di With lift coefficient C L The linear term is proportional to the airfoil; K1 is the drag factor; A is the effective aspect ratio of the three-dimensional airfoil; π is pi; k is the error correction coefficient; R = A / (A+2) is the three-dimensional effect coefficient of the airfoil; α is the angle of attack; and θ is the angle between the aerodynamic vector of the two-dimensional airfoil and the vertical axis of the airframe coordinate system.

3. The method for reducing overall aircraft drag according to claim 2, characterized in that, The design of step S1 to reduce zero-lift drag and reduce or eliminate induced drag also includes: constructing a mathematical model of total induced drag, which is based on the sum of two parts of total induced drag in flight: first induced drag and second induced drag. The expression for the total induced drag in flight is: C Di =C Di1 +C Di2 =K1*C L +K2*C L 2 =Tan(α+θ)*C L +sign(K1)*(kR / πA)*C L 2 (2) In equation (2), C Di1 It is the first induced resistance, C Di2 It is the second induced resistance; C L K is the lift coefficient; K1 is the drag factor; k is the error correction coefficient; R is the airfoil three-dimensional effect coefficient; A is the effective aspect ratio of the three-dimensional airfoil.

4. The method for reducing overall aircraft drag according to claim 1, characterized in that, The method for designing the airfoil geometry in step S2 also includes: By changing the relative orientation of the airfoil and fuselage, and altering the relative orientation of the aerodynamic vector with respect to the track coordinate system, induced drag can be reduced or even eliminated, and induced drag can be converted into induced thrust. The methods for changing the relative orientation of the airfoil and the fuselage include one or a combination of two of the following: reversing the orientation of the leading and trailing edges of the airfoil, and adjusting the wing installation angle in the negative direction.

5. The method for reducing overall aircraft drag according to claim 4, characterized in that, The methods for eliminating induced drag include: using dipole wings or counter-switch wings; The dipole wing consists of two identical wings mounted on the fuselage, with the leading edge of one wing pointing forward and the trailing edge of the other pointing forward. When the aircraft moves forward, the aerodynamic vectors are opposite in direction to the flight speed, and their projection components on the horizontal axis of the track system cancel each other out, resulting in zero induced drag.

6. The method for reducing overall aircraft drag according to claim 5, characterized in that, The counter-dipole wing is a single wing mounted on the fuselage, forming a quasi-dipole wing pair with the existing wing layout to counteract induced drag.

7. The method for reducing overall aircraft drag according to claim 1, characterized in that, The method for designing the airfoil geometry in step S2 also includes: Add a leading-edge slotted suction attachment to the existing airfoil to generate leading-edge suction.

8. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the program is executed by the processor, it implements the steps of the method for reducing overall aircraft drag as described in any one of claims 1-7.

9. A computer device, the computer device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the program, it implements the steps of the method for reducing overall aircraft drag as described in any one of claims 1-7.

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