Aviation dynamic pressure difference all-wing aircraft

Abstract

The invention relates to an aviation dynamic differential pressure all-wing aircraft. An airfoil profile of the aircraft comprises an upper airfoil surface and a lower airfoil surface. The section lines of the airfoils are all composed of curves of a monotonically decreasing function from the front edge to the rear edge. The structure is simple, and when the lower wing surface does work on windward airflow, the flow speed is reduced, the static pressure intensity is increased, and upward thrust is generated; the upper airfoil does not do work, airflow flows back, static pressure is reduced, upward pulling force is generated by the airfoils, and therefore the double-airfoil lift airfoil profile with the upper airfoil and the lower airfoil simultaneously and consistently upward lift force is formed, the sum of the lift force of the two surfaces is larger than that of an existing single-face lift airfoil profile, and the airfoil efficiency is further improved by optimizing the lift-drag ratio. By adopting the wing section line and the spanwise line rear corner, shock waves are weakened to be low-sound-velocity object plane static pressure, damage of the shock waves is avoided, and the strong resistance effect is reduced.

Description

Technical Field

[0001] The present invention belongs to the field of aviation manufacturing technology, and relates to aviation full-wing aircraft, such as full-wing aircraft, flying-wing aircraft, etc. Background Art

[0002] Existing all-wing aircraft only have wings from the outside, and their airfoils are the same as those of fixed-wing aircraft. The lower surface of the airfoil is flat, while the upper wing is curved upward. At a zero-degree elevation angle, lift is provided by the principle of velocity difference lift. The upper wing has a fast flow velocity, and the static pressure decreases, forming a unilateral lift force. In addition, when the lower wing has an elevation angle, the flow velocity slows down, the static pressure increases, and the airflow generates thrust on the wing in an upward direction. If the forces of the upper and lower wing surfaces can be superimposed and manifested at the same time, the airfoil efficiency will be improved.

[0003] In the Chinese patent "202121320870.9 A ship dynamic pressure differential propeller", the principle of dynamic pressure differential lift of incompressible water fluid is applied to achieve an airfoil that can generate lift on both the upper and lower wing surfaces at the same time, and the efficiency is significantly improved. Therefore, it is promoted and applied to compressible air fluids, so that the airfoil of the aviation full-wing aircraft also has a double-sided lift wing surface at the same time, which can also improve its efficiency. Summary of the invention

[0004] The object of the present invention is to provide a double-sided lift airfoil capable of improving aviation air propulsion efficiency, and an aviation dynamic pressure difference full-wing aircraft, so as to solve the above-mentioned problems.

[0005] In order to achieve the above object, the technical solution of the present invention is achieved as follows:

[0006] An aerodynamic pressure differential all-wing aircraft, a wing cross-section diagram, includes an upper wing surface and a lower wing surface, characterized in that the cross-sections of the upper wing surface and the lower wing surface are both curved surfaces composed of monotonically descending curves from the leading edge to the trailing edge.

[0007] The main wing surface of the upper wing is a descending surface, and the surface is composed of a monotonically descending function. That is, from the leading edge to the trailing edge, the height of the vertical section of the surface decreases gradually, and generally there is no inflection point.

[0008] The main wing surface of the lower wing is a descending surface, and the surface is composed of a monotonically descending function. That is, from the leading edge to the trailing edge, the height of the vertical section of the surface decreases gradually, and generally there is no inflection point.

[0009] At near-sonic and supersonic speeds, the leading edge of the upper wing and the leading edge of the lower wing directly intersect.

[0010] At near-sonic and supersonic speeds, the rearward turning angle of the leading edge point of the lower wing section line is equal to or greater than the required normal angle of the leading edge point.

[0011] At near-sonic and supersonic speeds, the section line of the lower wing surface is the rearward turning angle from the leading edge to the trailing edge, and the rearward turning angle of the latter point is equal to or greater than the rearward turning angle of the former point.

[0012] At near-sonic and supersonic speeds, the rear turning angle of the leading edge point of the spanwise line is equal to or greater than the required normal angle of the point.

[0013] At near-sonic and supersonic speeds, the spanwise section line is a rearward turning angle from the leading edge to the trailing edge such that the rearward turning angle of the latter point is equal to or greater than the rearward turning angle of the former point.

[0014] Increasing the normal angle of the wing section can improve the lift-to-drag ratio.

[0015] Increasing the normal angle of the spanwise section can reduce drag.

[0016] Compared with the prior art, the advantages of the embodiments of the present invention are as follows: the present invention has a simple structure, the lower wing does work on the windward airflow, the flow rate is slowed down, the static pressure is increased, and the wing surface generates an upward thrust; the airflow on the upper wing surface flows back, the static pressure is reduced, and the wing surface generates an upward pull, thereby simultaneously forming the upper and lower wing surface lifts that are consistent and upward, and the sum of the two lifts is greater than the existing single-surface airfoil lift. After optimizing the lift-to-drag ratio, the airfoil efficiency is further improved. When the airflow is near the speed of sound and supersonic speed, the rear angle of the wing section line and the spanwise line is used to weaken the shock wave into a low-speed sound surface static pressure, thereby reducing flight resistance. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 Schematic diagram of horizontal non-thick pipe

[0018] Figure 2 Schematic diagram of horizontal straight uniform flow

[0019] Figure 3 Schematic diagram of static pressure calculation on the frontal surface

[0020] Figure 4 Schematic diagram of static pressure calculation on the back flow surface

[0021] Figure 5 Schematic diagram of the cross section of the full wing

[0022] Figure 6 Horizontally symmetrical diagram of the left part

[0023] Figure 7 Transition diagram after the horizontal length is doubled

[0024] Figure 8 Schematic diagram of the new airfoil with a length of 75%

[0025] Fig. 9 Schematic diagram after optimization and improvement DETAILED DESCRIPTION

[0026] The following will clearly and completely describe the technical solutions of the embodiments of the invention in conjunction with the figures. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0027] Example

[0028] 1. Lift principle of all-wing aircraft due to aviation dynamic pressure difference

[0029] The lift principle, main reasoning and conclusions of the aerodynamic pressure difference all-wing aircraft can apply the lift principle, main reasoning and conclusions of the water fluid dynamic pressure difference given in the Chinese patent "202121320870.9 A ship dynamic pressure difference propeller". Because water and air are both incompressible fluids (when it is lower than 0.2 times the speed of sound, air can be considered as an incompressible gas), it is enough to replace the water parameters with incompressible air parameters and make the same reasoning on the concepts. When the air flow rate is equal to or greater than 0.2 times the speed of sound, the above formulas, reasoning and conclusions can be expanded with compressible fluids, perfect gases (air is considered to be a perfect gas), isentropic, and high-speed one-dimensional steady flow, and can be applied to the airfoils of low-sonic aerodynamic pressure difference all-wing aircraft; at near-sonic and supersonic speeds, combined with the influence of shock waves, the lower wing section and the spanwise section rear angle are applied to weaken the surface shock wave static pressure to the low-sonic surface static pressure, thereby reducing flight resistance.

[0030] 2. The full-wing aircraft moves in the air flow field and does work

[0031] Basic formula for work in pipeline fluid dynamics

[0032] Assuming that the ideal irrotational flow field is a constant flow, the mass force is only gravity, the flow is along the element flow (streamline), and the gas flow is incompressible, the integral of the differential equation of motion along the streamline, plus the input mechanical energy, such as the fan, etc., the total flow Bernoulli equation is obtained as:

[0033]

[0034] For a straight smooth circular pipe without thickness, placed horizontally in the gas flow field, the position term can be ignored, and we get:

[0035]

[0036] In the formula: z is the height from a certain reference surface; u is the flow velocity; p is the static pressure; g is the acceleration due to gravity; ρ is the density, subscript 1 is the initial section, 2 is the final section, H m is the mechanical energy obtained by unit weight of fluid through fluid machinery, is the external energy, and α is the kinetic energy correction coefficient, where 1 is taken.

[0037] Multiply both sides of the equation by ρg, and let ρg H m =H, H is the input mechanical energy, called external pressure, and after sorting, we get:

[0038]

[0039] The parameters in (1) and (2) are the centroid parameter values. For a circular pipe, the centerline of the pipe is the centroid. The above formula is used for incompressible air fluids with a velocity lower than 0.2 times the speed of sound. It can be used for aviation fixed-wing airfoils and aviation full-wing airfoils. Air parameters should be substituted when applying.

[0040] Assume that a straight pipe with limited length, smooth, round shape and small radius R is placed horizontally in the air flow field. The pipe is stationary, the coordinate axis is fixed on the pipe, and the horizontal axis coincides with the center line of the pipe. B is the inlet end, C is the external machine, A is the outlet end, and the air flow in the pipe flows from right to left. The schematic diagram of the horizontal non-thick pipe is as follows: Figure 1 shown.

[0041] When the air velocity in the flow field is zero and the steady velocity in the pipe is u, any point in the inlet pipe BC is selected as section 1, and any point in the outlet pipe AC is selected as section 2.

[0042] At section 1, there is a velocity u, flowing from right to left. Its kinetic energy is converted from the static energy of the flow field (also called background static energy). Therefore, the static energy of pipe BC is twice as low as the background static energy. At section 2, there is a velocity u, but the air needs to be discharged to the outlet. Therefore, the static energy must be equal to or greater than the background static energy to discharge the gas smoothly. Here, it is a constant velocity and is equal to. Substituting into formula (1), the work done is:

[0043]

[0044]

[0045]

[0046]

[0047] This means that as long as the air velocity in the tube is greater than zero, the external machine will always do work, and the amount of work done is one times the kinetic energy. 背 is the static pressure intensity of the background flow field.

[0048] If we imagine the flow field as a pipe of infinite diameter with a horizontal straight pipe of finite length and small diameter inside, then, Figure 1 The small tube and the large tube form a fluid circuit. As long as the airflow passes through the external mechanical part C, the energy will be increased.

[0049] In this way, the entire flow field consists of three parts: outlet pipe, large pipe (background flow field), and inlet pipe. The flow direction is: the inlet pipe section sucks the flow field fluid from the inlet end, and after the energy is increased by external force, it is discharged into the outlet pipe section, flows into the background flow field through the outlet end, and then flows to the inlet end through the background flow field.

[0050] The energy of the entire flow field is divided into two areas, among which the energy in the outlet pipe AC is one area, which has twice the kinetic energy than other areas.

[0051] If formula (2) is used and the effect of velocity is taken into account, the static pressure of the entire flow field can be divided into three regions: the outlet section has the highest static pressure (when the fluid is set to a stationary coordinate, the velocity is the dynamic pressure, which can be converted into static pressure), the background flow field has the second highest static pressure, and the inlet section has the lowest static pressure, with the difference being one times the dynamic pressure.

[0052] Use a non-thick rigid flat sheet with a radius equal to the inner radius R of the pipe, perpendicular to the paper surface, and drive the airflow to move to the left along the center line of the pipe at a constant speed u, replacing the external mechanical work. The amount of work can be calculated using formula (1). At this time, the airflow on the right surface of the sheet, due to the leftward movement of the sheet, the inlet pipe section becomes longer, and the incompressible airflow has not yet flowed in from the inlet end. The airflow and the sheet are separated, and the separation speed is u. This space is a vacuum. At the inlet end, the airflow is the static pressure of the flow field, and the right side of the sheet is a vacuum. Under the action of the static pressure difference between the two, the flow field airflow flows into the inlet pipe section to fill the vacuum area. When the flow speed reaches u, it is equal to the separation speed of the sheet, reaching flow equilibrium, and the vacuum disappears. From formula (2), it is known that the static pressure of the inlet pipe section is twice the dynamic pressure of the static pressure of the flow field.

[0053] When the pipe is extremely short, formula (1) still applies; when it tends to be infinitely short, it still applies. When it is equal to zero, it is equivalent to the individual motion of the sheet without the pipe section. For the sheet with and without the pipe, assume that both are doing the same horizontal motion from right to left, with a speed of u. The static pressure of the flow field on the left side of the two sheets is the same, which is equal to the static pressure of the background flow field. Due to the movement of the sheet, the air is separated from the sheet on the back side of the right side of the two sheets, and a vacuum is generated on the back side of the sheet. The gas on the right side of the sheet is the static pressure of the background flow field. The air between the static pressure of the flow field and the vacuum moves to the vacuum area under the action of the static pressure of the flow field. When the flow balance is reached, the flow velocity is u, the vacuum area disappears, and the static pressure on the right side of the sheet is the dynamic pressure of the static pressure of the flow field dropping by one times. Therefore, the two sheets are subjected to the same force, and the work done by both is equal. The amount of work done can be calculated using formula (1) with pipes. The motion of the sheet without pipes can be regarded as the motion of the sheet with pipes. The two sheets do the same horizontal motion in the flow field and the amount of work done is equal.

[0054] When a non-thick rigid flat sheet moves alone in a static flow field, the gas energy in front of it is higher and the static pressure on the back is the lowest.

[0055] The all-wing aircraft moves in the air fluid and can be considered as a rigid body. Its surface can be divided into countless small pieces. Formula (1) can be applied to each small piece to calculate the micro-work. Finally, by summing them up, we can get the work done by the all-wing aircraft.

[0056] In the flow calculation of pipeline fluid, when the air velocity is at an angle with the centerline of the pipeline, the velocity needs to be decomposed into two components, one parallel to the centerline and the other perpendicular to it. The parallel component forms the pipeline flow and constitutes the work flow, while the perpendicular flow does not do work. Therefore, we only need to find the velocity component parallel to the centerline and substitute it into the formula for calculation.

[0057] After changing the conditions of formula (1) to compressible gas, adding perfect gas, isentropic, and one-dimensional high-speed steady flow, the formula for pipeline work is:

[0058]

[0059] Where γ is the specific heat ratio (the specific heat ratio of air is 1.4), subscripts 1 and 2 are the corresponding parameters of pipe sections 1 and 2 respectively, and the usage is the same as formula (1).

[0060] Similarly, when an aeroplane moves in a compressible air flow field, its surface can be divided into countless small pieces. Formula (3) can be applied to each small piece to calculate the micro-work (quantity). Finally, by summing them up, we can get the work done by the aeroplane.

[0061] 3. Aeroplane full wing airfoil section static pressure strength

[0062] It is divided into incompressible airfoil section static pressure and compressible airfoil section static pressure.

[0063] The incompressible full-wing airfoil section is the wing airfoil section of a fixed-wing aircraft. Under incompressibility, they are the same as the airfoil section of the lower wing of a ship, so the fluid formula and results of the underwater wing can be directly applied, but the parameters are air parameters. The compressible one needs to use the incompressible theory extension. At low speed, there is no need to calculate the rear turning angle; but at near-sonic and supersonic speeds, it is necessary to calculate the rear turning angle of the lower wing and the spanwise section to achieve the purpose of weakening the shock wave to the static pressure of the low-sonic object surface and reducing drag.

[0064] Incompressible airfoil section static pressure strength.

[0065] There is no left and right (front and back) static pressure difference for a thick rigid flat sheet.

[0066] The coordinate origin is fixed to the center point of the sheet, the X axis is placed horizontally, the direction is right, and the Y axis is upward. The air flow in the flow field becomes a horizontal straight uniform flow, moving from left to right. Figure 2 shown.

[0067] According to the incompressible Bernoulli level formula in the flow field:

[0068]

[0069] Find the static pressure p on the left side of the film 片左 Select the far left end of the flow field as section 1, the left side of the sheet as section 2, and substitute into formula (4) to obtain:

[0070]

[0071]

[0072] That is, the left side of the film converts the dynamic pressure of the gas at the far end into static pressure. The static pressure on the left side of the film is p 片左 The dynamic pressure is twice as much as the static pressure of the flow field.

[0073] Find the static pressure p on the right side of the plate 片右 The air in the flow field moves horizontally to the right at a speed of u, and the gas separates from the right side of the sheet at a speed of u. The separation area is a vacuum. Under the action of the static pressure difference between the static pressure of the flow field and the vacuum, the air quickly moves along the horizontal line to the right back side of the sheet. When dynamic equilibrium is reached, the speed of the air moving to the back side is u, and the static pressure on the back side of the sheet is the dynamic pressure of the static pressure of the flow field down by one times, that is:

[0074]

[0075]

[0076] The difference in static pressure between the left and right sides of the sheet is 2 times the dynamic pressure:

[0077] p 片左右差 =p 片左 -p 片右 =ρu 2 .

[0078] The cross section of an aviation wing is a closed smooth transition curve, which we represent with a circular curve. By finding the static pressure at any point on the curve, we can apply it to find the static pressure at any point in the cross section of the wing. When calculating the area, take 1 unit length in the span direction.

[0079] Assume that the section of the wing cross-section diagram is a circular curve with a radius of R. The circular curve is parallel to the surface. The infinitely far horizontal parallel air flow flows from left to right at a velocity u. The origin of the coordinate axis is fixed at the same position as the center of the circle. The horizontal axis X is placed horizontally and points to the right. The vertical axis Y is vertically upward. The normal direction of the curve is positive within the static pressure of the air. Point A is the intersection of the straight uniform airflow and the circumference. Line B is the normal to point A. Line C is the tangent to point A. α is the angle between the straight uniform airflow and the normal. According to the meaning of formulas (2) and (4), on the frontal surface, in the second and third quadrants, the fluid collides head-on, which is the normal component. The normal velocity becomes zero, and the dynamic pressure will be completely converted into static pressure. The direction of the increased static pressure points to the center of the circle. The tangential component has no collision and remains unchanged. The schematic diagram of the static pressure calculation on the frontal surface is as follows: Figure 3 shown.

[0080] According to the meaning of formulas (2) and (4), the normal flow velocity = u × cosα, and the sum of the background static pressure and the added static pressure is the static pressure value at point A on the frontal surface:

[0081]

[0082] The second and third quadrants are symmetrical figures with the horizontal axis. The static pressure values ​​at the symmetrical points are equal in magnitude and along the inner normal direction. The newly added static pressure is thrust, which is equal in magnitude and also along the inner normal direction. The newly added static pressure component formed in the horizontal direction is resistance, which is equal in magnitude and in the same direction. The newly added static pressure component formed in the vertical direction is lift, which is equal in magnitude. In the second quadrant, the component is negative lift, and in the third quadrant, it is positive lift.

[0083] The back flow surface is composed of the first and fourth quadrant curves of the circular curve. Point A is the intersection of the reverse straight uniform airflow (the velocity is equal to that of the straight uniform airflow, but the direction is opposite) and the circumference. Line B is the normal line of point A, line C is the tangent line of point A, and α is the angle between the reverse straight uniform airflow line and the normal line. Figure 4 shown.

[0084] According to the meaning of formulas (2) and (4), the normal flow velocity = u × cosα, and the normal flow velocity on the backflow surface is converted from the static pressure of the background flow field, which reduces the static pressure. The difference between the background static pressure and the newly added static pressure is the static pressure value of point A on the backflow surface:

[0085]

[0086] The direction of the static pressure is the same as the inner normal, pointing toward the center of the circle, but it is decreasing.

[0087] The first and fourth quadrants are symmetrical figures with the horizontal axis. The static pressure values ​​at the symmetrical points are equal in magnitude and along the inner normal direction. The newly added static pressure is pulling force (suction force), which is equal in magnitude and along the outer normal direction. The newly added static pressure component formed in the horizontal direction is resistance force, which is equal in magnitude and in the same direction. The newly added static pressure component formed in the vertical direction is pulling force, which is equal in magnitude. In the first quadrant, the component is positive lift, and in the fourth quadrant, it is negative lift.

[0088] In summary, the static pressure value of a point A on the wing section line is:

[0089]

[0090] The plus sign in the formula is the upstream surface, and the minus sign is the downstream surface. The formula shows that when the upper wing surface is used, the first quadrant intercept should be used, and the intercept is a monotonically decreasing function curve without an inflection point; when the lower wing surface is used, the third quadrant intercept should be used, and the intercept is a monotonically decreasing function curve without an inflection point, so as to form a consistent double-sided positive lift airfoil, that is, a dynamic pressure difference airfoil.

[0091] For the convenience of discussion, we limit the lift-to-drag ratio of a point on the wing section line to the first and third quadrants, and consider it as the ratio of the absolute value of the lift increased by the dynamic pressure at that point to the absolute value of the drag increased by the dynamic pressure. Then the static pressure increment in formula (5) is multiplied by the static pressure of the smallest unit area at this point, and then projected in the vertical and horizontal directions respectively, and the ratio is calculated, which is the lift-to-drag ratio of the point:

[0092]

[0093] It means that the larger the angle, the higher the lift-to-drag ratio. Under the condition of satisfying lift, the angle should be increased as much as possible to improve the efficiency of the wing; the larger the angle, the larger the rear turning angle, the lower the static pressure on the surface, and the more it can avoid shock wave damage. The micro-segment of the cross section of the point can be considered as a straight line. If the straight line is extended only in the horizontal direction, the normal angle becomes larger, the tangent value increases, and the lift-to-drag ratio improves. If each point extends the wing cross section in the horizontal direction, it is an effective way to improve the lift-to-drag ratio. When adopting a horizontal straight uniform flow design, the chord of the wing is horizontal, and relatively non-horizontal, the normal angle of its cross section is large, the tangent value is large, and the lift-to-drag ratio is relatively high. The above is the reasoning and conclusion of the dynamic pressure difference lift principle in an incompressible fluid when the air flow is lower than 0.2 times the speed of sound.

[0094] Compressible airfoil section static pressure strength.

[0095] In the flow field, the parameter relationship between sections 1 and 2 of compressible, perfect gas, isentropic, one-dimensional high-speed steady flow is:

[0096]

[0097]

[0098] The static pressure value p on the left side of the plate 片左 When the airflow flows from the far left end to the left side of the sheet, it is an isentropic compression process. The far left end of the flow field is selected as section 1, and the parameters are flow field parameters; the left side of the sheet is section 2, and the velocity parameter is zero. Substituting into formula (6) yields:

[0099]

[0100]

[0101] When the airflow at point A of the tangent line has a normal angle, the static pressure intensity p α Substituting the normal velocity u cosα, we get:

[0102]

[0103] The static pressure value p on the right side of the plate 片右 When the gas on the right side of the sheet moves to the right along with the velocity of the flow field, it separates from the sheet and a vacuum space appears. Under the action of these two forces, the gas between the static pressure of the flow field and the vacuum fills the vacuum interval in an isentropic expansion process. When equilibrium is reached, formula (6) gives:

[0104]

[0105] When the airflow at point A of the tangent line has a normal angle, the static pressure intensity p α Substituting the normal velocity u cosα, we get:

[0106]

[0107] Formula (7) is used for the upstream surface, and formula (8) is used for the downstream surface.

[0108] The formula shows that the upper wing surface should use the first quadrant tangent, and the tangent height from the leading edge to the trailing edge decreases monotonically without an inflection point; the lower wing surface should use the third quadrant tangent, and the tangent height from the leading edge to the trailing edge decreases monotonically without an inflection point; only in this way can a consistent positive lift bihedral wing be formed at the same time, forming a dynamic pressure difference airfoil. The leading edge point of the dynamic pressure difference airfoil tangent is the highest point of the tangent, and the leading edge point of the centerline of the full wing is at the front tip of the nose; the highest leading edge point of the spanwise tangent is the front end of the nose, and the wing leading edge point and the spanwise leading edge point intersect at the origin of the coordinate system.

[0109] At low sound speeds, the above formula can complete the lift design and improvement of the cross-section.

[0110] However, at near-sonic and supersonic speeds, the impact of shock waves must also be considered. It can reduce the strong surface static pressure of the normal shock wave to the low-sonic surface static pressure by the angle of the normal line of the oncoming surface. All-wing aircraft are divided according to the maximum speed limit specified for low-sonic speeds, and the maximum limit has different values ​​according to different materials.

[0111] When the normal angle rotates from 0 degrees to α, the tangent also rotates α accordingly, which is called the tangent back rotation angle, or back rotation angle for short. The back means rotation along the velocity direction. The object plane back rotation angle is when the normal angle rotates from 0 degrees to α, the object plane also rotates α accordingly. Here it is called the object plane back rotation angle. The back rotation angle, object plane back rotation angle, and normal angle are equal. At near-sonic and supersonic speeds, increasing the back rotation angle will reduce the normal speed, work, and surface static pressure, and the resistance will also decrease. The tangent back rotation angle and the spanwise back rotation angle are used to weaken the point that exceeds the static pressure to a point that is equal to or less than the static pressure required for low sound speed.

[0112] The straight uniform air flow velocity u at point A is within the range of 5 times the supersonic speed, and the result calculated by the above formula is generally considered to be within an acceptable error range.

[0113] At near-sonic and supersonic speeds, shock waves will be generated on the windward surface, and the normal angle of the windward surface needs to be calculated so that the static pressure of the object surface reaches the low sound speed value we require. There is no shock wave on the leeward side, so there is no need to calculate it. From formula (7), the normal angle of the windward surface is obtained:

[0114]

[0115] We limit the range of α to 0 to 90 degrees. The larger the value, the smaller the normal velocity and the lower the static pressure on the surface. Based on practice, we often recommend the highest specified value of the low-sonic speed value. Using this value, we can calculate the highest specified value of the static pressure on the surface at that point. If it is equal to or less than this value, it is considered to meet the requirements of the static pressure on the surface at low sound speeds, eliminating the adverse effects of shock waves and not damaging the aircraft body. The spanwise line is a line perpendicular to the straight uniform flow velocity line. When it is near the sound speed and supersonic speed, it will also encounter the same strong static pressure generated by the positive shock wave as the wing section line. It is necessary to reduce the static pressure on the surface. Using the same derivation process of the wing section line angle, the same spanwise line angle formula (9) can be derived. Increasing the normal angle in the spanwise direction can reduce drag and increase flight speed.

[0116] We assume that the nose of the all-wing aircraft is facing left, the fuselage is placed horizontally (the fuselage is the wing, here it refers to the horizontal chord of the wing), the leading edge and the highest point of the nose are the origin of the coordinate system, the centerline of the horizontal chord of the aircraft is horizontally aligned with the axis X, the direction of the axis X is to the right, and parallel to the book; the span line is aligned with the axis Y, and the direction is toward us; the axis Z is vertical and upward, and the straight uniform flow is parallel to the axis X and in the same direction.

[0117] At near-sonic and supersonic speeds, the angle of the leading edge point of the lower wing section line is substituted into formula (7) with the highest low-sonic speed specified value and the cosine zero value. After calculating the highest low-sonic specified static pressure value, this static pressure value and the actual speed value of the point are substituted into formula (9) to calculate the α value. When the rear turning angle of each subsequent point is equal to or greater than this angle, the normal velocity and static pressure value of each point will not be higher than that of the leading edge point; if the rear turning angle of the subsequent point is equal to or greater than the rear turning angle of the previous point, no secondary shock wave will be generated. In this way, the static pressure value of each point on the wing section line does not exceed the highest specified value of the static pressure value of the low-sonic object surface.

[0118] Similarly, at near-sonic and supersonic speeds, the angle of the leading edge point of the spanwise line is substituted into formula (7) with the highest low-sonic speed specified value and the cosine zero value. After calculating the highest low-sonic specified static pressure value, this static pressure value and the actual speed value of the point are substituted into formula (9) to calculate the α value. When the rear turning angle of each subsequent point is equal to or greater than this angle, the normal velocity and static pressure value of each point will not be higher than that of the leading edge point; if the rear turning angle of the subsequent point is equal to or greater than the rear turning angle of the previous point, no secondary shock wave will be generated. In this way, the static pressure value of each point on the spanwise section line does not exceed the highest specified value of the static pressure value of the low-sonic surface. As a result, the static pressure value of the entire wing surface is lower than the highest specified value of the static pressure value of the low-sonic surface. The spanwise section view is a horizontal plane, and we only check the horizontal plane of the coordinate origin here, and do not check other horizontal planes. Because the leading edge points of the wing are all on the Y axis, the others can basically meet the requirements, otherwise they also need to be checked.

[0119] Different airfoils have different maximum speed limits for low-speed sound. Here we take a metal wing as an example. The static pressure value at the wing point is not higher than 0.95 (to 0.97, here we take 0.95) times the speed of sound u 0.95 The static pressure value p 0.95 , we can convert u 0.95 Substitute the cosine value equal to 1 into formula (7) to find the p of the point 0.95 ; Then the real speed of the leading edge point and p 0.95 Substitute into formula (9) to find the normal angle. When this angle is reached, the static pressure value of the point is equal to p 0.95 , when the angle is greater than this, the static pressure value is less than p 0.95 If applied to the wing section and span section of a full-wing aircraft, the entire wing surface will reach p 0.95 The requirements are met, achieving the purpose of weakening the shock wave to the static pressure strength of the low-speed sound object surface.

[0120] The straight uniform flow velocity of the flow field is less than u 0.95 The speed of sound is low, which is greater than or equal to u 0.95 The speeds are near the speed of sound and supersonic.

[0121] The above is the lift principle of aviation dynamic pressure difference all-wing aircraft, which is also the lift principle of aviation all-wing maneuvering pressure difference. It is applied below.

[0122] 4. Aerodynamic pressure difference full wing airfoil

[0123] In the design of the dynamic pressure difference airfoil (the chord length of the dynamic pressure difference airfoil in this paper refers to the distance between the projections of the leading and trailing edges on the chord line, and the chord line is a horizontal line), with a horizontal straight uniform flow and a chord elevation angle of zero degrees, according to formulas (7) and (8), the first quadrant tangent is selected for the upper wing surface, and the tangent is a monotonically decreasing function curve; the third quadrant tangent is selected for the lower wing surface, and the tangent is a monotonically decreasing function curve; the two are connected at the leading edge through a smooth inner arc at low sound speed to form a wing, and the trailing edge is directly connected to form a wing, forming a dynamic pressure difference wing; at near-sonic speed and supersonic speed, the upper wing and the lower wing are directly connected at the leading edge point; the angle at the leading edge point of the lower wing surface tangent line is specified by the maximum low sound speed u. 0.95 Substitute the zero-degree value of cosine into formula (7) to calculate the maximum low-speed static pressure value p 0.95 Then, substitute this static pressure value and the straight uniform flow velocity value into formula (9) to calculate the α value. When the rear turning angle of each subsequent point is equal to or greater than this angle, the static pressure value of each point will not be higher than that of the leading edge point; if the rear turning angle of the subsequent point is equal to or greater than the rear turning angle of the previous point, no secondary shock wave will be generated. In this way, the static pressure value of each point on the wing section line does not exceed the maximum specified value of the static pressure value of the low-sonic surface, avoiding the influence of shock waves. As a result, both the upper and lower wing surfaces will generate positive lift at the same time, and then according to the effective method of point lift-to-drag ratio, a more efficient section line is selected to further improve the lift-to-drag ratio.

[0124] The current wing based on velocity difference lift principle is changed to a wing based on dynamic pressure difference

[0125] Assume that the airfoil of a full-wing aircraft is designed for a horizontal straight uniform flow field, the upper wing surface is an upward arc, point A is the leading edge, point B is the trailing edge, the lower wing surface is horizontal, AB is the chord length, C is the highest point, D is the foot of C, and the distance between point D and point A is 25% of the length of AB. If the straight uniform airflow flows horizontally from left to right, curve AC is the upstream surface and curve CB is the downstream surface. Therefore, the upstream surface of the upper wing curve ACB should be removed, leaving the downstream surface. The schematic diagram of the cross-section of the full-wing aircraft airfoil is as follows: Figure 5 shown.

[0126] Take the left part ACD and make a symmetrical image with the horizontal line as the symmetry axis as part ADE. The lift of the two curves in the symmetrical image is equal in magnitude but opposite in direction; the horizontal drag is equal. Figure 6 shown.

[0127] Draw a graph of the ADE part, the length of AD is 25% of AB, extend the AD horizontal line twice, get the AF horizontal line, AF is 75% of AB length, through point F, draw a vertical line, take point G, and make GF equal to the length of DE. Make the AG curve of the AFG part (shown by the dotted line) have the same curve type as the AE curve. Due to the horizontal extension, the normal angle of the AG curve is larger, the tanα value is higher, and the lift-to-drag ratio is higher. The transition diagram after the horizontal length is extended by twice is as follows Figure 7 shown.

[0128] Will Figure 5 The CB curve is used as the upper wing surface, and Figure 7 The AG curve is used as the lower wing surface, that is, point C is connected to point A, point B is connected to point G, the horizontal lengths of the two are equal, and the vertical heights are equal, forming a new dynamic pressure difference airfoil, where the solid line represents the upper wing surface, and the dotted line represents the lower wing surface. The schematic diagram of the new airfoil with a length of 75% is shown in the figure. Figure 8 shown.

[0129] The new dynamic pressure difference airfoil generates positive lift on both the upper and lower surfaces, forming a double-surface lift wing with an additional lift surface, which significantly improves the lift-to-drag ratio.

[0130] Increasing the normal angle can further improve the lift-to-drag ratio of the airfoil.

[0131] We can further improve the tangent value to optimize the wing efficiency. Figure 8 The leading edge point C (A) in the figure is horizontally extended to the left by 25% of the length of AB, and the upper wing section line CB (solid line) and the lower wing section line AG (dotted line) are extended according to their similar line types. The extended CB (solid line) and AG (dotted line) form a smoother and more efficient new airfoil, further improving the lift-to-drag ratio of the airfoil. This optimized dynamic pressure difference airfoil is the aerodynamic pressure difference full-wing airfoil. The schematic diagram after optimization and improvement is as follows: Fig. 9 shown.

[0132] 5. Aerodynamic pressure differential all-wing aircraft

[0133] At the origin of the full-wing aircraft coordinates, on the span line, marks are placed every unit length on both sides, and a cross-sectional view of the full-wing aircraft with aerodynamic pressure difference is made at each mark. Then, the leading edge point is placed at the corresponding mark, the chord is kept horizontal, and the figure is vertical, and the cross-sectional view of the full-wing aircraft with low-sonic aerodynamic pressure difference is made. At low sound speeds, the leading edge points on both sides of the span line are connected by a smooth inner arc.

[0134] However, at near-sonic and supersonic speeds, the angle between the normal line at the leading edge of the lower wing section line is the maximum subsonic speed specified value u 0.95 Substitute the zero-degree value of cosine into formula (7) to calculate the maximum low-speed static pressure value p 0.95Then, substitute this static pressure value and the straight uniform flow velocity value into formula (9) to calculate the α value. When the rear turning angle of each subsequent point is equal to or greater than this angle, the static pressure value of each point will not be higher than that of the leading edge point; if the rear turning angle of the subsequent point is equal to or greater than the rear turning angle of the previous point, no secondary shock wave will be generated, and this is considered to be the completion of the production of one wing profile; for every unit length on both sides of the span, a profile is produced until all profiles are completed. The angle between the spanwise line and the leading edge point normal of the coordinate origin is the specified value u of the highest subsonic speed. 0.95 Substitute the zero-degree value of cosine into formula (7) to calculate the maximum low-speed static pressure value p 0.95 Then, substitute this static pressure value and the straight uniform flow velocity value into formula (9) to calculate the α value. When the rear turning angle of each subsequent point is equal to or greater than this angle, the static pressure value of each point will not be higher than that of the leading edge point; if the rear turning angle of the subsequent point is equal to or greater than the rear turning angle of the previous point, no secondary shock wave will be generated. The two sides of the spanwise line are directly connected at the leading edge point. After the spanwise line is completed, the leading edge points of the wing section diagram are matched one by one to the marked positions on the spanwise line, the chord is horizontal, and the diagram is vertical. At this point, the unfolded plane diagram of the aerofoil of the aviation full wing is completed.

[0135] If you want to complete the three-dimensional diagram, you can use a smooth curved surface to connect the wing sections of each section to form a three-dimensional diagram of the full-wing aircraft with aviation dynamic pressure difference.

Claims

1. An aerodynamic pressure differential all-wing aircraft, the cross-sectional development diagram of which includes an upper wing surface and a lower wing surface, characterized in that: The upper wing surface and the lower wing surface section lines are both curved surfaces consisting of monotonically descending curves from the leading edge to the trailing edge.

2. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: The main wing surface of the upper wing is composed of a monotonic descending function surface from the leading edge to the trailing edge, and generally has no inflection point.

3. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: The main wing surface of the lower wing is composed of a monotonic descending function surface from the leading edge to the trailing edge, and generally has no inflection point.

4. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: At near-sonic and supersonic speeds, the leading edge of the upper wing and the leading edge of the lower wing directly intersect.

5. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: At near-sonic and supersonic speeds, the rearward turning angle of the leading edge point of the lower wing section line is equal to or greater than the required normal angle of the leading edge point.

6. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: At near-sonic and supersonic speeds, the section line of the lower wing surface is the rearward turning angle from the leading edge to the trailing edge, and the rearward turning angle of the latter point is equal to or greater than the rearward turning angle of the former point.

7. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: At near-sonic and supersonic speeds, the rearward turning angle of the leading edge point of the spanwise line is equal to or greater than the required normal angle of the point.

8. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: The spanwise section line is a rear angle from the leading edge to the trailing edge such that the rear angle at the rear point is equal to or greater than the rear angle at the front point.

9. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: Increasing the normal angle of the wing section can improve the lift-to-drag ratio.

10. The aerodynamic pressure differential all-wing aircraft according to claim 1, characterized in that: Increasing the normal angle of the spanwise section can reduce drag.

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