Flight method for short-distance take-off and landing by means of angle-of-attack lift force

Abstract

The invention provides a short-distance take-off and landing flight method depending on attack angle lift force, relates to the technical field of aircrafts, points out the essence that the attack angle lift force comes from air acceleration, and provides a calculation method and characteristic indexes of short-distance take-off and landing flight depending on the attack angle lift force. The flight principle, the trimming mechanism and the control method of the flat-plate-wing aircraft are systematically elaborated, the dynamic mechanical analysis process is simplified, the aerodynamic layout design key points of the A-wing aircraft are summarized, the flight theory of short-distance take-off and landing depending on the attack angle lift force is provided, and accordingly the two-dimensional maneuvering characteristics and the market application prospect of the A-wing aircraft are analyzed. The direction is pointed out for aerodynamic layout planning and preliminary design calculation of the A-wing aircraft; according to the method provided by the invention, theoretical analysis, configuration design, parameter calculation, take-off and landing modes and flight control of the A-wing aircraft are covered, and a complete technical scheme is formed.

Description

Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to a flight method that relies on angle-of-attack lift for short takeoff and landing. Background Technology

[0002] Fixed-wing aircraft, also known as fixed-wing aircraft, primarily generate lift from airfoil lift. Airfoil lift relies on a certain airflow speed to produce sufficient lift, requiring the aircraft to take off and land on a runway. This makes them heavily dependent on runways and limits their applicability.

[0003] Vertical takeoff and landing aircraft, also known as rotorcraft, use an upward-moving propeller to pull the aircraft vertically for takeoff and landing, such as helicopters and multi-rotors. Rotary aircraft are convenient and fast, but they are complex in structure, consume a lot of energy, produce significant noise and vibration, have poor reliability, and their flight efficiency is far lower than that of fixed-wing aircraft.

[0004] Therefore, there is an urgent need for a new solution to address the above problems, enabling aircraft to break free from runway dependence and overcome the drawbacks of existing vertical takeoff and landing aircraft. Summary of the Invention

[0005] The purpose of this invention is to provide a new type of aircraft that can both get rid of runway dependence and overcome the shortcomings of existing vertical take-off and landing aircraft. It utilizes a flight method that relies on angle of attack lift for short take-off and landing, thus forming an aerodynamic layout scheme for a winged aircraft.

[0006] This invention relates to a fixed-wing aircraft that flies using angle-of-attack lift. Unlike fixed-wing aircraft that fly using airfoil lift, which require a certain speed for takeoff and landing, angle-of-attack lift is related to acceleration, reducing the dependence on speed. Therefore, this invention enables a fixed-wing aircraft to reduce its dependence on runways during takeoff and landing, and can achieve short takeoff and landing with a takeoff and landing distance of 0.

[0007] The present invention relates to an aircraft that relies on angle-of-attack lift for short takeoff and landing. This aircraft has a low wing loading and a wide chord wing with a wing surface load of no more than 30 kg per square meter and an aspect ratio of no more than 3.5. When the thrust-to-weight ratio is no less than 1.2 and the angle of attack is no less than 30°, more than half of the takeoff weight is borne by the vertical component of the thrust. At the same time, the angle-of-attack lift generated by the large-area wing during acceleration lifts part of the takeoff weight to achieve short takeoff and landing.

[0008] Let the takeoff weight of the A-wing aircraft be mg, and the wing area be S. When the thrust-to-weight ratio is equal to 1.2 and the angle of attack is equal to 30°,

[0009] The vertical component of the thrust, Ty = T * sinθ = 1.2mg * sin30° = 0.6mg.

[0010] The horizontal component of the thrust, Tx, = T * cosθ = 1.2mg * cos30° = 1.04mg.

[0011] When a winged aircraft performs short takeoff and landing, there is an acceleration a in the vector direction of the thrust T, i.e., T=ma. Given T=1.2mg, a=1.2g. Therefore, ax=a*cosθ=1.2g*cos30°=1.04g. When the wing moves forward with acceleration ax, the air in front of the wing generates inertial drag Gx. The volume of the air in front of the wing is Dk*S*sinθ, where Dk is the shape drag coefficient of the wing. The Dk coefficient is 0.7 for a circular wing and 0.6 for a square wing.

[0012] When the aspect ratio is less than 3.5, the shape of the wing is approximately square. Taking Dk = 0.6, θ = 30°, and air density ρ = 1.2, the inertial drag Gx is:

[0013] Gx = ax*ρ*Dk* S*sinθ ;

[0014] We can calculate that Gx = 1.04g * 1.2 * 0.6 * S * sin30° = 0.37gS;

[0015] The inertial drag Gx creates wind pressure Wx under the wing surface, Wx = Gx / S / sin30° = 0.74g. This wind pressure Wx is transmitted vertically to the wing surface, creating lift F at the angle of attack, i.e.: F = Wx * S = 0.74gS.

[0016] The lift force F at the angle of attack can be decomposed into a horizontal component Fx and a vertical component Fy at the angle of attack θ, as follows:

[0017] Fx = F * sinθ = 0.74gS * sin30° = 0.37gS,

[0018] Fy = F * cosθ= 0.74gS* cos30°=0.64gS,

[0019] Fy and Ty have the same direction, both vertically upwards, therefore:

[0020] Fy+Ty=0.64gS+0.6mg=0.64*9.8*S+0.6mg=6.3S+0.6mg;

[0021] If the wing surface load is 30 kg per square meter, i.e., mg / S = 30, then S = mg / 30, which means:

[0022] Fy+Ty=6.3S+0.6mg = 6.3* mg / 30+0.6mg = 0.21mg+0.6mg=0.81mg,

[0023] At this point, the thrust is basically enough to overcome the takeoff weight (mg), and a short runway is sufficient for takeoff.

[0024] The takeoff distance for short takeoff is calculated by accelerating from 0 to Vbe with acceleration ax, therefore the takeoff time t is:

[0025] ,

[0026] Where Vbe is the sustained flight speed of the armored aircraft, ax is the horizontal component of acceleration, mg is the takeoff weight, and S is the area of ​​the armored wing.

[0027] The distance s required for the gliding run is:

[0028] ,

[0029] Where t is the time required for takeoff and landing, Vbe is the sustained flight speed of the wingless aircraft, mg is the takeoff weight, and S is the area of ​​the wingless aircraft.

[0030] Since the vertical component of the thrust, Ty, bears 0.6 mg (18 kg) of the takeoff weight, and the takeoff weight requiring lift at the angle of attack is 12 kg, the takeoff run time t is: t = 0.13 * =1.41 seconds; the distance s required for takeoff is: s=0.18*12*9.8=21 meters. For a heavy-load aircraft with a wing surface load of 30 kg per square meter, the fuselage length of the A-wing aircraft is greater than 6 meters, so it only needs to take off from 4 parking positions.

[0031] If the wing surface load mg / S is 20 kg per square meter, and the thrust-to-weight ratio is 1.2 with an angle of attack greater than 30°, since the vertical component of the thrust Ty bears 0.6 mg (12 kg) of the takeoff weight, the takeoff weight that needs to be lifted by the angle of attack lift is 8 kg. Therefore, the takeoff time t required is: t = 0.13 * =1.15 seconds; the distance s required for takeoff is: s=0.18*8*9.8=14 meters. For a heavy-load aircraft with a wing surface load of 20 kg per square meter, the fuselage length of the A-wing aircraft is greater than 5 meters, so it only needs to take off from 3 parking positions.

[0032] If the wing load is 20 kg per square meter, short takeoff and landing with a takeoff distance of 0 can be achieved when the thrust-to-weight ratio is 1.2 and the pitch angle is greater than 60°, or when the thrust-to-weight ratio is greater than 2 and the pitch angle is greater than 30°, or when the thrust-to-weight ratio is greater than 1.5 and the pitch angle is greater than 45°.

[0033] If the wing load is 10 kg per square meter, when the thrust-to-weight ratio is greater than 1.2 and the pitch angle is greater than 30°, Fy + Ty = 1.23 mg. At this time, the lift is greater than the takeoff weight, and the aircraft can take off immediately with a takeoff distance of 0.

[0034] It is evident that the smaller the wing loading, the shorter the takeoff distance. When the angle of attack remains constant, the larger the thrust-to-weight ratio, the shorter the takeoff distance, and the smaller the thrust-to-weight ratio, the longer the takeoff distance. When the thrust-to-weight ratio remains constant, the larger the angle of attack, the shorter the takeoff distance, and the smaller the angle of attack, the longer the takeoff distance.

[0035] The aircraft's wing is flat, without an airfoil, and does not generate airfoil lift; it is called wing A. The lift of wing A comes from angle-of-attack lift. The angle-of-attack lift comes from the inertial drag of the air in front of the wing and is related to acceleration. Due to the elastic collision between the lower wing surface and the air quanta, the inertial drag of the air quanta is transmitted vertically to the wing surface, thereby generating lift. The lift is closely related to the wing's angle of attack, hence it is called angle-of-attack lift.

[0036] When an airfoil accelerates from 0 to 0 at an angle of attack θ in still air, the initial velocity of the oncoming airflow is 0. The acceleration of the airfoil forces the airflow to wash downwards, causing the downwash airflow to accelerate. When an airfoil moves horizontally at a constant velocity V in still air at an angle of attack θ, the air washes downwards at a constant velocity V under the action of the airfoil surface, also causing the downwash airflow to accelerate. The reaction force generated by the acceleration of the downwash airflow is the lift F of the airfoil at the angle of attack.

[0037] Airfoil lift is related to speed. Fixed-wing aircraft rely on airfoil lift to fly, and airfoil lift depends on speed. The greater the speed, the greater the airfoil lift. Therefore, fixed-wing aircraft rely on runway takeoff and landing to increase lift for takeoff and landing. Angle of attack lift is related to acceleration. The greater the acceleration, the greater the angle of attack lift. Therefore, the aircraft can obtain angle of attack lift by generating acceleration through burst thrust when the speed is 0, so as to achieve ultra-short takeoff and landing, or even takeoff and landing on the spot with a runway distance of 0, i.e., hop takeoff.

[0038] The low wing loading and large-area wing of the mech-wing aircraft create conditions for short takeoff and landing. The large-area wing can generate lift at the angle of attack to help the mech-wing aircraft take off during takeoff, and can also generate drag to help brake the mech-wing aircraft during landing. The takeoff distance is directly proportional to the wing surface load mg / S, inversely proportional to the angle of attack θ, and inversely proportional to the thrust T. When the wing surface load mg / S is not greater than 10 kg / m², the thrust-to-weight ratio is not less than 1.2, and the angle of attack θ is not less than 30°, the duration of short takeoff of the mech-wing aircraft is 0, the takeoff distance is 0, and it immediately switches to fixed-wing flight after takeoff.

[0039] When the drive mechanism uses a propeller, the preferred installation position is located at the leading edge of the wing, and it is set as a forward-pull layout; the preferred landing gear layout is a tricycle landing gear, the preferred takeoff and landing angle of attack is 30°, the nose landing gear is the main landing gear, located below the cabin, the propeller is located above the cabin, increasing the propeller height to avoid ground scraping, the rear landing gear is the auxiliary landing gear, located on both sides of the trailing edge of the wing, improving takeoff and landing stability, and the load weight in the cabin is moved forward as much as possible for balance, improving flight stability.

[0040] The aircraft with the aforementioned wing structure is called a wingless aircraft, which is a type of flying wing aircraft. The wingless aircraft has a two-dimensional maneuverability, namely forward maneuver and vertical maneuver, which is between the three-dimensional maneuverability of a rotorcraft (i.e., forward maneuver, vertical maneuver, and horizontal maneuver) and the one-dimensional maneuverability of a fixed-wing aircraft, namely forward maneuver, as shown in the table below.

[0041]

[0042] The so-called axle-wing aircraft, as a cross-border aircraft between one-dimensional and three-dimensional maneuvering aircraft, has a flight state between that of a fixed-wing aircraft, a transitional flight state, and a rotorcraft. Its advantages and disadvantages also lie between those of rotorcraft and fixed-wing aircraft. For mass-market aircraft like flying cars that need to fly at low altitudes, low speeds, high frequencies, and short distances in densely populated areas, the axle-wing aircraft has significant advantages. On the one hand, the axle-wing aircraft's flight efficiency, speed, altitude, and range are inferior to those of fixed-wing aircraft, so it is only suitable for short- to medium-distance flights. On the other hand, the axle-wing aircraft's vertical takeoff and landing performance and three-dimensional maneuverability are inferior to those of rotorcraft, but its safety, economy, reliability, and comfort are all superior. Therefore, rotorcraft are more geared towards military use, while axle-wing aircraft are suitable for civilian use. For mass-market aircraft that fly at low altitudes and short distances, the axle-wing aircraft's advantages are significant while its disadvantages are acceptable, making it the preferred configuration for flying cars.

[0043] The thickness of the wing is less than one-twentieth of the chord length, the aspect ratio is less than 3.5, and the basic airfoil is a thin-shell flat plate without curvature. Simplifying the large-area shell-shaped wide-chord wing into a flat plate greatly reduces production difficulty and maintenance costs. The preferred shape of the wing is rectangular, rhomboid, trapezoidal, triangular, elliptical, etc. The wing surface load of the aircraft is less than 30 kg per square meter, and the preferred wing surface load is 3 kg to 10 kg per square meter. It flies by relying on the angle-of-attack lift generated by the large-area wing.

[0044] Let the area of ​​the wing be S, the thrust be T, the takeoff mass be m, the takeoff weight be mg, and the angle of attack be θ.

[0045] When a winged aircraft performs short takeoff and landing, there is an acceleration *a* in the vector direction of the thrust *T*, i.e., *T* = *ma*. The horizontal component of acceleration *a* is *ax*, so *ax* = *a*cosθ. The horizontal component *ax* of acceleration *a* is generated by the horizontal component *Tx* of the thrust *T*, giving the winged aircraft a tendency to move horizontally, i.e., *Tx* = *m*ax*. Since the air in front is stationary, the horizontal movement of the winged aircraft will be hindered by the inertia of the stationary air, generating inertial drag *Gx*. The volume of the inertial drag air is equal to the frontal area of ​​the wing multiplied by the shape drag coefficient *Dk* of the wing. For a circular wing, *Dk* is 0.7, and for a square wing, *Dk* is 0.6. Therefore, the volume of the inertial drag air is *Dk* *S*sinθ, and the mass of the inertial drag air is *ρ* *Dk* *S*sinθ. Therefore, the inertial drag *Gx* is:

[0046] Gx = ax*ρ*Dk* S*sinθ ;

[0047] Due to the quantum properties and elastic collision characteristics of air, the inertial drag Gx is transmitted vertically to the wing surface, forming the angle-of-attack lift F, that is: F = Gx = ax*ρ*Dk* S*sinθ;

[0048] The vertical component of the thrust T is: Ty = T * sinθ. When Ty = mg, the takeoff distance of the wingless aircraft is 0, achieving a jump. That is, the condition for the wingless aircraft to jump is: sinθ ≥ mg / T. Even without considering the ground effect and the lift F generated by the inertial drag Gx, the wingless aircraft can jump using only the vertical component of the thrust T. If T = 2mg, the wingless aircraft jumps when θ ≥ 30°; if T = 1.5mg, the wingless aircraft jumps when θ ≥ 45°; if T = 1.2mg, the wingless aircraft jumps when θ ≥ 60°; if T = 1mg, the wingless aircraft jumps when θ = 90°.

[0049] When the A-wing aircraft is flying at level speed, its speed is V. The stationary air in front of it is washed down at a constant speed of V under the action of the wing surface, which also causes the air under the wing to have acceleration a and angle of attack lift F.

[0050] If, within time t, oncoming air with velocity V is washed down at a constant velocity V by the wing surface, then the downwash acceleration a = V / t; let the mass of the downwash air be Mk, then within time t seconds, we have:

[0051] Mk = *V*t*S*sinθ, Let the air density be denoted as . = 1.2 kg / m3;

[0052] That is, the lift force F at the angle of attack is: F = Mk*a = V / t * *V*t*S*sinθ= 1.2 *V 2 *S*sinθ;

[0053] It is evident that the lift F at the angle of attack is directly proportional to the square of the velocity V and the sine of the angle of attack θ. When the velocity V remains constant, increasing the angle of attack θ increases the lift; when the angle of attack θ remains constant, increasing the velocity V increases the lift. When the angle of attack θ = 0, the lift F at the angle of attack is zero regardless of the velocity V.

[0054] When a mech-wing aircraft is in level flight, the oncoming airflow creates wind pressure that acts on the lower surface of the wing. This fluid pressure forces the air downwash at an acceleration *a*, generating lift at the angle of attack. The faster the flight speed *V*, the greater the wind pressure, the greater the downwash acceleration *a*, and the greater the lift at the angle of attack *F*. Once the lift at the angle of attack reaches the takeoff weight (F=mg), it will not increase further. Continuing to increase the flight speed *V* only raises the mech-wing aircraft, reducing the angle of attack; the faster the speed *V*, the smaller the angle of attack *θ*.

[0055] To simplify the aerodynamic analysis of the armored wing aircraft, the wing is simplified to a square shape, and the drive system is simplified to a single forward-pushing propeller, with the propeller exhaust airflow divided by the wing. Let the chord length of the wing be c, the wingspan be l, the propeller diameter be D, the area of ​​the wing be S, and the horizontal frontal area (minimum drag area) of the armored wing aircraft be Sd. Therefore, the aspect ratio λ is: λ = l / c, and the area S is: S = c * l. Then:

[0056] The lift-to-drag ratio K of the winged aircraft is: K = S / Sd;

[0057] The thin-shell flat-plate wing is very thin, less than one-twentieth of the chord length. This means that the frontal area during horizontal flight is less than 0.05 times the wing area S. Even after adding the drag area of ​​the fuselage, the minimum drag area Sd of the wing is still much less than 0.1 times the wing area S. Therefore, the lift-to-drag ratio K = S / Sd > 10. By reducing the wing thickness and reducing the drag area of ​​the fuselage through drag reduction design, the lift-to-drag ratio K of the wing can reach more than 30.

[0058] When a mech-winged aircraft flies horizontally with a thrust T and an angle of attack θ, the oncoming wind pressure W is: W = T / Sd. This oncoming wind pressure compresses the air under the wing of the mech-winged aircraft. The pressure of the compressed gas is equal in all directions, thus generating lift F in the normal direction of the wing, F = S*W, that is: F = S*W = S*T / Sd = T*S / Sd = T*K. Therefore, the mech-winged aircraft obtains an angle-of-attack lift F that is K times T with a thrust of T. The mech-winged aircraft flies precisely by utilizing this angle-of-attack lift mechanism.

[0059] The surface loading of the wingless aircraft, mg / S, is the ratio of its takeoff weight, mg, to its wing area, S. When the wingless aircraft maintains a constant altitude and speed with thrust T and an angle of attack θ, the thrust T is δ times its takeoff weight, mg, i.e., T = δmg, where δ is the thrust-to-attack ratio of the wingless aircraft.

[0060] ;

[0061] When the winged aircraft flies horizontally at a constant angle of attack θ and velocity V, the propeller generates thrust T along the wing chord direction, which can be decomposed into horizontal thrust Tx and vertical thrust Ty under the action of the angle of attack θ.

[0062] Tx = T * cosθ , Ty = T * sinθ ,

[0063] The vertical thrust Ty offsets part of the weight of the winged aircraft, while the horizontal thrust Tx drives the winged aircraft to move horizontally, generating a frontal wind pressure W. When the winged aircraft flies horizontally at a constant speed of θ and V, the frontal wind pressure W is:

[0064] .

[0065] The oncoming wind pressure W acts on the lower surface of the straight airfoil of the A-wing aircraft, generating fluid pressure in the normal direction of the A-wing. This pressure is the lift force F at the angle of attack. Therefore:

[0066] ;

[0067] The lift force F at the angle of attack can be decomposed into a horizontal component Fx and a vertical component Fy at the angle of attack θ, where Fx = F * sinθ and Fy = F * cosθ, i.e.:

[0068] ;

[0069] When the angle of attack θ > 10°, sin θ = 0.17, which is much larger than 1 / K = Sd / S = 0.03. Therefore, the minimum drag area Sd can be ignored, and the calculation formula can be simplified to:

[0070] ,

[0071] Therefore: That is, Fx = Tx. It can be seen that the horizontal component Fx of the lift F at the angle of attack of the A-wing aircraft is the wind resistance. The wind resistance Fx and the horizontal thrust Tx cancel each other out, are equal in magnitude and opposite in direction.

[0072] The vertical component Fy of the angle-of-attack lift F of the winged aircraft is:

[0073] ,

[0074] If the propeller generates a thrust T along the chord of the wing, pulling the A-wing aircraft to fly horizontally at a constant speed of θ and V, that is, the lift of the A-wing aircraft equals its weight, then:

[0075] Fy + Ty = mg

[0076] Substituting the formulas for Fy and Ty into the above formula, we get: ,

[0077] Substituting the thrust-to-attack ratio δ = T / mg into the equation, we have:

[0078] ;

[0079] The thrust-to-attack ratio δ is a dimensionless parameter representing the ratio of the thrust T required by the drive system to the takeoff weight mg for a winged aircraft to maintain a constant altitude and speed at an angle of attack θ. For θ < 10°, δ < 0.1; for θ > 60°, δ > 0.9, it exhibits a hyperbola with relatively steep extremes. See the table below:

[0080]

[0081] It can be seen that the thrust-to-attack ratio δ decreases rapidly as the angle of attack θ decreases, which means that the thrust T of the drive device required to maintain flight altitude decreases rapidly. For an armored aircraft to maintain flight altitude at the wing angle of attack θ, the thrust T of the drive device needs to satisfy: T = δmg.

[0082] Let Tl be the level flight thrust of the A-wing aircraft, and θl be the level flight angle of attack of the A-wing aircraft. When the thrust T=Tl remains constant, θ=θl means the A-wing aircraft is flying at level; θ>θl means the A-wing aircraft is descending, with a decrease in speed and altitude; θ<θl means the A-wing aircraft is climbing, with an increase in speed and altitude. When the angle of attack θ=θl remains constant, T=Tl means the A-wing aircraft is flying at level; T<Tl means the A-wing aircraft is descending, with a decrease in speed and altitude; T>Tl means the A-wing aircraft is climbing, with an increase in speed and altitude.

[0083] When the thrust-to-tank ratio of the A-wing aircraft is δ = 1 / K, the thrust required to maintain the aircraft's flight altitude is the minimum level flight thrust Tmin: Tmin = δmg = mg / K. Even if the angle of attack θ is further reduced to 0 ≦ δ ≦ 1 / K, the drag area of ​​the A-wing aircraft has already reached the minimum drag area Sd. Therefore, even if the angle of attack θ is further reduced, the drag area will not decrease further, and thus the thrust will not decrease further. Therefore, when designing an A-wing aircraft, increasing the wing area S or reducing the minimum drag area Sd can increase the lift-to-drag ratio K, thereby improving the flight efficiency of the A-wing aircraft. From the perspective of the aerodynamic layout of the A-wing aircraft, optimizing K to 30 or higher is feasible.

[0084] The minimum level flight thrust Tmin corresponds to the constant horizontal flight speed, which is the minimum horizontal flight speed of the armored wing aircraft, also known as the long-range flight speed Vbe. The wind pressure corresponding to the long-range flight speed is equal to the wing surface load, i.e.:

[0085] ,

[0086] C is the drag coefficient. For a rectangular wing with an angle of attack, c = 0.3. However, for an A-wing aircraft, the drag from the fuselage and frame needs to be added, so c = 0.9 is used. Let the air density be denoted as . = 1.2 kg / m 3 ,have to:

[0087] ,

[0088] The wind pressure corresponding to the long-range speed Vbe is equal to the wing surface load, and the angle of attack corresponding to the long-range speed Vbe is the long-range angle of attack θbe.

[0089] When 90°>θ>θbe, the winged aircraft is in a transitional flight state. The horizontal component of the flight speed V, Vl, is less than the sustained flight speed Vbe, that is: Vl<Vbe. Increasing the thrust T can only increase the flight altitude, but cannot increase the horizontal flight speed Vl, and cannot perform sprint flight. The larger θ is, the smaller Vl is.

[0090] When θ=90°, Vl=0, and the A-wing aircraft is in rotorcraft flight mode.

[0091] When θ≦θbe, the A-wing aircraft is in fixed-wing flight mode, and its flight speed is proportional to its thrust T. If δ<1 / K, the A-wing aircraft stalls at a speed equal to its sustained flight speed Vbe. The A-wing aircraft continuously descends to maintain a horizontal speed equal to Vbe. Let β be the thrust-to-weight ratio of the A-wing aircraft during its sprint flight, and Tmcp be the thrust corresponding to the maximum sustained output power of the A-wing aircraft. Then we have: Tmcp=βmg, and β>δ. If β≧1 / K, the A-wing aircraft enters sprint flight, and increasing the thrust at this time will increase its horizontal flight speed.

[0092] Let Wbe be the long-range wind pressure of the armor-wing aircraft. The long-range wind pressure is equal to the wing surface loading:

[0093] Wbe = Tmin / (S / K) = mg / S;

[0094] Let the sprint wind pressure of the winged aircraft be Wmax, and the sprint thrust of the winged aircraft be β times its takeoff weight. Since Tmcp = βmg, then the sprint wind pressure Wmax = Tmcp / (S / K) = βKmg / S.

[0095] According to the general formula for wind speed and wind pressure:

[0096] ,

[0097] The maximum horizontal sprint speed Vmax can be calculated as follows:

[0098] ;

[0099] That is, the maximum horizontal sprint speed Vmax is equal to the sustained flight speed Vbe. times, that is: , where β is the multiple of the maximum sprint thrust Tmcp to the takeoff weight mg, i.e.: Tmcp=βmg, and β>δ.

[0100] Since the long-range speed Vbe is determined by the wing loading, the sprint speed Vmax of the armored wing aircraft is determined by the sprint thrust Tmcp, lift-to-drag ratio K, and wing loading mg / S, and the sprint speed is proportional to all three.

[0101] When the A-wing aircraft is in rotorcraft flight mode and transitional flight mode, its uniform horizontal flight speed is less than or equal to the long-range flight speed Vbe. Increasing thrust can only increase the climbing speed and cannot exceed the long-range flight speed Vbe. Only when the A-wing aircraft's angle of attack decreases to θ≦θbe, allowing it to enter fixed-wing aircraft flight mode, can increasing thrust increase the horizontal flight speed and achieve sprint flight. The sprint speed is proportional to the thrust T.

[0102] The typical wing loading of a winged aircraft is 10 kg / m. 2 Assuming an air drag coefficient of c = 0.9, the corresponding stall speed is 48.6 kilometers per hour.

[0103] ,

[0104] Considering the economy of the propulsion system and the comfort of hop-off flight for manned aircraft, the multiple of sprint thrust to takeoff weight is taken as β≦1.2, and the lift-to-drag ratio is taken as K=15. Calculations show that the maximum flight speed of the armor-wing aircraft is approximately 206 kilometers per hour.

[0105] ,

[0106] Trial is a crucial prerequisite for aircraft flight. The flight trim of a winged aircraft is determined by the propeller position, the overall center of gravity, and the aerodynamic focus of the wing. The propeller wake has a significant impact on the stress on the wing surface and must be considered during trim. The propeller position is the main variable affecting the pressure generated by the propeller wake on the upper wing surface.

[0107] If the center of gravity of the aircraft is set in front of and below the aerodynamic focus of the wing, and the propeller axis coincides with the wing surface, then the line connecting the center of gravity of the aircraft and the aerodynamic focus is the horizontal line for the wing aircraft to fly at level. The closer the center of gravity of the aircraft is to the wing surface, the smaller the angle of attack during level flight.

[0108] If the propeller is positioned above the leading edge of the wing with an eccentricity e1, the center of gravity CG of the entire aircraft is located below the wing at a distance e2 from the wing and at a distance d1 from the propeller disk, the aerodynamic focus AC of the wing is located on the wing surface at a distance d2 from the center of gravity CG, F is the lift at the angle of attack of the wing, Fh is the lift of the leading edge strake of the wing, and Ft is the pressure generated by the propeller wake on the upper surface of the wing. The magnitude of the wake pressure Ft is related to the propeller thrust T, the eccentricity e1, and the wing chord length. The positional change of the wake pressure Ft is related to the tilt angle of the aileron. When the above installation position, the wing chord length b, the propeller radius R, and the propeller thrust T are all determined, the propeller wake pressure Ft is:

[0109] , where Ω is the tail pressure coefficient;

[0110] Let the eccentricity coefficient be ψ, and e1 = ψR, then the tail pressure coefficient Ω is:

[0111] ;

[0112] When e1=0, ψ=0, the pressure formed by the propeller wake above and below the wing is equal, Ω=0, therefore the pressure Ft formed on the wing surface is 0; when 0<E1<R, the pressure formed by the propeller wake above and below the wing is not equal, therefore a pressure Ft is formed on the wing surface, which is equal to the propeller wake pressure multiplied by the area where the propeller wake is tangent to the wing; when E1>R, ψ=0, the propeller wake pressure is equal to the disk load, but the area where the propeller wake is tangent to the wing is 0, therefore the pressure Ft formed on the wing surface is 0. Choosing ψ=0.0, 0.1~0.9, 1.0, the tail pressure coefficient Ω, which generates downward pressure at the aerodynamic focus of the propeller wake on the wing surface, can be calculated as:

[0113]

[0114] It can be seen that when ψ > 0.5, Ω > 0.5; when ψ = 0.7, Ω is at its maximum, Ωmax = 0.58. When b = 4R and ψ = 0.2, Ft = 0.6T; when ψ = 0.3, Ft = 0.9T; when b = 6R and ψ = 0.2, Ft = T.

[0115] The propeller thrust T is parallel to the wing, and the propeller wake pressure Ft acts on the geometric center of the wing plane and is perpendicular to the wing. When b=6R, ψ=0.2, and Ft=T, if T=1.2mg and the wing angle of attack θ=60°, the wing-mounted aircraft is in torque balance and can maintain a slow, forward hovering flight. When b=6R, ψ=0.2, and Ft=T, if d1=d3, T=1mg, and the wing angle of attack θ=90°, the wing-mounted aircraft is also in torque balance and can maintain a slow, forward hovering flight.

[0116] The pitching moment of the axle-winged aircraft is generated solely by the propeller wake pressure Ft, which is somewhat insufficient. It is necessary to add vortices at the leading edge of the axle to generate lift Fh from the leading edge strake to improve the pitching moment and ensure flight stability and safety.

[0117] An aileron is provided at the rear of the wing. The aileron chord length is approximately one-third of the wing chord length, and the aileron span is slightly less than one-half of the wing span. When the aileron is dipping upwards, the propeller wake pressure Ft increases and its position moves rearward. When the aileron is dipping downwards, the propeller wake pressure Ft decreases and its position moves forward. When the wing is in rotorcraft flight mode, the magnitude, direction, and position of the propeller wake pressure Ft are adjusted by the aileron to control flight. When the wing is in fixed-wing flight mode and transitional flight mode, the magnitude, direction, and position of the angle of attack lift F and the propeller wake pressure Ft are adjusted simultaneously by the aileron to control flight.

[0118] The aforementioned winged aircraft, as a cross-domain aircraft, can continuously transition between rotorcraft and fixed-wing aircraft states, consisting of three stages: rotorcraft flight state, transitional flight state, and fixed-wing aircraft flight state. The degree of transition can be measured by the suspension-to-flight ratio ζ, which is a function of the angle of attack θ.

[0119] ,

[0120] Since the A-wing has no airfoil, it does not require airfoil lift and is not subject to the stall problems of fixed-wing aircraft. In fact, the flight state of the A-wing is the same as the stall state of a fixed-wing aircraft. The sustained flight speed Vbe of the A-wing is equal to the stall speed Vs, and the stall wind pressure Ws is equal to the sustained flight wind pressure Wbe. The critical stall condition for the A-wing in level flight is that the oncoming wind pressure equals the wing surface load: Wbe = mg / S.

[0121] When a winged aircraft flies horizontally at a constant speed with an arbitrary angle of attack θ, the oncoming wind pressure W corresponding to the flight speed V is:

[0122] , and ,

[0123] Given Wbe = mg / S, we can calculate:

[0124] ;

[0125] Let ζ be the suspension-to-flight ratio: Then we have: W = ζWbe.

[0126] Therefore, the ratio of horizontal flight speed V to stall speed Vs satisfies the following relationship:

[0127] That is: V 2 =ζVs 2 ;

[0128] The hover ratio ζ is equal to the ratio of the oncoming wind pressure during transition flight to the stall wind pressure during level flight, or the square of the ratio of the horizontal speed during transition flight to the stall speed.

[0129] ζ=W / Ws=V 2 / Vs 2 ;

[0130] As can be seen, the hovering ratio ζ is a dimensionless parameter, representing the degree to which the flight state of the winged aircraft gradually transitions from level flight to hovering flight state as the angle of attack θ increases while maintaining a constant speed and altitude. See the table below:

[0131]

[0132] As a cross-domain aircraft, the flight state of the armored wing transitions continuously between rotorcraft and fixed-wing aircraft as the angle of attack θ of the armored wing changes. The horizontal component Vl of the armored wing's flight velocity V decreases as the angle of attack θ increases: Vl = Vbe * If ζ=0, the A-wing aircraft is in rotorcraft flight mode, Vl=0; if 0<ζ<1, the A-wing aircraft is in transition flight mode, Vl= Vbe * <Vbe; ζ=1, then the A-wing aircraft is in fixed-wing flight state. When δ=1 / K, Vl = Vbe; when δ>1 / K, it enters sprint flight. The sprint speed is proportional to the thrust T. When Tmcp =βmg, Vl = Vmax>Vbe, and .

[0133] The hangar ratio ζ is a hyperbola with steep ends as a function of the angle of attack θ. When θ = 30°, then... = 0.9, δ=0.5, θ=80° then =0.32, δ=0.99, therefore increasing the angle of attack does not rapidly reduce the horizontal flight speed, but instead rapidly increases the thrust, thereby reducing the flight efficiency. When designing a winged aircraft, the stall speed should be reduced by decreasing the wing surface load, and the flight efficiency should be improved by decreasing the angle of attack.

[0134] Although the power reserves of the winged aircraft are insufficient to support the stable flight of the rotorcraft in hovering mode, it still has positive practical significance as an emergency take-off measure or flexible landing measure in emergency situations.

[0135] The wide-chord wing of the aircraft has the disadvantage of high induced drag and two-dimensional maneuvering characteristics, resulting in a large pitching moment and difficult control. In addition, the wing only has angle-of-attack lift and no airfoil lift. Therefore, when the angle of attack is 0, the lift disappears completely, making it prone to instability and violent nose-down at low speeds. Therefore, some special measures are needed to achieve the goals of stability enhancement, lift enhancement and drag reduction. These specific measures for stability enhancement, lift enhancement and drag reduction can be combined with the wing to form a lifting body. This can be achieved by setting flaps or strakes on the leading edge of the wing, or by bending the leading edge of the wing and adding canards to increase the control moment, or by forming an S-shaped airfoil with the wing and trailing edge ailerons to counteract each other, or by lowering the center of gravity of the aircraft to create a vertical trim moment to improve flight stability.

[0136] A flap is provided at the leading edge of the wing and combined with the wing to form a lifting body. The flap chord length is less than one-third of the wing chord length, and the flap span is not greater than the wing span. The airfoil of the flap is matched with the cruising speed of the aircraft. Generally, a low-speed airfoil is adopted, or it is curved into an appropriate airfoil surface.

[0137] Integrating a leading-edge extension (LES) with the wing to form a lifting body effectively improves flight stability. The LES chord length is less than half the wing chord length, and its span is no greater than the wing span. The longer the LES chord length or the larger the angle of attack, the greater the aircraft's lift and pitching moment, while simultaneously increasing form drag, decreasing induced drag, and reducing stall speed. With constant LES chord length and angle of attack, the faster the aircraft's speed, the greater the lift and pitching moment. The LES shape can be rectangular, rhomboid, elliptical, or triangular. The LES surface is located at the leading edge of the wing... When wing surfaces intersect, the angle between the leading-edge extension (LES) surface above the wing surface and the wing surface is the forward tilt angle, and the angle between the leading-edge extension surface below the wing surface and the wing surface is the sweep angle. Preferably, the leading-edge extension is planar, meaning the forward tilt angle and sweep angle are perpendicular. Preferably, the leading-edge extension has a triangular shape, and preferably its lower edge coincides with the leading edge of the wing and is dihedral. Dihedral leading-edge extensions reduce induced drag on the upper wing surface, while dihedral leading-edge extensions reduce induced drag on the lower wing surface. When using a diamond-shaped leading-edge extension, the long side of the diamond-shaped extension is parallel to the wing surface, and the forward tilt angle and sweep angle are equal, equivalent to... Combining a triangular upper dihedral wing and a triangular lower dihedral wing can increase lift while reducing induced drag on the wing surface, thus lowering the aircraft's stall speed. The leading-edge extensions can also be configured as movable leading-edge extensions with variable tilt angles. By changing the tilt angle γ, the aircraft's stall speed Vs is altered, thereby changing its sprint speed Vmax. When the leading-edge extension tilts to increase the frontal area, the aircraft's stall speed decreases; conversely, when the tilting angle decreases the frontal area, the aircraft's stall speed increases. Symmetrical leading-edge extensions resemble a rhombus. When the leading-edge extension (LEA) or elliptical leading-edge extension (LEA) is tilted, the aerodynamic center (trim) of the aircraft remains unchanged. A preferred control method is to divide the LEA into two parts, left and right, and tilt them separately. A vertically symmetrical stabilizer is set between the left and right LEAs. The simultaneous tilting of the left and right LEAs increases the frontal area, which can increase the lift of the wing and reduce the stall speed. The simultaneous reduction of the frontal area can reduce the lift of the wing and increase the stall speed. The differential tilting of the left and right LEAs can increase the lift difference between the left and right wing surfaces, thereby realizing the roll control and yaw control of the aircraft. At this time, the left and right horizontal tails can be combined into one and used only for pitch control.

[0138] The various specific measures for stabilizing, increasing lift, and reducing drag of the wing can be combined with the wing to form a lifting body. When the aircraft is used for manned purposes or has a large takeoff weight, multiple lifting bodies are integrated in a way that is connected in parallel laterally, or in series longitudinally, or in a staggered manner, so that the lift of each lifting body remains unchanged while the total lift increases, thereby improving the aircraft's load-bearing capacity.

[0139] The drive unit of the winged aircraft can be a closed propeller such as a turbine, a semi-closed propeller such as a duct, or an open propeller such as a rotor, and the installation position is not limited; when using an open propeller, the preferred installation position is located at the leading edge of the wing, and it is set as a forward-pulling layout.

[0140] When multiple propellers are distributed for propulsion, each propeller is paired with a wing to form a lifting body. Multiple lifting bodies are connected in parallel along the wingspan to form a single wing layer. Each wing layer is staggered vertically to form a multi-wing layout or staggered aft to form a waist-thrust wing layout. The multi-propeller distributed propulsion layout increases the effective area of ​​the propeller wake on the wing surface, helping to fully utilize the vortex lift generated by the propeller wake on the wing surface. It also reduces the difference in airflow velocity between takeoff, landing, and cruise flight states, reducing induced drag and improving flight stability. When the aircraft adopts a multi-wing layout, grouping the propellers allows for vertical takeoff and landing using multi-rotor flight control. When an aircraft adopts a multi-wing configuration, flight trim can be flexibly arranged near the combined aerodynamic center of multiple wings, offering greater flexibility than with a single wing. Furthermore, the tailwheel landing gear can be modified to have two adjustable height settings. Increasing the tailwheel landing gear height reduces the wing angle of attack for runway takeoffs and landings, while decreasing it increases the wing angle of attack for short takeoffs and landings. Alternatively, the tricycle landing gear can be expanded into a quintuple landing gear configuration, with the front three wheels for runway takeoffs and the rear four wheels for short takeoffs and landings, meeting the needs of cross-domain flight. Finally, the large-area flat wing of the aforementioned mecha-wing aircraft is ideally suited for photovoltaic installation, and a lift-to-drag ratio greater than 10 supports efficient flight, enabling solar-powered endurance.

[0141] If a wing-mounted aircraft with a wing loading of 10 kg / m² takes off and lands with a thrust of T=1.2mg and an angle of attack of θ=30°, the distance required for short takeoff is 0 meters. After takeoff, it switches to fixed-wing flight, and the stall speed for level flight is 13.5 m / s. For short landing, the wing-mounted aircraft descends with a speed of Vs=13.5 m / s, a thrust of T=0, and an angle of attack of θ=30°, and the distance required for landing is 16.9 meters. For manned wing-mounted aircraft with a wing chord length of more than 5 meters, taking off and landing within a distance of two or three aircraft positions is already considered ultra-short takeoff and landing.

[0142] The short takeoff and landing (STOVL) method utilizing angle-of-attack lift provided by this invention achieves four significant advantages through the application of a low-wing-loading, large-area shell-shaped wing: First, by utilizing angle-of-attack lift, the STOVL ensures sufficient lift is generated even at low speeds, breaking the dependence of fixed-wing airfoil lift on horizontal speed. Second, by employing hop takeoff and landing, the time window for STOVL is compressed, significantly reducing or even eliminating the required time. This not only reduces efficiency losses but, crucially, the takeoff and landing operation occurs at altitudes close to zero and speeds close to zero, greatly minimizing the risk exposure for flight safety. Third, a method for improving the flight efficiency of STOVL is proposed, providing direction for the design and application of low-wing-loading, large-area fixed-wing aircraft. Fourth, by utilizing the deflection of the propeller wake, the overlap between the flight envelopes of STOVL and horizontal flight at low speeds is increased, making flight control faster, more efficient, and more reliable.

[0143] The present invention achieves the following technical effects compared to the prior art:

[0144] This invention provides a novel flight mechanism for short takeoff and landing (STOVL) relying on angle-of-attack lift. While this mechanism has been applied in tethered aircraft such as kites, its autonomous flight mechanism has not been extensively studied, especially in manned aircraft. Existing aircraft wings all have airfoils, and their flight theory research focuses on airfoil lift, with insufficient attention paid to angle-of-attack lift. The significance of this invention lies in proposing a flight method for a winged aircraft, establishing a flight theory for an autonomous aircraft that relies entirely on angle-of-attack lift, encompassing design calculations, takeoff and landing methods, and flight control, forming a complete technical solution.

[0145] This invention provides a short takeoff and landing (STOVL) flight method that relies on angle-of-attack lift. Through an integrated lift design of propeller, wing armor, and stability enhancement, lift enhancement, and drag reduction measures, an independent lifting body structure is formed. This breaks the dependence of fixed-wing airfoil lift on horizontal speed, obtains sufficient lift during the low-speed takeoff phase, reduces vertical takeoff and landing time, allows direct jump into fixed-wing flight, improves flight efficiency, reduces risk exposure, and ensures flight safety.

[0146] The short takeoff and landing (STOVL) method provided by this invention overcomes the major defect of existing tiltrotor vertical takeoff and landing (TVTOL) aircraft where the flight control law changes abruptly when switching between horizontal and vertical flight modes. In three different flight states—horizontal flight mode, STOVL mode, and vertical takeoff and landing mode—the control parameters and control effects of the tiltrotor aircraft remain consistent, improving the stability and safety of flight mode transitions. Although the power reserve of the tiltrotor aircraft is insufficient to support stable flight in rotorcraft flight mode (vertical takeoff and landing mode), it still has positive practical significance as an emergency takeoff measure or flexible landing measure in emergency situations.

[0147] The flight method for a fixed-wing aircraft provided by this invention covers design calculations, takeoff and landing methods, and flight control, forming a complete technical solution. It proposes a method to improve the flight efficiency of fixed-wing aircraft and points the way for the design and application of low wing loading, large-area fixed-wing aircraft.

[0148] The wing provided by this invention is a large-area, low-wing-loading, thin-shell-shaped, wide-chord wing, which is equivalent to a rigid parachute that opens at zero altitude, ensuring a safety baseline.

[0149] The wing aircraft provided by this invention is compatible with three takeoff and landing modes: vertical, short takeoff and landing, and runway. It overcomes the low efficiency of traditional vertical takeoff and landing aircraft and avoids the discomfort caused to the crew by tail-seat vertical takeoff and landing aircraft. When used as a manned aircraft, it has outstanding advantages.

[0150] The wing aircraft provided by this invention has a flat wing surface, low wing loading, and large area, which simplifies the manufacturing process and provides ample space for photovoltaic installation. Due to the improved flight efficiency, it can achieve continuous flight using only half of the solar power generated, thus becoming a solar-powered aircraft capable of hop-on takeoff and landing.

[0151] The armored wing aircraft provided by this invention is specifically developed for low-altitude, low-speed, short-distance flight applications. Its general economic cruising speed is around 100 km / h, and its maximum flight speed is less than 200 km / h. Although its speed is relatively low compared to many existing fixed-wing aircraft, it has the aforementioned numerous advantages. As a flying car for daily short-distance commuting, it can still demonstrate significant advantages.

[0152] In summary, this invention makes aircraft operation more efficient, flight safer, structure more rational, performance more stable, and wider application more advantageous. Attached Figure Description

[0153] Figure 1 This is a side view of the short takeoff and landing (STOL) winged aircraft of the present invention.

[0154] Figure 2 This is a front view schematic diagram of the short takeoff and landing (STOL) winged aircraft of the present invention taking off.

[0155] Figure 3 This is a side view of the short takeoff and landing (STOL) winged aircraft of the present invention in level flight.

[0156] Figure 4 This is a frontal view of the short takeoff and landing (STOL) winged aircraft of the present invention in level flight.

[0157] Figure 5 This is a top-down view of the short takeoff and landing (STOL) winged aircraft of the present invention in level flight.

[0158] Figure 6 This is a comparison table of aircraft maneuverability characteristics according to the present invention.

[0159] Figure 7 This is a schematic diagram of the force analysis of the A-wing aircraft of the present invention flying at an angle of attack θ.

[0160] Figure 8 This is a schematic diagram of the horizontal flight trim analysis of the anti-delta headed wing aircraft of this invention.

[0161] Figure 9 This is a side view schematic diagram of the short takeoff and landing (STOVL) of the single-seat biplane armored aircraft of the present invention.

[0162] Figure 10 This is a side view of the single-seat biplane armored aircraft of the present invention in horizontal flight.

[0163] Figure 11 This is a frontal view of the single-seat biplane armored aircraft of the present invention in horizontal flight.

[0164] Figure 12 This is a top-down view of the single-seat biplane armored aircraft of the present invention in horizontal flight.

[0165] Figure 13 This is a top-down view of the single-seat, waist-thrust armored wing aircraft of the present invention in horizontal flight.

[0166] Figure 14 This is a side view schematic diagram of the single-seat, waist-thrust armored wing aircraft for short takeoff and landing according to the present invention.

[0167] Figure 15 This is a schematic diagram of the force analysis during takeoff of the A-wing aircraft according to the present invention.

[0168] Figure 16 This is a schematic diagram of the force analysis of the A-wing aircraft in level flight according to the present invention.

[0169] Figure 17 This is a schematic diagram of the hovering flight trim analysis of the A-wing aircraft of the present invention.

[0170] Figure 18 This is a frontal view of the twin-engine, multi-seat, monoplane armored aircraft of the present invention in horizontal flight.

[0171] Figure 19 This is a side view of the twin-engine, multi-seat, monoplane armored aircraft of the present invention in horizontal flight.

[0172] Figure 20 This is a top-down view of the twin-engine, multi-seat, monoplane armored aircraft of the present invention in horizontal flight.

[0173] Figure 21 This is a side view of the short takeoff and landing fixed-wing aircraft of the present invention.

[0174] Figure 22 This is a graph showing the relationship between the angle of attack γ of the leading edge spar wing and the stall speed Vs of the present invention.

[0175] In the diagram: 1-wing A; 2-fuselage; 3-drive unit; 4-leading slat; 5-vertical tail; 6-horizontal aileron; 7-nose landing gear; 8-rear landing gear; 9-angle of attack for takeoff and landing; 10-wingtip wing; 11-low-speed airfoil. Detailed Implementation

[0176] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0177] Example 1: As Figures 1-17 As shown, this embodiment provides a flight method for short takeoff and landing relying on angle of attack lift, including a low wing loading fixed wing with a straight airfoil configuration, called an A-wing aircraft.

[0178] This embodiment utilizes the principle of short takeoff and landing using angle-of-attack lift as follows:

[0179] This embodiment uses a square-shaped wing with a surface load of 30 kg per square meter, an aspect ratio of 3.5, and a thrust-to-weight ratio of 1.2, i.e., a thrust T = 1.2 mg. The angle of attack is 30°. Therefore, the horizontal component Tx and the vertical component Ty of the thrust are:

[0180] Tx=Tcos30°=1.2mg*0.866=1.04 mg; Ty=Tsin30°=1.2mg*0.5=0.6 mg;

[0181] The inertial drag of air, Gx, is: Gx = ax*ρ*Dk* S*sinθ;

[0182] Given air density ρ = 1.2 and Dk = 0.6 for a square wing, the following can be calculated:

[0183] Gx = 1.04g*1.2*0.6*S* sin30°= 0.37gS,

[0184] The inertial drag Gx creates wind pressure Wx under the wing surface, Wx = Gx / S / sin30° = 0.74g. This wind pressure Wx is transmitted vertically to the wing surface, creating lift F at the angle of attack, i.e.: F = Wx * S = 0.74gS.

[0185] The lift force F at the angle of attack can be decomposed into a horizontal component Fx and a vertical component Fy at the angle of attack θ, as follows:

[0186] Fx = F * sinθ = 0.74gS * sin30° = 0.37gS,

[0187] Fy = F * cosθ= 0.74gS* cos30°=0.64gS,

[0188] Fy and Ty have the same direction, both vertically upwards, therefore:

[0189] Fy+Ty=0.64gS+0.6mg=0.64*9.8*S+0.6mg=6.3S+0.6mg;

[0190] The wing surface load in this embodiment is 30 kg per square meter, i.e., mg / S = 30, then S = mg / 30, that is:

[0191] Fy+Ty=6.3S+0.6mg = 6.3* mg / 30+0.6mg = 0.21mg+0.6mg=0.81mg

[0192] At this point, the thrust is basically enough to overcome the takeoff weight (mg), and a short runway is sufficient for takeoff.

[0193] The takeoff distance for short takeoff is calculated by accelerating from 0 to Vbe with acceleration ax, therefore, the takeoff time t is:

[0194] ,

[0195] The distance s required for the gliding run is:

[0196] ,

[0197] In this embodiment, the wing surface load mg / S is 30 kg per square meter. Since the vertical component of the thrust Ty bears 0.6 mg, or 18 kg, of the takeoff weight, and the takeoff weight that needs to be lifted by the angle of attack is 12 kg, the takeoff time t required for the taxiing run is: The required distance s for takeoff is: s = 0.18 * 12 * 9.8 = 21 meters. For a heavy-load aircraft with a wing surface load of 30 kg per square meter, the fuselage length of the A-wing aircraft is greater than 6 meters, so it only needs to take off from 4 parking positions.

[0198] It is evident that the lower the wing loading of a mech, the shorter the takeoff distance. When the wing loading remains constant, the takeoff distance is related to the takeoff angle of attack and the thrust-to-weight ratio. When the angle of attack remains constant, a higher thrust-to-weight ratio results in a shorter takeoff distance, while a lower thrust-to-weight ratio results in a longer takeoff distance. When the thrust-to-weight ratio remains constant, a higher angle of attack results in a shorter takeoff distance, while a lower angle of attack results in a longer takeoff distance.

[0199] The principle of lift flight utilizing angle of attack in this embodiment is as follows:

[0200] When the A-wing aircraft is in level flight with a speed of V, the stationary air in front of it is washed down at a constant speed of V by the wing surface, which also causes the air under the wing to have an acceleration a and an angle-of-attack lift F as follows:

[0201] F = Mk*a = V / t * *V*t*S*sinθ= 1.2 *V 2 *S*sinθ;

[0202] When the speed V is constant, increasing the angle of attack θ can increase the lift. When the angle of attack θ is constant, increasing the speed V can increase the lift. When the angle of attack θ = 0, no matter how large the speed V is, the lift F at the angle of attack is 0.

[0203] When a mech-winged aircraft flies horizontally with a thrust T and an angle of attack θ, the oncoming wind pressure W is: W = T / Sd. This oncoming wind pressure compresses the air under the wing of the mech-winged aircraft. The pressure of the compressed gas is equal in all directions, thus generating lift F in the normal direction of the wing, F = S*W, that is: F = S*W = S*T / Sd = T*S / Sd = T*K. Therefore, the mech-winged aircraft obtains an angle-of-attack lift F that is K times T with a thrust of T. The mech-winged aircraft flies precisely by utilizing this angle-of-attack lift mechanism.

[0204] The surface loading of the wingless aircraft, mg / S, is the ratio of its takeoff weight, mg, to its wing area, S. When the wingless aircraft maintains a constant altitude and speed with thrust T and an angle of attack θ, the thrust T is δ times its takeoff weight, mg, i.e., T = δmg, where δ is the thrust-to-attack ratio of the wingless aircraft.

[0205] ;

[0206] The thrust-to-attack ratio δ is a dimensionless parameter representing the ratio of the thrust T required by the drive system to the takeoff weight mg for a winged aircraft to maintain a constant altitude and speed at an angle of attack θ. For θ < 10°, δ < 0.1; for θ > 60°, δ > 0.9, it exhibits a hyperbola with relatively steep extremes. See the table below:

[0207]

[0208] It can be seen that the thrust-to-attack ratio δ decreases rapidly as the angle of attack θ decreases, which means that the thrust T of the drive device required to maintain flight altitude decreases rapidly. For an armored aircraft to maintain flight altitude at the wing angle of attack θ, the thrust T of the drive device needs to satisfy: T = δmg.

[0209] Let Tl be the level flight thrust of the A-wing aircraft, and θl be the level flight angle of attack of the A-wing aircraft. When the thrust T=Tl remains constant, θ=θl means the A-wing aircraft is flying at level; θ>θl means the A-wing aircraft is descending, with a decrease in speed and altitude; θ<θl means the A-wing aircraft is climbing, with an increase in speed and altitude. When the angle of attack θ=θl remains constant, T=Tl means the A-wing aircraft is flying at level; T<Tl means the A-wing aircraft is descending, with a decrease in speed and altitude; T>Tl means the A-wing aircraft is climbing, with an increase in speed and altitude.

[0210] The thrust required to maintain the flight altitude of the winged aircraft is the minimum level flight thrust Tmin: Tmin = δmg = mg / K. After this, even if the angle of attack θ is further reduced to 0≦δ≦1 / K, the thrust will not decrease further.

[0211] The minimum level flight thrust Tmin corresponds to the constant horizontal flight speed, which is the minimum horizontal flight speed of the armored wing aircraft, also known as the long-range flight speed Vbe. The wind pressure corresponding to the long-range flight speed is equal to the wing surface load, i.e.:

[0212] ,

[0213] C is the air drag coefficient; for the A-wing aircraft, c = 0.95. With an air density of 1.2 kg / m³ 3 ,have to:

[0214] ,

[0215] The wind pressure corresponding to Vbe is equal to the wing surface load, and the angle of attack corresponding to Vbe is the long-range aircraft's angle of attack θbe.

[0216] When 90°>θ>θbe, the winged aircraft is in a transitional flight state. The horizontal component of the flight speed V, Vl, is less than the sustained flight speed Vbe, that is: Vl<Vbe. Increasing the thrust T can only increase the flight altitude, but cannot increase the horizontal flight speed Vl, and cannot perform sprint flight. The larger θ is, the smaller Vl is.

[0217] When θ=90°, Vl=0, and the A-wing aircraft is in rotorcraft flight mode.

[0218] When θ≦θbe, the A-wing aircraft is in fixed-wing flight mode, and its flight speed is directly proportional to its thrust T. If δ<1 / K, the A-wing aircraft stalls at a speed equal to its sustained flight speed Vbe. The A-wing aircraft continuously descends to maintain a horizontal speed equal to Vbe. Let β be the thrust-to-weight ratio of the A-wing aircraft during its sprint flight, and let Tmcp be the thrust corresponding to the maximum sustained output power of the A-wing aircraft, Tmcp=βmg, and β>δ. If β≧1 / K, the A-wing aircraft enters sprint flight, and increasing the thrust at this point will increase its horizontal flight speed.

[0219] Let Wbe be the long-range wind pressure of the armor-wing aircraft. The long-range wind pressure is equal to the wing surface loading:

[0220] Wbe = Tmin / (S / K) = mg / S;

[0221] Let the sprint wind pressure of the winged aircraft be Wmax, and the sprint thrust of the winged aircraft be β times its takeoff weight. Since Tmcp = βmg, then the sprint wind pressure Wmax = Tmcp / (S / K) = βKmg / S.

[0222] According to the general formula for wind speed and wind pressure:

[0223] ,

[0224] The maximum horizontal sprint speed Vmax can be calculated as follows:

[0225] ;

[0226] That is, the maximum horizontal sprint speed Vmax is equal to the sustained flight speed Vbe. times, that is: , where β is the multiple of the maximum sprint thrust Tmcp to the takeoff weight mg, i.e.: Tmcp=βmg, and β>δ.

[0227] Since the long-range speed Vbe is determined by the wing loading, the sprint speed Vmax of the armored wing aircraft is determined by the sprint thrust Tmcp, lift-to-drag ratio K, and wing loading mg / S, and the sprint speed is proportional to all three.

[0228] As a cross-domain aircraft, the A-wing can continuously transition between the flight states of a rotorcraft and a fixed-wing aircraft, consisting of three stages: rotorcraft flight state, transitional flight state, and fixed-wing aircraft flight state. The degree of transition can be measured by the suspension-to-flight ratio ζ, which is a function of the angle of attack θ.

[0229] ;

[0230] The flight state of the A-wing aircraft transitions continuously between rotorcraft and fixed-wing aircraft as the angle of attack θ of the A-wing changes. The horizontal component Vl of the A-wing aircraft's flight velocity V decreases as the angle of attack θ increases, Vl = Vbe * If ζ=0, the A-wing aircraft is in rotorcraft flight mode, Vl=0; if 0<ζ<1, the A-wing aircraft is in transition flight mode, Vl= Vbe * <Vbe; ζ=1, then the A-wing aircraft is in fixed-wing flight state. When δ=1 / K, Vl = Vbe. When β>δ=1 / K, it enters sprint flight. The sprint speed is proportional to the thrust T. When Tmcp =βmg, Vl = Vmax>Vbe, and ;

[0231] It can be seen that when ζ=0, the A-wing aircraft is in the state of rotorcraft flight, and Vl=0. When the thrust-to-weight ratio of the A-wing aircraft is greater than 1, the thrust is greater than the takeoff weight, and both hovering and vertical landing with a takeoff distance of 0 can be achieved. Although the power reserve is insufficient to support the stable flight of rotorcraft in the state of hovering flight, it still has positive practical significance as an emergency takeoff measure or flexible landing measure in emergency situations.

[0232] When the A-wing aircraft is in rotorcraft flight mode and transitional flight mode, its uniform horizontal flight speed is less than or equal to the long-range flight speed Vbe. Increasing thrust can only increase the climbing speed and cannot exceed the long-range flight speed Vbe. Only when the A-wing aircraft's angle of attack decreases to θ≦θbe, allowing it to enter fixed-wing aircraft flight mode, can increasing thrust increase the horizontal flight speed and achieve sprint flight. The sprint speed is proportional to the thrust T.

[0233] The wing load of this embodiment is 30 kg / m. 2 Given Wbe = mg / S, and taking the air drag coefficient c = 0.9, the corresponding stall speed is approximately 84 km / h.

[0234] ;

[0235] Considering the economy of the propulsion system and the comfort of hop-off flight for manned aircraft, the multiple of sprint thrust to takeoff weight is taken as β≦1.2, and the lift-to-drag ratio is taken as K=10. Calculations show that the maximum flight speed of the armor-wing aircraft is approximately 290 kilometers per hour.

[0236] ;

[0237] Trial is a crucial prerequisite for aircraft flight. The flight trim of a mech-wing aircraft is determined by the propeller position, the overall center of gravity, and the position of the mech's aerodynamic focus. In this embodiment, the flight trim is assumed to have a propeller eccentricity coefficient ψ=0, resulting in zero propeller wake pressure. The flight trim is determined by the overall center of gravity and the position of the mech's aerodynamic focus. By forward and downward positioning of the battery compartment, the center of gravity is adjusted to be below and in front of the aerodynamic focus, improving flight stability.

[0238] The method of using angle-of-attack lift for flight provided in this embodiment is a novel flight mechanism. This mechanism has been applied in tethered aircraft such as kites, but its flight mechanism has not yet been studied in depth, especially in autonomous aircraft. Existing aircraft wings all have airfoils, and their flight theory research focuses on airfoil lift, with angle-of-attack lift playing a secondary role. The significance of this invention lies in proposing a flight method for a winged aircraft, establishing a flight theory for an autonomous aircraft that relies entirely on angle-of-attack lift, covering design calculations, takeoff and landing methods, and flight control, forming a complete technical solution.

[0239] The wing provided in this embodiment is a large-area, low-wing-loading, thin-shell-shaped, wide-chord wing, which is equivalent to a rigid parachute that opens at zero altitude, ensuring a safety baseline.

[0240] The flight method for the axle-winged aircraft provided in this embodiment makes aircraft operation more efficient, flight safer, structure more rational, and performance more stable.

[0241] Example 2, as Figures 1-17 As shown, this embodiment uses a square-shaped wing with a wing surface load of 20 kg per square meter, an aspect ratio of 2, and a thrust-to-weight ratio of 1.2, i.e., a thrust T = 1.2 mg. The angle of attack is 30°, so Fy + Ty = 0.92 mg. At this point, the thrust is close to the takeoff weight. Since the vertical component of the thrust, Ty, bears 0.6 mg (12 kg) of the takeoff weight, the takeoff weight that needs to be lifted by the angle of attack is 8 kg. Therefore:

[0242] The time t required for takeoff roll is: Second;

[0243] The distance s required for the gliding run is: s = 0.18 * 8 * 9.8 = 14 meters;

[0244] For heavy-load aircraft with a wing surface load of 20 kg per square meter, the fuselage length of the A-wing aircraft is greater than 5 meters, so only 3 taxiing positions are required for takeoff. The remaining structure and principles are the same as in Example 1.

[0245] Example 3, as Figures 1-17 As shown, this embodiment uses a square wing with a wing surface load of 10 kg per square meter, an aspect ratio of 1, and a thrust-to-weight ratio of 1.2, i.e., a thrust T = 1.2 mg. The pitch angle is 30°, so Fy + Ty = 1.23 mg. At this point, the thrust is greater than the takeoff weight, allowing for immediate takeoff with a takeoff run of 0. The remaining structure and principle are the same as in Embodiment 1.

[0246]

[0247] Example 12: As Figures 1-5 , Figures 7-17 As shown, this embodiment provides a method for a mech-wing aircraft to achieve vertical takeoff and landing by controlling trim. The trim of the mech-wing aircraft is determined by the propeller position, the overall center of gravity position, and the aerodynamic focus position of the mech. By controlling trim, this embodiment can achieve vertical takeoff and landing.

[0248] Let the span of wing A be l, the chord length be b, and the area of ​​wing A be S: S = l * b. Let the propeller diameter be D, the radius be R, the area of ​​the propeller disk be A, A = π * R * R, and the eccentricity between the propeller shaft center and the surface of wing A be E1. Then the propeller disk is divided into an upper part A1 and a lower part A2 by wing A, where A1 > A2. Therefore, the pressure of the propeller wake on the surface of wing A is unequal, with the pressure on the surface of wing A being greater than the pressure below the surface. Thus, a downward pressure Ft is generated at the aerodynamic focus of wing A. If the eccentricity E1 = 0, that is, wing A bisects the propeller disk, then Ft = 0. As E1 increases, Ft increases. On the other hand, if the eccentricity E1 ≥ R, that is, wing A is tangent to the propeller disk or not, then the propeller wake leaves the surface of wing A, and Ft = 0. That is, when the eccentricity 0 < E1 ≤ R, the pressure Ft of the propeller wake on the wing surface is not 0, and the downward pressure Ft generated by the propeller wake at the aerodynamic focus of the wing surface is:

[0249] ,

[0250] Let Ω be the tail pressure coefficient of the armor-wing aircraft. When b, T, and R are all determined, the pressure Ft in the propeller wake actually becomes a function of the eccentricity E1, that is: ;

[0251] Let the eccentricity E1 be: E1 = ψR. When ψ = 0.0, 0.1~0.9, and 1.0, the tail pressure coefficient Ω, which generates downward pressure at the aerodynamic focus of the propeller wake on the wing surface, can be calculated as:

[0252]

[0253] It can be seen that when ψ > 0.5, Ω > 0.5; when ψ = 0.7, Ω is at its maximum, Ωmax = 0.58.

[0254] When b=4R and ψ=0.2, Ft=0.6T; when ψ=0.3, Ft=0.9T; when b=6R and ψ=0.2, Ft=T.

[0255] The takeoff mass of the A-wing aircraft is m, the gravity mg is perpendicularly downward along the center of gravity, the angle between gravity and the A-wing is equal to 90° minus the angle of attack of the A-wing, and the distance between the center of gravity of the A-wing aircraft and the surface of the A-wing is e2.

[0256] The propeller thrust T is parallel to the wing, and the pressure of the propeller wake on the wing is Ft. Ft is a function of the propeller thrust T and is proportional to the propeller thrust T. Ft = f(T). Ft acts on the aerodynamic focus and is perpendicular to the wing.

[0257] When b=6R, ψ=0.2, and Ft=T, if T=1.2mg and the angle of attack of the wing A is θ=60°, the moment balance of the wing A aircraft is achieved, and it can maintain a slow forward hovering flight at the pitch angle.

[0258] When b=6R, ψ=0.2, and Ft=T, if d1=d3, T=1mg, and the angle of attack of the wing A is θ=90°, the moment balance of the wing A aircraft is achieved, and it can also maintain a slow forward hovering flight at the angle of attack.

[0259] When a winged aircraft hovers, the wing surface does not move relative to the air; therefore, the equilibrium condition is independent of the aerodynamic center AC, but depends on the center of gravity CG, i.e.:

[0260] Torque balance is given by: T*(e1+e2) = Ft*(d3+d4).

[0261] Force balance gives: mg = T*sinθ - Ft*cosθ,

[0262] When the armored aircraft is flying horizontally, a strake is installed on the leading edge of the armored wing, which moves the angle-of-attack lift forward to before the aerodynamic center AC of the wing surface. Therefore, the equilibrium condition can be briefly described as follows:

[0263] Torque balance is given by: T*e1 = mg*d5 + Ft*(d3+d5)

[0264] Force balance gives: mg = F - Ft,

[0265] Therefore, once the geometry of the A-wing aircraft is determined, the tail pressure coefficient Ω is also determined. The lever arm (distance from the aerodynamic center) of the propeller thrust T and the wake pressure Ft is also determined. The torque exerted by the propeller thrust T and gravity mg on the A-wing aircraft is the pitching torque, and the torque exerted by the propeller wake pressure Ft on the A-wing aircraft is the pitching torque. When the A-wing aircraft is in level flight, the ailerons are level with the A-wing surface, and the distance between the propeller wake pressure Ft and the aerodynamic center CG is 0. As the ailerons of the A-wing aircraft dihedral, the distance between the propeller wake pressure Ft and the aerodynamic center CG increases to d4, and the pitch angle of the A-wing aircraft reaches θ. At this point, the trim torque of the A-wing aircraft needs to satisfy the following equation:

[0266] .

[0267] It is evident that the tilting of the aileron alters the position of the propeller wake pressure Ft, thus changing the trim moment and consequently causing a change in the pitch angle θ of the wingless aircraft.

[0268] The axle-wing aircraft provided in this embodiment can achieve hovering flight and vertical take-off and landing through control trim. Although the power reserve of the axle-wing aircraft is not sufficient to support the stable flight of the rotorcraft in hovering flight mode, it still has positive practical significance as an emergency take-off measure or flexible landing measure in emergency situations.

[0269] Example 13: As Figure 1-5 , Figure 7 , Figure 8 , Figure 15 - 17、 Figure 22 As shown, this embodiment provides a flight method for an armored wing aircraft to increase lift using a leading-edge extension wing lifting body. The structure and parameters of the armored wing aircraft provided in this embodiment are the same as those in Embodiment 1. The armored wing aircraft includes an armored wing 1, a fuselage 2, a drive unit 3 located at the leading edge of the armored wing 1, a leading-edge extension wing 4, a vertical tail 5, a horizontal aileron 6, a front landing gear 7, and a rear landing gear 8.

[0270] A strake 4 is provided on the leading edge of the wing 1. The chord length of the strake 4 is less than half the chord length of the wing 1, and the span of the strake 4 is slightly less than the span of the wing 1. The larger the chord length of the strake 4, the greater the lift and the greater the torque, which can effectively improve the flight stability of the wing 1. The preferred shape of the strake 4 is triangular, and it is preferably set as a fixed wing with dihedral angle. The larger the dihedral angle, the greater the lift and the smaller the induced drag, but the greater the form drag. The strake 4 can also be set as a movable strake 4 with variable dihedral angle to achieve more complex flight control. A preferred control method is to divide the strake 4 into two parts, left and right, and tilt them separately. The left and right strake 4 tilt simultaneously to increase the forward tilt angle, which can increase the lift of the wing 1 and reduce the stall speed. The left and right strake 4 tilt simultaneously to decrease the forward tilt angle, which can decrease the lift of the wing 1 and increase the stall speed. The left and right strake 4 tilt differentially to increase the lift difference between the left and right wing surfaces, which can achieve roll control of the wing 1. At this time, the left and right horizontal tails can be combined into one and used only to control the pitch of the wing 1.

[0271] The rest of the structure is the same as in Example 1.

[0272] Example 14: As Figures 1-5 , Figures 7-17 , Figure 21 As shown, a method for enhancing lift in a mech-wing aircraft using flaps as lifting bodies is described. Flaps are installed on the leading edge of the mech 1. The flap chord length is less than one-third of the mech chord length, and the flap span is slightly less than the mech span. The flaps have a certain low-speed airfoil. During level flight, the flap airfoil generates lift to counteract the nose-hing tendency of the mech-wing aircraft, thus improving its flight stability. The remaining structure and principle are the same as in Embodiment 13.

[0273] Example 15: As Figure 1-5 , Figures 8-17 As shown, a method for solar-powered flight of a mech-wing aircraft is provided. Because the mech-wing aircraft provided by this invention has high flight efficiency, and the mech 1 has a flat wing surface, low wing loading, and large area, it provides ample space for photovoltaic installation. Solar cells can be installed on the surface of the mech 1, enabling solar-powered flight and making the mech-wing aircraft provided by this invention a solar-powered aircraft capable of short takeoff and landing. The remaining structure is the same as in Embodiment 1.

[0274] Example 16: As Figures 1-17 , Figure 21 As shown, this embodiment provides a flight method for short takeoff and landing relying on angle-of-attack lift, including a low-wing-loading fixed wing with a flying wing configuration featuring a low-speed airfoil.

[0275] The low-speed airfoil in this embodiment is as follows: Figure 21 As shown.

[0276] The remaining structure and principle of the low wing loading fixed wing provided in this embodiment are the same as those in Embodiment 1.

[0277] Example 17: As Figures 1-17 As shown, this embodiment uses a circular wing with a surface load of 25 kg per square meter, a thrust-to-weight ratio of 1.5, an angle of attack of 60°, and the vertical component Ty of the thrust T is:

[0278] Ty = T * sin60°= 1.5mg* sin60°= 1.3mg.

[0279] At this point, the pull Ty completely overcomes the takeoff weight mg, allowing for takeoff from the spot.

[0280] The remaining structure and principle of the winged aircraft provided in this embodiment are the same as those in Embodiment 1.

[0281]

[0282] Example 23: As Figure 9-12 As shown, this embodiment provides an aerodynamic layout method for a manned biplane armored aircraft. The armored aircraft includes an armored wing 1, a fuselage 2, and a drive unit 3 located at the leading edge of the armored wing 1. The armored wing 1 is rectangular, the fuselage 2 is a single-seat manned cabin, and the drive unit 3 consists of multiple propellers distributed for drive. Each propeller is matched with one armored wing to form a lifting body. The multiple armored wings are connected in parallel along the wingspan to form a single wing layer. The two wing layers are staggered vertically to form a double-wing layout. The manned cabin is located in front of and below the aerodynamic center of the double-wing. The manned biplane armored aircraft adopts a layout with multiple propellers distributed for drive, which increases the effective area of ​​the propeller wake on the wing surface. This helps to fully utilize the vortex lift generated by the propeller wake on the armored wing surface, while reducing the difference in airflow velocity on the wing surface during takeoff, landing, and cruise, thus improving flight stability. The rear two-point landing gear of the tricycle landing gear is expanded into a quintuple landing gear, with the front three wheels being runway landing gear and the rear four wheels being short takeoff landing gear, meeting the cross-border flight requirements of the armored aircraft. The remaining structure and principle of the winged aircraft provided in this embodiment are the same as those in Embodiment 13.

[0283] Example 24: As Figures 9-12As shown, this embodiment provides an aerodynamic layout method for a bi-deck manned wing aircraft. When the manned wing aircraft adopts a bi-deck or multi-deck wing layout for vertical takeoff and landing, the front drive unit can increase the throttle to lift the front of the fuselage during takeoff with the support of the rear landing gear when the front landing gear leaves the ground. After the angle of attack reaches the required takeoff angle, the rear drive unit increases the throttle to complete the takeoff. During landing, the throttle of the rear drive unit is immediately reduced after the rear landing gear touches the ground. After the aircraft is firmly seated, the throttle of the front drive unit is gradually reduced to slowly lower the front landing gear for landing. This makes the operation of the front and rear drive units more intuitive and smooth, easier to control, and the takeoff and landing process smoother and more gentle, reducing the discomfort caused by tail-seat vertical takeoff and landing.

[0284] Jump-wing aircraft can be prioritized for applications in water sports, and can also take off and land in small ponds or squares in remote villages; of course, when taxiing conditions are available, take-off and landing should be mainly based on runway take-off and landing with an angle of attack θ of less than 30°, with short take-off and landing as a supplement, and vertical take-off and landing only as an emergency backup.

[0285] The remaining structure and principle of the winged aircraft provided in this embodiment are the same as those in Embodiment 23.

[0286] Example 25: As Figure 13 , 14 As shown, this embodiment provides a waist-thrust aerodynamic layout method for a single-layer manned armored wing aircraft. The armored wing aircraft includes an armored wing 1, a fuselage 2, and a drive unit 3 located at the leading edge of the armored wing 1. The fuselage 2 is a multi-seat manned cabin. The remaining structure and principle of the armored wing aircraft provided in this embodiment are the same as those in Embodiment 1. Multiple propellers are distributed for drive, with each propeller matched to one armored wing. The multiple armored wings are staggered along the chord direction to form a waist-thrust layout. The layout of multiple propellers distributed for drive increases the effective area of ​​the propeller wake on the wing surface, which helps to fully utilize the vortex lift generated by the propeller wake on the armored wing surface, while reducing the difference in airflow velocity on the wing surface during takeoff, landing, and cruise, thereby improving flight stability.

[0287] Example 26: This example provides an aerodynamic layout method for a twin-engine, multi-seat, single-layer wing manned armored aircraft, such as... Figure 2 , Figures 4-6 , Figures 18-20 As shown, the wingless aircraft includes a wing 1, a fuselage 2, and a drive unit 3 located at the leading edge of the wing 1. The fuselage 2 is a multi-seat cabin. The remaining structure and principle of the wingless aircraft provided in this embodiment are the same as those in embodiment 1.

[0288] Example 27: This example uses a variable-angle diamond-shaped strake wing, such as... Figure 22As shown, the rhomboid leading-edge extension is planar, with its long side parallel to the surface of the wing A. It is configured as a movable leading-edge extension with variable tilt angle. The chord length of the rhomboid leading-edge extension is one-third of the chord length of the wing A, and its span is four-fifths of the wing A's span. The long side of the rhomboid leading-edge extension coincides with the surface of the wing A. It is connected to the wing A via a hinge and can tilt around the wing A by an angle γ, where γ3 > γ > γ1. The stall speed Vs of the wing A decreases as the angle of attack γ of the leading-edge extension increases. By changing the tilt angle γ of the leading-edge extension, the stall speed Vs of the wing A is changed, thereby changing the sprint speed Vmax of the wing A. That is, when γ3>γ2>γ1, Vs1>Vs2>Vs3, then Vmax1>Vmax2>Vmax3. That is, when other flight parameters remain unchanged, the flight speed of the A-wing aircraft can be increased by reducing the leading-edge slat tilt angle γ. When the diamond-shaped leading-edge slat tilts, the aerodynamic center (trimming) of the A-wing aircraft remains unchanged, realizing more complex flight control. The remaining structure and principle of the A-wing aircraft provided in this embodiment are the same as those in embodiment 13.

[0289] Example 28: As Figure 22 As shown, this variable tilt elliptical strake wing has an elliptical strake wing, the major axis of which coincides with the surface of the wing A. The rest of the structure and principle are the same as in embodiment 27.

[0290] Example 29: As Figure 22 As shown, this variable tilt angle triangular strake wing has a triangular strake wing, with the long side of the triangular strake wing coinciding with the surface of the first wing. The rest of the structure and principle are the same as in embodiment 27.

[0291] Specific examples have been used to illustrate the principles and implementation methods of this invention. The descriptions of the above embodiments are only for the purpose of helping to understand the method and core ideas of this invention. Furthermore, those skilled in the art will recognize that, based on the ideas of this invention, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of this invention.

Claims

1. A method of flying short take-off and landing by angle of attack lift, characterised by: The aircraft has a low wing with low wing load and wide chord, the wing surface load is not greater than 30 kg per square meter, and the aspect ratio is not greater than 3.5; when the thrust-weight ratio is not less than 1.2 and the angle of attack is not less than 30°, more than half of the take-off weight is supported by the vertical component of the thrust and part of the take-off weight is lifted by the angle of attack lift generated by the large area wing during acceleration to realize short take-off and landing; the smaller the wing surface load, the shorter the sliding distance; when the angle of attack is constant, the larger the thrust-weight ratio, the shorter the sliding distance, and the smaller the thrust-weight ratio, the longer the sliding distance; when the thrust-weight ratio is constant, the larger the angle of attack, the shorter the sliding distance, and the smaller the angle of attack, the longer the sliding distance.

2. The method of short takeoff and landing by lift dependent angle of attack according to claim 1, characterized in that: The wing of the aircraft is a thin shell rigid wing or a thin shell wing covered on a rigid frame, the wing surface is a plane without airfoil curve, and no airfoil lift is generated, which is called "A wing"; the lift of the A wing is derived from the angle of attack lift; the angle of attack lift is generated due to the inertial resistance of the air in front of the wing acting on the lower surface of the wing, which is related to acceleration; when the wing starts to accelerate from 0 in the static air at an angle of attack θ, the initial speed of the oncoming airflow is 0, and the acceleration of the wing forces the airflow to wash down, so that the washing down airflow generates acceleration; when the wing moves at a constant speed V horizontally in the static air at an angle of attack θ, the air washes down at a constant speed V under the action of the wing surface, which also makes the washing down airflow generate acceleration; the reaction force generated by the acceleration of the washing down airflow is the angle of attack lift of the wing; when the driving mechanism is a propeller, the preferred installation position is located at the leading edge of the wing; the preferred landing gear layout is a front three-point landing gear; if the wing surface load is 20 kg per square meter, when the thrust-weight ratio is 1.2 and the angle of attack is not less than 30°, short take-off and landing with a sliding distance of less than 3 aircraft can be realized; if the wing surface load is 20 kg per square meter, when the thrust-weight ratio is 1.2 and the angle of attack is not less than 60°, or when the thrust-weight ratio is 1.5 and the angle of attack is not less than 45°, or when the thrust-weight ratio is 2 and the angle of attack is not less than 30°, or when the wing surface load is not greater than 10 kg per square meter, the thrust-weight ratio is not less than 1.2 and the angle of attack is not less than 30°, the take-off sliding distance is 0.

3. The method of short takeoff and landing by lift dependent angle of attack according to claim 2, characterized in that: The aircraft composed of the A wing is called A wing aircraft; the A wing aircraft is a flying wing layout aircraft, its maneuvering dimension is two-dimensional maneuvering, i.e. forward maneuvering and up-down maneuvering, which is between one-dimensional maneuvering of fixed wing aircraft, i.e. forward maneuvering, and three-dimensional maneuvering of rotary wing aircraft, i.e. forward maneuvering, up-down maneuvering and left-right maneuvering; as a cross-border aircraft between one-dimensional maneuvering aircraft and three-dimensional maneuvering aircraft, the flight state of the A wing aircraft is between fixed wing aircraft flight state, transition flight state and rotary wing aircraft flight state, the advantages and disadvantages of the A wing aircraft are also between rotary wing aircraft and fixed wing aircraft, and its characteristics are exactly suitable for the application scene of flying car; the large area low wing load thin shell rigid wing of the A wing aircraft is equivalent to a rigid parachute opened at 0 height, which has essential safety, even if the horizontal speed is 0, the wing surface load close to the parachute ensures the vertical landing and low speed gliding of the A wing aircraft, and the gliding flight has strong controllability.

4. The method of short takeoff and landing by lift dependent angle of attack according to claim 3, characterized in that: The thickness of the wing is less than one-twentieth of the chord length, and the basic wing type is a thin shell flat plate without camber; the preferred shape of the wing is rectangular, diamond, trapezoidal, triangular, or oval; the preferred wing surface load of the wing machine is less than 10 kg per square meter, and the wing machine relies on the large-area wing to generate the angle of attack lift to fly; the lift-drag ratio of the wing machine is K, K=S / Sd, wherein S is the area of the wing, and Sd is the minimum wind resistance area of the wing machine, that is, the windward area of the wing machine in horizontal flight; the thickness of the thin shell flat plate type wing is very small, the windward area in horizontal flight is very small, Sd is much smaller than S, so the lift-drag ratio K=S / Sd of the wing machine is greater than 10, and the highest can reach more than 30.

5. The method of short takeoff and landing by lift dependent angle of attack according to claim 4, characterized in that: The horizontal flight of the wing machine is related to the thrust T and the angle of attack θ; if the take-off mass of the wing machine is m, the take-off weight is mg, and the wing surface load is mg / S, when the wing machine maintains a high horizontal uniform speed flight with a thrust T and an angle of attack θ, the thrust T is δ times of the take-off weight mg, that is: T=δmg, and δ is the thrust angle ratio of the wing machine: When the thrust T=Tl is kept unchanged, θ=θl, the wing machine flies horizontally, θ>θl, the wing machine descends, the speed decreases, and the height decreases, and θ<θl, the wing machine climbs, the speed increases, and the height increases; when the angle of attack θ=θl is unchanged, T=Tl, the wing machine flies horizontally, T<Tl, the wing machine descends, the speed decreases, and the height decreases, and T>Tl, the wing machine climbs, the speed increases, and the height increases.

6. The method of short takeoff and landing by lift dependent angle of attack according to claim 5, characterized in that: When the thrust angle ratio δ of the wing machine is 1 / K, the thrust that maintains the flight height of the wing machine is the minimum horizontal flight thrust Tmin: Tmin=δmg=mg / K, and even if the angle of attack θ continues to decrease to make 0≦δ≦1 / K, the thrust will not decrease again; the uniform horizontal flight speed corresponding to the minimum horizontal flight thrust Tmin is the minimum horizontal flight speed of the wing machine, that is, the stall speed Vs, and this speed is also the endurance speed Vbe of the wing machine, that is, Vbe=Vs; the endurance speed Vbe of the wing machine is proportional to the wing surface load mg / S, that is: The angle of attack corresponding to the flight speed Vbe is the flight angle of attack θbe. When 90° > θ > θbe, the A-wing aircraft is in a transitional flight state, and the horizontal component Vl of the flight speed V is less than the flight speed Vbe, i.e., Vl < Vbe. Increasing the thrust T can only increase the flight altitude, but cannot increase the horizontal flight speed Vl, and it cannot perform a sprint flight. The larger θ is, the smaller Vl is. When θ = 90°, Vl = 0, and the A-wing aircraft is in a rotorcraft flight state. Let β be the thrust-to-weight ratio of the A-wing aircraft during sprint flight. When θ ≦ θbe and β ≧ 1 / K, the A-wing aircraft is in a fixed-wing flight state and can perform a sprint flight. The sprint speed is proportional to the thrust T, and the maximum horizontal sprint speed Vmax is the flight speed Vbe. times, that is: , where β is the multiple of the sprint thrust Tmcp to the takeoff weight mg, i.e.: Tmcp=βmg, and β≧1 / K.

7. The method of short takeoff and landing by lift dependent angle of attack according to claim 6, characterized in that: The flight trim of the wing machine is determined by the position of the propeller, the position of the center of gravity of the whole machine, and the position of the aerodynamic focus of the wing; the center of gravity of the whole machine is set below and in front of the aerodynamic focus of the wing, if the propeller axis coincides with the wing surface, the line connecting the center of gravity of the whole machine and the aerodynamic focus is the horizontal line of the wing machine in horizontal flight, the closer the center of gravity of the whole machine to the wing surface, the smaller the horizontal flight angle; if the propeller is set above the leading edge of the wing, the pressure generated by the propeller wake on the upper wing surface of the wing is Ft, and Ω is the tail pressure coefficient, when the chord length b of the wing, the radius R of the propeller, and the propeller thrust T are determined, there is: ; the eccentricity between the propeller shaft and the wing surface of the bow is E1, the eccentricity coefficient is ψ, E1=ψR, and the tail pressure coefficient Ω is: 。 8. The method of short takeoff and landing by lift dependent angle of attack according to claim 7, characterized in that: The wing machine serves as a cross-border aircraft, and its flight state can continuously transit between the rotorcraft and the fixed-wing aircraft, and sequentially passes through three stages of rotorcraft flight state, transition flight state, and fixed-wing aircraft flight state, and the degree of transition can be measured by the hang ratio ζ, which is a function of the angle of attack θ: ； The flight state of the wing-in-ground-effect vehicle continuously transitions between the rotorcraft and the fixed-wing aircraft as the wing angle of attack θ changes, and the horizontal component Vl of the wing-in-ground-effect vehicle flight speed V decreases as the angle of attack θ increases, Vl = Vbe ; ζ = 0, the wing-in-ground-effect vehicle is in the rotorcraft flight state, Vl = 0; 0 < ζ < 1, the wing-in-ground-effect vehicle is in the transition flight state, Vl = Vbe ; ζ = 1, the wing-in-ground-effect vehicle is in the fixed-wing aircraft flight state, β = 1 / K, Vl = Vbe, β > 1 / K, the wing-in-ground-effect vehicle enters the sprint flight, the sprint speed is proportional to the thrust T, when Tmcp = βmg, Vl = Vmax > Vbe, and .

9. The method of short takeoff and landing by lift dependent angle of attack according to any of claims 1 or 8, characterized in that: The wide-chord wing of the aircraft has the defect of large induced drag, and the two-dimensional maneuvering flight feature makes the pitching moment of the aircraft large, and the aircraft is difficult to control, in addition, the lift of the wing is only the angle of attack lift without the airfoil lift, so when the angle of attack is 0, the lift disappears completely, and the aircraft is prone to instability at low speed, and a sharp nose dive occurs, so some special measures are needed to achieve the purpose of stability augmentation, lift increase and drag reduction; the measures are combined with the wing to form a lifting body, such as setting a flap or a slat wing on the leading edge of the wing, or bending the leading edge of the wing, and / or adding a canard to increase the control moment, and / or forming an S-wing type with the wing and the trailing edge aileron to counter each other, and / or lowering the height of the center of gravity of the aircraft to form a vertical trim moment to improve flight stability; the slat wing set on the leading edge of the wing can effectively improve the stability, the chord length of the slat wing is less than half of the chord length of the wing, the span length of the slat wing is not greater than the span length of the wing, the longer the chord length of the slat wing or the larger the angle of attack, the greater the lift of the aircraft and the greater the head-up moment, while the form drag increases, the induced drag decreases, and the stall speed decreases; when the chord length and the angle of attack of the slat wing are constant, the faster the speed of the aircraft, the greater the lift and the greater the head-up moment; the shape of the slat wing can be rectangular, diamond, elliptical or triangular; the slat wing surface intersects with the wing surface at the leading edge of the wing, the angle between the slat wing surface above the wing surface and the wing surface is the forward rake angle, the angle between the slat wing surface below the wing surface and the wing surface is the backward sweep angle, preferably the slat wing is a plane, i.e. the forward rake angle and the backward sweep angle are complementary angles, preferably the shape of the slat wing is triangular or diamond, preferably the lower edge of the slat wing coincides with the leading edge of the wing and is inverted upward, when the diamond slat wing is used, the long side of the diamond slat wing is parallel to the wing surface, and the forward rake angle is equal to the backward sweep angle; the slat wing can also be set as a movable slat wing with variable inclination angle, when the slat wing inclination angle is turned to increase the windward area, the stall speed of the aircraft decreases, when the slat wing inclination angle is turned to reduce the windward area, the stall speed of the aircraft increases, the aerodynamic center (trim) of the aircraft remains unchanged when the symmetrical slat wing such as the diamond slat wing or the elliptical slat wing is turned; the slat wing can be divided into left and right slat wings which are turned respectively, and a symmetrical vertical stabilizer is arranged between the left and right slat wings to realize the roll control and yaw control of the wing aircraft.

10. The method of short takeoff and landing by lift dependent angle of attack according to claim 1 or 9, characterized in that: The various specific measures of the stability increasing, lift increasing and drag reducing can be combined with the wing to form a lifting body. When the aircraft is used for carrying people or has a large take-off weight, multiple lifting bodies are used in a transverse parallel, longitudinal series, and / or staggered arrangement. The lift of each lifting body remains unchanged, and the total lift increases, thereby improving the load capacity of the aircraft. The driving device of the aircraft is selected from a closed propeller such as a turbine, a semi-closed propeller such as a duct, or an open propeller such as a rotor, and the installation position is not limited. When an open propeller is used, the preferred installation position is located at the leading edge of the wing, and is arranged in a front-pull layout. In order to meet the needs of cross-border flight, the rear two-point landing gear of the front three-point landing gear of the aircraft is improved to have two adjustable heights, i.e., a high height and a low height. Increasing the height of the rear two-point landing gear reduces the wing angle of attack for taxiing and landing, and reducing the height of the rear two-point landing gear increases the wing angle of attack for short take-off and landing. Alternatively, the front three-point landing gear is expanded to a front five-point landing gear, the front three wheels are used for taxiing and landing, and the rear four wheels are used for short take-off and landing. The large-area flat wing of the aircraft can also be provided with photovoltaic cells. An lift-drag ratio greater than 10 can support efficient flight of the aircraft, and realize sunlight-powered endurance.

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