RU2742496C2 - Vertical take-off and landing aircraft - Google Patents

Abstract

FIELD: aviation.SUBSTANCE: invention relates to aviation, in particular to structures of vertical take-off and landing aircraft. The vertical take-off and landing aircraft (VTOLA) comprises a fuselage, a wing and an air propeller of a variable step, the axis of which is arranged along the longitudinal axis of the fuselage. The VTOLA also comprises a tail chassis which provides for the vertical position of the longitudinal axis of the fuselage during take-off and landing in the form of long, hingedly attached hydraulic cylinders to the fuselage. The air propeller is arranged in the nose part of the fuselage and is provided with a percussion rifle. The tail fairing, which is devoid of tail fin, is capable of remotely controlled folding and laying in flight, and the hydraulic cylinders of the chassis are capable of being removed in flight under the above-mentioned tail fairing.EFFECT: ensured is reduction in aerodynamic resistance of the aircraft and increase in speed, economy and range of the flight.3 cl, 6 dwg

Description

The invention relates to aerodromeless aviation and can be used in small aircraft to create vehicles capable of making a worthy alternative to existing light aircraft, helicopters and passenger cars.

The idea of vertical take-off and landing accompanies the entire period of aviation's existence. In this case, the goal is to create a vehicle that combines all the advantages of a car, an airplane and a helicopter, while eliminating their disadvantages. The disadvantage of a car is the need for roads and a relatively low speed of movement, for an airplane it is a need for airfields and the inability to make an emergency landing anywhere, and for a helicopter it is low efficiency of movement and a fundamentally limited speed. I would like to have a vehicle that combines high speed and economy of movement, like an airplane, with the ability to do without airfields and roads, like a helicopter, and at the same time provide sufficient security for civilian use.

Until the middle of the twentieth century, the solution to this problem was constrained by the lack of a sufficiently light engine. With the advent of a gas turbine engine, the necessary weight and energy parameters can be provided. Operating VTOL aircraft have been created, for example, HARRIER and Yak-141 with thrust from a turbojet engine, as well as VTOL aircraft with propeller thrust from a theater engine, taking off from the tail (see for example: masterok.livejornal.com/3435854/.html). However, all of them are not suitable for widespread civilian use. do not provide sufficient safety during take-off and landing. The reason for this is the use of gas-dynamic rudders to control the landing and landing and transient modes, acting from the air taken from the gas turbine propulsion engine, which can always fail.

Attempts were made to use Wankel and others rotary piston engines, which are quite light, in small aircraft. However, they do not provide an acceptable service life.

In recent years, ways have been outlined for creating ultralight piston engines using the well-proven classic piston-cylinder group with a crank mechanism. From the published information on this area of work, it is still possible to refer only to the patent of the Russian Federation No. 2645369, priority from 2015. In this regard, the development of possible options for the use of ultralight internal combustion engines in non-aerodrome aviation is underway.

The subject of the present invention is a schematic diagram of an aircraft that allows you to solve this problem. In modern aviation, schemes have been used that solve only individual problems. For example, the wing perfectly, with high efficiency, solves the problem of cruising flight, and this is used in airplanes. The main rotor does a good job of taking off and landing. However, the level flight speed of the main rotor, if used in the oblique blowing mode, is limited to a value that is approximately one quarter of the terminal peripheral rotor speed. And since the terminal speed in terms of aerodynamics and strength can only approach the speed of sound, then the maximum flight speed of a classic helicopter is about 250 km / h. In addition, since the main flight of the helicopter takes place in the mode of oblique blowing of the main rotor, the helicopter, as a vehicle, is not economical, which is due to the operation of the elements of the main rotor, on average, in non-optimal modes of flow, both in angle of attack and in speed (the angle of attack of the blade varies greatly around a circle, and the speed varies over a wide range - both around the circle of rotation and along the radius, so that the optimal value is available only in a small percentage of cases.

In aircraft convertible in flight - tiltrotors, the transient flight regime is not satisfactorily solved, i.e. the acceleration process that takes place from a helicopter takeoff to a speed gain sufficient for wing operation. Acceleration occurs horizontally, and the air flow around the vehicle rapidly changes from vertical to horizontal direction. The wing and propeller must be oriented with the flow, keeping track of its angle. In this case, it is possible to reorient the entire apparatus, or only its supporting part - the wing and the propeller, without changing the orientation of the fuselage. But this raises two problems. Firstly, the overclocking process is delayed, because at the initial stage, only a small horizontal component of the propeller thrust is responsible for acceleration. Second, acceleration occurs near the ground, making it problematic to ensure safety in the event of engine, support, or control system failure. At the same time, there is a dangerous combination of low altitude and not high enough speed. Thirdly, the successful reorientation of the tiltrotor bearing systems is hindered by the large inertia of the moving parts of the vehicle and the rapid change in the characteristics of the controls.

For the above reasons, the most preferable for solving the task at hand are vehicles with an airplane layout, which is not substantially transformable in terms of bearing systems and in terms of operating modes. These are schemes of VTOL aircraft taking off from the tail. This scheme is applicable only for small devices. However, this fully satisfies the needs of independent movement, as proven by the practice of motorization, where passenger cars are most developed.

Closest to the proposed version of the scheme of an aerodromeless aircraft is VTOL according to US Pat. RF No. 2430859, priority from 2008. It contains a fuselage with a small wing located on it, and also contains a tail landing gear that provides vertical orientation of the longitudinal axis of the fuselage in take-off and landing modes. In the tail section of the fuselage, a variable-pitch pushing propeller is installed in the casing, as well as a cruciform (bomb) stabilizer with rudders that provide control over all three channels - roll, course and pitch. Safety at all stages of flight is ensured by a small orienting parachute located in the nose of the fuselage, as well as by the fact that the tail landing gear is made in the form of long telescopic hydraulic cylinders that provide parachuting speed damping with a sufficiently wide landing base during wind drift. At the same time, in cruising mode, the hydraulic cylinders are oriented along the stream, thereby minimizing their aerodynamic resistance, which is necessary to achieve high flight speeds.

In the process of further development of this VTOL aircraft scheme, the following disadvantages were identified:

1. Influence of the proximity of the ground on the characteristics of the rudders located in the tail, which creates a large non-stationary interference in the heading, pitch and roll control channels in the first second of takeoff and in the last seconds of landing. This complicates control not only in manual but also in automatic mode.

2. Increased aerodynamic drag created by the propeller casing, casing attachment pylons, chassis hydraulic cylinders attachment brackets, as well as the aircraft's cruciform stabilizer at high flight speeds. Moreover, even the landing gear hydraulic cylinders located along the direction of flight give great resistance, because the direction of flow is different near the fuselage surface.

3. Limited possibilities to increase the screw diameter.

The listed disadvantages of the known scheme increase the required engine power, reduce transport efficiency, flight speed and range, and complicate safety during takeoff and landing.

The aim of the invention is to eliminate the above disadvantages of the known scheme of VTOL aircraft.

A VTOL aircraft is proposed, containing a fuselage, as well as a wing and a variable-pitch propeller fixed on it, the axis of which is located along the longitudinal axis of the fuselage. There is also a tail landing gear, which provides a vertical position of the longitudinal axis of the fuselage during take-off and landing, as well as a system serving to mitigate the impact during an emergency landing, made in the form of long hydraulic cylinders pivotally attached to the fuselage, connected at the ends by stretch marks. The purpose of the invention is achieved in that the specified propeller is located in the nose of the fuselage and is equipped with a swash plate, and the specified hydraulic cylinders of the tail landing gear are placed in a tapered tail fairing, made with the possibility of operational folding and unfolding in flight.

The specified implementation of the VTOL aircraft allows you to increase the diameter of the propeller with a corresponding decrease in the required power for takeoff and landing. The swashplate, mounted on a propeller with an increased diameter, allows the aircraft to be controlled in pitch and heading, freeing the tail of the aircraft from this function. In addition, due to the forward position of the propeller, the roll control can be assigned to the ailerons of the wing, which is blown in this case by the propeller. Thus, the tail can be made with a pointed fairing, without any rudders and stabilizers. The tail landing gear cylinders can be retracted under the fuselage tail fairing. This reduces the aerodynamic drag of the non-bearing elements of the aircraft, which, at high flight speeds, constitutes the majority of the total drag of the aircraft. Thus, an increase in the aerodynamic perfection of the aircraft is provided, which makes it possible to increase the flight speed and increase the efficiency of flight, the required engine power for vertical take-off and landing decreases, and the payload and maximum flight range increase.

In a particular embodiment of the proposed VTOL aircraft, the tail fairing is made in the form of a flexible cone-shaped inflatable shell. In this case, the possibility of deploying the tail fairing before landing and folding after take-off is ensured by the fact that the flexible shell of the tail fairing is provided with cuts located along the generatrix of the cone and a mechanism for their tight automatic connection and disconnection.

In a particular embodiment, the specified mechanism for tightly connecting and separating the edges of the shell contains teeth located on one of the joined edges of the shell, and on the other - corresponding depressions. Moreover, the teeth are provided on the mating surfaces with through holes with chamfers, and the axes of the holes, when connecting the edges of the shell, coincide with each other, forming a channel to which a connecting element can move, made, for example, in the form of a cable with a rounded end connected to a fixed base drum, providing the ability to push the cable into the channel and pull it out. When the cable is pushed into the channel, the connecting edges of the shell will be connected under the action of the push-in rope, the end of which self-centers when pushing in, resting on the chamfers of the holes in the teeth. Thus, it is possible to remotely connect and disconnect the edges of the shell, i.e. fastening and unbuttoning as with a zipper. However, in comparison with "lightning", here, in addition to the possibility of automating the process, high strength and reliability of the connection is provided, according to the principle of the so-called multi-lug connection. The joint is also sufficiently sealed by two continuous edges running parallel along the entire joint. The said edges also provide rotation orientation of the shells to be connected, which is necessary to obtain a smooth outer surface of the fairing without protrusions and depressions.

The invention is illustrated by the following description of structural examples and six figures.

FIG. 1 shows the proposed VTOL aircraft in flight configuration.

FIG. 2 shows a nose view in flight configuration.

FIG. 3 shows the proposed VTOL aircraft in section in a configuration corresponding to parking on the ground.

The single-dot dash-and-dot line shows the location of the landing gear cylinders inside the tail fairing.

The dash-and-dot line with two dots shows the position of the hydraulic cylinders before landing or immediately after takeoff.

The dashed line shows the maximum possible deflection of the hydraulic cylinders achieved after an emergency landing without an engine.

FIG. 4 shows the design of the mechanism for connecting and disconnecting the edges of the flexible inflatable shell of the tail fairing.

FIG. 5 is a sectional view taken along the plane A-A shown in FIG. 4.

FIG. 6 shows the trajectory of the full flight cycle of the proposed VTOL aircraft with vector diagrams of forces acting on the vehicle in each of the characteristic phases of the flight.

T is the propeller thrust force. Fy is the lift of the wing. Fx is the total aerodynamic drag force. P is the force of the weight of the apparatus.

The proposed VTOL aircraft contains a fuselage 1, which has a drop-like shape with a tapered tapered tail section. Wing 2 is equipped with deflectable surfaces 3 and 4 with independent two-channel control, which makes it possible for them to simultaneously combine two functions - aileron function and flap function. In the forward part of the fuselage, an air propeller 5 is installed with a swashplate 7 located in its sleeve 6, which provides the ability to controllably change the total pitch of the propeller, as well as the so-called cyclic pitch of the propeller, as is the case with the main rotor of a helicopter. In principle, the swashplate device is no different from a helicopter one. The propeller blades 5 have a twist that is optimized for cruising flight speed.

The screw 5 is connected by means of a gearbox to the engine 8. It can be either a turboshaft engine or a piston internal combustion engine. Optimum is the use of a radial piston-type liquid-cooled internal combustion engine with the use of elements of the outer skin of the fuselage as a radiator, so as not to impair the streamlining of the fuselage. The part of the fuselage fairing located in front of the propeller is rigidly connected to the main part of the fuselage and it houses the parachute 9 (Fig. 3), which serves to decelerate and orient the vehicle in case of engine failure. Parachute 9 should not be made large, so as not to increase the wind drift angle of the vehicle during an emergency landing, which is fraught with the danger of capsizing. The emergency landing speed of the vehicle must be at least 20 m / s. The corresponding parachute area for a two-seater VTOL aircraft will be 10 square meters.

In the tail section of the VTOL aircraft there is a chassis made in the form of long hydraulic cylinders 10 pivotally attached to the power bottom 11 of the fuselage (Fig. 3). The lower ends of the rods of the hydraulic cylinders are connected to each other by power braces 12 of constant length, which are located radially when parked on the ground. In flight, the hydraulic cylinders 10 are retracted under the tail fairing 13 (see dot-and-dash with one point). To move and extend the hydraulic cylinders 10, they are equipped with servo drives 14. To ensure compact folding of the guy wires 12, auxiliary cables 15 are attached to them, connected to non-power uncontrolled winches 16, made, for example, with spring motors. When retracting the hydraulic cylinders under the fairing, winches 16 pull the power extensions 12 upward, eliminating their sagging.

The tail fairing of the fuselage is made in the form of a conical flexible inflatable shell 17 (Fig. 1, Fig. 3). Along the generatrix of the cone, opposite the location of the respective hydraulic cylinders 10, the shell 17 has cuts along the generatrix of the cone, extending from the base of the cone to the top, where the shell is attached to the braces 12 at their intersection (Fig. 3). At the edges of the cuts there are teeth 18 (see Fig. 4), the location of which corresponds to the depressions on the counterpart of the cut. The teeth 18 on the mating surfaces have chamfered holes 19. These holes, when tightly closed with each other, form a continuous channel, in which a flexible connecting element 20, made, for example, in the form of a cable with a rounded end, is disposed for movement along the channel. The flexible connecting element 20 is wound on a drum 21 with a pinch roller 22. Both the roller and the drum are provided with a rotation drive. The teeth 18 are made in one piece with the sealing strips 23, which also perform the function of orienting the teeth in rotation relative to the axis of the connecting element 20. This makes it possible to make the outer surface of the connection without protrusions and depressions, which is necessary for streamlining (see Fig. 5). This mechanism for connecting and disconnecting the edges of the casing cut provides high strength and reliability of the connection, as well as the ability to remotely control the process of opening and closing the casing cuts in flight (as opposed to zipper locks).

Under the bottom 11, in the center, there is an additional emergency landing braking system. It takes effect in the last braking section, when the hydraulic cylinders have diverged more than 30 degrees, at which their action becomes ineffective. In the embodiment shown in FIG. 3, the additional system is designed as a bellows 24 equipped with an automatic pressure limiting valve.

The action of the proposed VTOL aircraft differs from the action of a conventional aircraft, which has a thrust-to-weight ratio of more than one, in the method of creating control and stabilizing moments of force, which is due to the absence of a tail unit or the optionality of its presence or use in the proposed apparatus. In all standard flight modes, the vehicle operates in a flow close to longitudinal, i.e. with wing attack angles not exceeding critical (less than 14 degrees), and sliding angles close to zero. The task of piloting is to continuously maintain this condition. The difference from piloting a conventional aircraft is that the speed of the flow around the aircraft, in its change in absolute value, can decrease to zero. At the same time, an ordinary plane crashes because in an effort to resist gravity, the pilot is forced to move the wing to supercritical angles of attack in order to avoid acceleration towards the ground. In the proposed aircraft, which has a thrust reserve up to the weight of the aircraft, it may not do this, but simply add thrust, keeping the wing at the working angles of attack. In this case, the movement of the vehicle in space relative to the ground is determined by the vector sum of four forces: the force of gravity of the earth, the thrust force of the propeller directed along the longitudinal axis of the aircraft, the vector of the lifting force of the wing, directed almost perpendicular to the plane of the wing, and the vector of the total force of the aerodynamic drag of the vehicle, directed along the axis fuselage back. If the rotor thrust can be close to the magnitude of the gravity in absolute value, then the equilibrium of the apparatus in the air can be achieved at an arbitrarily small value of the wing lift. It is enough only to orient the apparatus vertically in relation to the ground. The wing at supercritically low flight speeds can be used for control needs to the extent of the remaining bearing capacity of the wing. However, in practice, the speed of axial flow around the wing in this case never reaches zero, because with a decrease in flight speed, the inductive speed of the rotor increases, reaching 20-30 m / s in hover mode. This is enough to compensate for the reactive torque of the engine applied from the propeller gearbox to the fuselage due to the ailerons 3 and 4 of the wing. So we get the balance of the apparatus in the mode of motionless hovering relative to the air. But this balance is unstable. A small change in any of these forces will lead to a deviation of the resulting force from zero, and there will be an acceleration of the vehicle caused by this force in the corresponding direction. The integral of this force over time will show the accumulation of speed, and the integral of the speed will show the accumulation of displacement. It is possible to counteract this process and keep the deviation of the apparatus along the movement within acceptable limits by having three control channels. If we use aircraft representations, then it is necessary to control the moments in pitch and heading, as well as to control the amount of propeller thrust. In an aircraft, the pitch moment is created by the tailplane's elevator, and the heading moment is created by the keel. However, at low flight speeds, the action of the rudders on the tail unit may not be effective enough. This is especially true in the case of the proximity of the ground, in particular during takeoff and landing. It is precisely because of the insufficient efficiency of the rudders of the tail assembly and in the absence of the possibility of controlling the thrust vector of the power plant, the aircraft can fall down and enter a spin at low flight speed. A large thrust-to-weight ratio does not always help to avoid this. In the VTOL aircraft, taken as a prototype, the blowing of the tail unit by the propeller located in the tail helps out. However, this does not work well near the ground. In the proposed apparatus, the pitch and heading control moments in all modes are equally effectively created by the swashplate. If the propeller blades are not hinged, then the action of the swashplate is reduced to a parallel displacement of the thrust vector relative to the center. In the case of hinged attachment of the blades, a cyclical swing deflection of the blade occurs with a corresponding blockage of the propeller throwing cone and, as a consequence, an angular deviation of the action of the total centrifugal force of the blades on the sleeve occurs with the appearance of a shoulder of this force relative to the center of inertia of the apparatus. In both cases, a pitching or yawing moment occurs, which, in combination with the control of the total propeller pitch, i.e. thrust can be used for full control of the vehicle both in hover mode, where control is no different from helicopter mode, and in airplane mode, where the same swashplate will operate no less effectively than the tail rudders. So, the tail unit turns out to be unnecessary from it, in order to reduce aerodynamic drag, which very critically affects the economic performance of flight at high speeds, should be abandoned, leaving only one ideal fuselage, the aerodynamics of which, in the case of the proposed VTOL aircraft, even the landing gear is not damaged, which completely retracts under the common fuselage fairing.

Some difference between piloting in horizontal flight from an airplane flight may consist in the fact that a propeller with an increased diameter creates perceptible gyroscopic forces, due to which the moment of force in pitch causes not acceleration in pitch, but angular velocity along the course, and, conversely, the moment of yaw causes the appearance of the rate of change pitch angle. Those. there is cross-talk along these control channels. However, in the case of the hinged fastening of the propeller blades and the presence of the flapping motion of the blades, the gyroscopic forces, as in a helicopter, are perceived by the air and are little transmitted to the fuselage. The positive influence of gyroscopic forces consists in stabilizing the direction of the longitudinal axis of the vehicle, independent of aerodynamic forces. This compensates for the loss of restoring forces present in the aerodynamic stabilizer when properly centered. Moreover, even an apparatus with instability of control along the course and pitch, at present, with the widespread use of autopilots, is quite applicable. However, in the case of the proposed apparatus, we are not dealing with instability, but only with a small value of the aerodynamic component of the stabilization forces. A small stabilizing force, in particular, can take place due to the installation of ailerons 3 and 4 at a small negative angle of attack, as is the case with tailless aircraft. However, this is not necessary in this unit.

Roll control, as already mentioned, is carried out by ailerons 3 and 4. That is, it remains conventional aerodynamic. However, due to the fact that the wing is located closer to the propeller blowing it than to the ground, and also due to the large shoulder of the ailerons, their effectiveness remains sufficient for all modes of flight with the engine.

In case of engine failure, parachute 9 should be immediately released, which will extinguish the residual airspeed of the vehicle to a small value, prevent the rotation of the vehicle and orient its longitudinal axis vertically.

Let's consider all stages of the flight cycle in sequence. The pre-launch configuration of the apparatus is shown in FIG. 3. Seats 25 and 26 are mounted on a hinge, which allows the crew to comfortably sit on and descend on the ground. For takeoff, both seats are overturned and then the doors 27 and 28 are closed. The engine is started at the set zero general and zero cyclic pitch of the propeller 5. Then the mode of automatic stabilization of the engine speed, as well as stabilization of the roll, course and pitch angles (in the coordinate system, associated with the aircraft). Then, the total pitch of the screw is quickly introduced until a thrust is obtained that exceeds the weight of the apparatus (for example, 1.25 times). In this case, the apparatus begins to accelerate upward with an acceleration equal to the difference between the weight force and the thrust force (for example, 2.5 m / s per second) (position "a" in Fig. 6). As the vertical flight speed increases, the pilot activates the wing, while controlling the pitch, as usual, but using the moments created by the swashplate instead of the tail unit. In this case, the angles of attack of the flow around the wing profile are kept within subcritical limits. Smoothly increasing flight speed causes a smooth increase in wing lift, and the flight path will tilt smoothly. It will tend to the horizontal (positions "b" and "c"). In this case, you can smoothly reduce the propeller thrust, and, at a certain speed, when the lift of the wing becomes equal to the weight, the flight becomes horizontal - airplane flight (position "d"). In order for horizontal flight to occur earlier, flaps must be introduced in advance, which, in this embodiment, are aligned with ailerons 3 and 4. That is, add a constant component to the differential aileron deflection angle. The decoupling of the aileron and flap channels is realized by the autopilot. Man controls only the master. When approaching a destination, reverse the flight path slope sequence. For this, the flight speed should be smoothly reduced to zero, without going beyond the working angles of attack. In this case, the flight angles will also change in the opposite direction than during takeoff. Those. the aircraft will pitch up, gaining altitude (position "f"). The difference from the takeoff regime here is that the thrust is first reduced and then introduced only within the weight of the vehicle. When the airspeed decreases to zero, the helicopter hover mode will be established (“g” position). Then it will blow off to make a decrease in hover mode. At the same time, it will blow off to correct the position of the apparatus relative to the ground to reach the specified landing point. This should be done at the expense of low horizontal speeds, which do not bring the propeller into the modes of a strong bevel of the blowing (especially if the blades do not have fly joints) (position "h"). To reduce the altitude of the helicopter descent and reduce the time spent on descent, and not to get into the vortex ring mode, the initial altitude of the pre-landing evolution should be minimized by approaching the destination in the airplane mode with descent (position "e"). This will save fuel, which is consumed faster in helicopter mode. In the helicopter mode of lowering, the sections of the shell of the tail fairing of the fuselage are opened and the hydraulic cylinders 10 of the chassis are opened using the servo drives 14.

If the engine, control system or load-bearing elements fail at a flight altitude above 20 meters, parachute 9 is thrown out, which brakes the vehicle relative to the air and orientates it to the vertical landing position. In this case, the control of the landing site is possible with the help of parachute 9, if there is a sufficient headroom.

If the equipment fails at an altitude of less than 20 meters, then the device is already in the takeoff and landing configuration, and the parachute is not required, because the speed of falling from this height is within the limits of its safe braking with the help of hydraulic cylinders 10, the braking stroke of which, together with the stroke of pneumatic bellows 24, is 2 m.

Flexible tail cone shells can be inflated using a high-velocity head taken from the forward fuselage. This pressure will be sufficient to avoid the appearance of aeroelastic vibrations of the shell, i.e. flutter plating. This is facilitated by the reinforcement of the casing with a high modulus cord.

Below are the estimated estimates of the technical parameters of the proposed VTOL aircraft, achieved in the above example of its implementation.

The engine is a piston radial two-stroke eight-cylinder with a turbofan over-purge and gasoline injection into the cylinder.

The maximum engine power is 200 kW.

Working volume - 2.5 liters.

Rated power 110 kW at 60 r / s.

Specific fuel consumption 0.3 kg / kWh.

The mass of the engine structure is 40 kg.

The diameter of the cylinder block (according to the bases of the candles) is 550 mm. (The engine parameters are obtained by scaling the parameters of the prototype of a slightly lower power).

The takeoff weight of the VTOL aircraft is 300 kg.

Number of seats - 2.

Payload ratio 0.5.

Screw diameter 2 m.

The blade chord for the two-bladed version is 0.18 m.

The thrust-to-weight ratio is 1.25.

The maximum propeller thrust is 3750 Newtons.

The maximum acceleration up on takeoff is 2.5 m / s sq.

Power on the propeller shaft at the beginning of vertical take-off with an acceleration of 2.5 m / s sq. with an air density of 1.2 kg / m3 and the efficiency of the propeller in hovering mode 60% - 160 kW.

Inductive flow velocity in the plane of the propeller in hovering mode - 25 m / s.

The minimum level flight speed with flaps is 40 m / s = 145 km / h.

The climb time for horizontal flight in vertical acceleration mode is 16 sec.

Fuel consumption in vertical take-off mode - 45 kg / h = 12.5 g / sec.

Fuel consumption for vertical take-off without smooth activation of the wing - 200 g.

The approximate total fuel consumption for takeoff and landing along the optimal glide path (i.e. with pitching up from the lowest possible height) is 0.5 kg.

Emergency parachuting speed - 20 m / sec.

The required area of the orienteering parachute with the Cx parachute is 1.25 - 10 square meters.

G-force during emergency landing without engine - 10 G.

Wingspan - 3 m.

Wing area - 1.6 sq. M.

Wing drag profile - Cx cr. - 0.01

The midsection of the fuselage is 0.8 square meters.

Fuselage Cx - 0, 07.

Taking an air density of 1.2 and a flight speed of 140 m / s = 500 km / h, we get:

The inductive speed of the wing bevel is 0.6 m / s.

The corresponding power of the wing inductive resistance is 1.8 kW.

The power of the wing profile resistance at 500 km / h is 27 kW.

The aerodynamic drag power of the fuselage is 94 kW.

Taking the interference of the wing with the fuselage to be zero for the midplane, we obtain the total power of the aerodynamic drag of flight at a speed of 500 km / h and an air density of 1.2 kg / m3:

1.8 + 27 + 94 = 123 kW.

If the twist of the propeller blades is optimized by 500 km / h, then the efficiency of the propeller can be 73%.

In this case, the shaft power will be 170 kW.

The corresponding fuel consumption per hour is 50 kg / h.

Specific fuel consumption - 10 kg / 100 km.

The flight range corresponding to a 30 kg of fuel on board will be 300 km.

However, if the flight altitude is increased to 7000 m, where the air density, and hence the aerodynamic drag, is half as much, then the specific fuel consumption will decrease to 5 kg per 100 km, and the range will increase to 600 km.

If in addition to this, the flight speed is halved, i.e. up to 250 km / h, then the specific fuel consumption will decrease by 4 times - to 1.25 kg per 100 km, and the range will increase to 1200 km. In these estimates, we can neglect the increase in the power of the inductive resistance of the wing, since it is, as stated above, very small. In addition, the increase in the wing profile Cx can also be ignored, since wing with an area of 1.6 sq. m. taken with a large margin of bearing capacity (the velocity head at the ground reaches 12000 Pa).

In view of the above, it is quite realistic to bring the flight altitude up to 10,000 m, where the air density is 3 times less than on the ground. At the same time, the specific fuel consumption will also decrease by 3 times, compared to the terrestrial, and will amount (at a speed of 500 km / h) 3.3 kg per 100 km, and the range will increase to 900 km.

Accordingly, when the flight speed is reduced to 250 km / h, the specific fuel consumption at 10,000 meters will decrease by another 4 times - to 0.83 kg per 100 km, and the flight range, with 30 kg of fuel on board, will increase to 3600 km ...

Thus, as a vehicle, VTOL aircraft are not only able to reduce the cost of road maintenance, but can have a significant superiority over a car, both in speed and in economy. In terms of safety, airplanes are already safer than cars. This is facilitated, in particular, by the more favorable opportunities for air traffic automation compared to ground traffic.

Claims (3)

1. A vertical take-off and landing aircraft (VTOL) containing a fuselage, a wing and a variable-pitch propeller attached to it, the axis of which is located along the longitudinal axis of the fuselage, and also containing a tail landing gear, providing a vertical arrangement of the longitudinal axis of the fuselage during takeoff and landing, made in the form of long hydraulic cylinders pivotally attached to the fuselage, connected at the ends with each other by stretch marks, characterized in that the specified propeller is located in the nose of the fuselage and is equipped with a swash plate, the tail fairing of the fuselage is made with the possibility of remotely controlled folding and unfolding in flight, and the landing gear hydraulic cylinders are made with the possibility of retraction in flight under the specified tail fairing. 2. VTOL aircraft according to claim 1, characterized in that said tail fairing is made in the form of a flexible inflatable shell with slots along generatrices located opposite to the corresponding hydraulic cylinders and is equipped with a mechanism for remotely controlled opening and connection of the edges of the slots. 3. VTOL aircraft according to claim 2, characterized in that said mechanism for opening and connecting the edges of the slots of the flexible inflatable shell of the tail fairing contains teeth located on one of the connected edges of the shell, and on the other - corresponding depressions, and the teeth are provided on the mating surfaces with through holes with chamfers, and the axes of the holes when connecting the edges of the shell coincide with each other, forming a channel through which a flexible connecting element connected to the drive drum can move.

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