WO2016084000A1

WO2016084000A1 - Safety system for controlling the attitude of aircrafts

Info

Publication number: WO2016084000A1
Application number: PCT/IB2015/059085
Authority: WO
Inventors: Simone Bianchi; Andrea ANDOLLINA
Original assignee: Simone Bianchi; Andollina Andrea
Priority date: 2014-11-26
Filing date: 2015-11-24
Publication date: 2016-06-02

Abstract

An auxiliary safety system for controlling the attitude of an aircraft (1) provides sensors placed on the aircraft (1) which identify stalling and/or other critical flying conditions that are imminent and/or have occurred, an electronic control unit (101, 201) connected to the sensors (106, 107, 108, 206, 207, 208), and propellers placed on the aircraft each suitable to generate a continuous thrust; the electronic control unit (101, 201) activates the propellers, in case the above mentioned conditions occur, in order to generate a continuous thrust for all the duration of the attitude correction and be able to recover the proper flight attitude of the aircraft. The system is reliable and easily controllable and has a low energy consumption.

Description

SAFETY SYSTEM FOR CONTROLLING THE ATTITUDE OF AIRCRAFTS

BACKGROUND OF THE INVENTION

It is an object of the present invention a safety system for controlling the attitude of aircrafts, in particular being able to identify stalling and/or other critical flying conditions that are imminent and/or have occurred.

KNOWWN ART

The lift is defined as the aerodynamic force acting in a direction perpendicular to the wind direction hitting a profile and allows an aircraft to take off. In other words, the lift is equivalent to the pressure difference existing between the lower and upper part of an aircraft. To the lift generation all the aircraft contributes but the higher contribution is given by the wing.

The generic airfoil shape allows obtaining a difference in the fluid velocity between the upper and lower part of the wing. Without going into details, but recalling the Bernoulli principle establishing a relationship between pressure and square of fluid velocity, in particular establishing that a velocity reduction causes a pressure increase, it is possible to say that on the lower part there is a higher pressure than the upper part. As mentioned, that pressure difference is called lift.

The lift is calculated by the general formula:

L =→V²SC_L

2

where P is the air density, V is the flight velocity, S is the reference surface (in case of aircrafts is the wing surface) and ^Lis a dimensionless coefficient being called lift coefficient.

The lift coefficient varies as a function of different parameters among which the wing geometrical shape and the angle of attack. In particular, for angles of attack lower than the stall angle it is possible to express the lift coefficient as:

C_L ~ C_L\a ⁽X

where ^a specifies the airfoil incidence (the angle that the profile chord forms with the unperturbed wind direction hitting it).

In fluid dynamics the stall is a reduction of the lift coefficient due to an increase of the incidence angle or to the reduction of the incidence velocity on an aerodynamic profile, such as for example an airfoil. The incidence angle minimum value for which the phenomenon appears is called critical incidence angle; that value, corresponding to the maximum lift coefficient, varies significantly as a function of the particular profile. For example, a mean value is around 15°.

The stall can be symmetrical or asymmetrical: in the first case, the aircraft loses altitude pitching in a nose-dive; in the second case, the aircraft loses altitude in a spin.

The aerodynamic stall happens when the flow separation point on the top, (that is where the current detaches from the top of the airfoil due to an incidence increase), is advancing with respect to the current direction, up to the point where the flow is separated almost over the entire wing top.

The stall can happen at limited speed but also at high speed (the high-speed stall, in the pilot jargon, is also called "power stall"). In case of very abrupt maneuvers, the angle of attack variation of the airfoil can vary too quickly to allow the limit layer adherence, thus exceeding the critical incidence. Differently from the low speed stall, the high-speed stall is more dangerous from a structural perspective due to how quickly the permitted limit load factors (n) of the aircraft structure can be exceeded.

Particularly interesting and critical it is the case where the stall happens at the wing tips because it involves the movable surfaces such as the ailerons; the stall never happens symmetrically on the two half-wings and thus the lift difference being generated between the already stalled half-wing and the other one still lifting produces a roll momentum leading to the so-called spin, which sometimes is emphasized by particular drift configurations: when to an asymmetrical stall it is added a rotation along the longitudinal axis of the aircraft (yaw), the airplane develops a self-sustained rotation with a resulting altitude lose.

In the past, the problem has been faced in several ways.

In particular, in the US patent document US2584298 an electro-mechanical mechanism is described which, once verified the stall condition are occurring, acts directly on the stabilizer in order to recover a proper flight velocity that allows the aircraft exiting the stall.

In the US patent document US2923499 an electro-mechanical mechanism is described which acts actively on the pilot controls, typically the stick, in order to regulate the stabilizer excursion, avoiding the occurrence of stall situations at the same time.

Beside the already mentioned documents, which are referring to problem solutions directly acting on the mobile control surfaces of the aircraft, other ones are highlighted, which describe solution mechanisms directly acting on the flow.

Actually, the US patent document US3149804 provides the use of an instrument, which, once the stall condition is occurred, provides to expel air on the wing top through distributed holes, at which diffusors are placed being made as nozzles. This allows extending the chord percentage where the airflow is attached to the wing itself.

Similarly to the preceding one, the US patent document US3237889 also provides to blow air on the wing by means of a continuous diffusor running near the wing leading and trailing edges.

Since the correction maneuvers are associated with the efficiency of the mobile control surfaces, the above-mentioned solutions could be non- satisfactory in that the control surfaces efficiency is strongly compromised in the stall conditions.

Another used system, not directly depending on the control surfaces, is the anti-spin parachute. However, it is not properly designed to recover the aircraft from symmetrical stall conditions and it is single-use. Moreover, since the parachute can also be ripped and/or damaged during use, it is not easily reusable. Its activation is not exact, with respect to the above-described situations, and once activated it is not controllable. The Russian patent document RU2368541 refers to a system including pulsejet engines (which are jet engines) placed on the fuselage. That solution causes reliability, control, noise and high fuel consumption problems.

The US patent document US 2009/0261206 concerns a system to control an undesired flow separation. The system provides one or more micro jets being placed in order to supply an auxiliary fluid flow in an area where a fluid separation is suspected to exist. This system is complex and not easily predictable and manageable.

PURPOSE OF THE INVENTION

The purpose of the present invention is to propose a system being able to obviate the above-described drawbacks.

BRIEF DESCRIPTION OF THE INVENTION

This purpose is obtained by means of a system according to claim 1.

BRIEF DESCRIPTION OF THE DRAWINGS In order to better understand the invention, in the following two exemplary and non- limiting embodiments thereof are described, which are shown in the attached drawings where:

- Figures 1, 2 and 3 are top, lateral and front views of an aircraft provided with a system according to a first embodiment of the invention, respectively;

- Figures 4, 5, 6 and 7 show particulars of the aircraft provided with the system according to the above-mentioned first embodiment;

- Figure 8 is a lateral view of an aircraft provided with a system according to a second embodiment of the invention;

- Figures 9-10- 11 show particulars of the aircraft according to the above- mentioned second embodiment;

- Figures 12 and 13 show a diagram of the system according to the first embodiment and a diagram of the system according to the second embodiment, respectively;

- Figure 14 shows the block diagram of the system activation logic according to the two embodiments of the invention.

DETAILED DESCRIPTION OF THE INVENTION

The two shown embodiments show a safety system to control the attitude of an aircraft 1.

The proposed system has the purpose of allowing recovering from critical and potentially catastrophic attitudes of an aircraft 1. The proposed system generates continuous correction thrusts, useful to recover a proper flight attitude, by means of compressed gas or monopropellant rocket engines.

Irrespective of the considered activation type, it is a purpose of the system to activate automatically when the aircraft 1 enters a critical flight condition. That event very often corresponds to the "stall" (symmetrical stall) or "spin" (losing altitude while spinning following an asymmetrical stall) condition.

If one of the two above-mentioned critical situations (or other critical flight conditions) has been just started or is going to start, then the system detects such a critical situation and activates, thus allowing the aircraft 1 to quickly recover (in few seconds) a proper flight attitude. The system comprises sensor means placed on the aircraft 1 (e.g. temperature, pressure, Mach, etc.) controlling the general attitude of the aircraft, in particular the orientation of the aircraft with respect to the horizon and the movement direction, and detecting stalling and/or other critical flying conditions that are imminent and/or have occurred (the sensor means will be cited in the following of the description).

The system comprises an electronic control unit 101 , 201, which can be independent from other ones being already installed on board of the aircraft 1.

That electronic control unit 101, 201 is connected to the sensor means to detect the above-mentioned stalling and/or other critical flying conditions that are imminent and/or have occurred. The system also provides propulsion means located on the aircraft 1, being able to generate one or more continuous forces acting directly on the aircraft in order to create one or more correction momenta with respect to the center of gravity of the aircraft in order to bring the aircraft back in the proper flight attitude. The control unit 101, 201 activates the propulsion means in the above-mentioned cases in order to generate a continuous thrust for the entire duration of the correction of the attitude and recover the proper flight attitude of the aircraft 1. Those propulsion means are different from the main propulsion means of the aircraft 1. Those propulsion means are intended to carry out functions for recovering the proper flight attitude of the aircraft 1. The attitude control system is independent from the main propulsion unit(s) of the aircraft 1.

As mentioned, the proposed system can be realized using different engine types among which, by way of example, the two embodiments being related to a compressed gas rocket engine and a monopropellant-gas rocket engine respectively will be described.

In a first embodiment, thus the system consists of an electro-pneumatic mechanism automatically activating when the aircraft 1 on which it is installed is flying in an attitude condition being detected as critical.

The system is shown in figures 1 to 7 and 12. In this embodiment, the system is mainly made of a highly compressed gas contained in a pair of tanks 102, a set of pneumatic connection pipes 105, a pair of nozzles 104, each connected to a respective tank 102 and the electronic control unit 101.

In case the chosen gas is air, the tank 102 can contain the compressed air quantity being necessary to continuously operate the system since the beginning or it can be pressurized by a compressor 112 activating to compress other air being tapped from outside through dedicated inlets.

When the air in a tank 102 is almost finished, the valves 103 allow the tank 102 being pressurized again. During the compressor 112 operation, the right valve 103 in fig.12 is closed in order to avoid the air flowing back towards the compressor 112, while the left valve 103 is open in order to connect the compressor 112 to the tank 102. When the compressor 112 is not operating, the valves 103 are closed in order to isolate the tank. If the auxiliary safety system must be activated, the right valve 103 is opened. The whole is managed by the electronic control unit 101, which acts based on the information provided by dedicated sensors and commands the valves 103 by means of electrical wirings 110 of the valves 103.

As mentioned, when a dangerous flight condition is detected, the system activates in order to bring the aircraft 1 back in a safe flight condition. The electronic control unit

101 operates in order to release part of or all the compressed gas contained in the tanks 102. Flowing through the pneumatic connection pipes 105, it reaches the nozzles 104 where it expands. The gas expansion increases the velocity thereof: in turn, the generated acceleration generates the continuous propelling force, which operates the correction maneuver (generally affecting the 6 degrees of freedom of the aircraft 1) being suitable to bring the aircraft 1 back in the proper flight attitude. In the second embodiment whit actuation by monopropellant rocket engine, as shown in figures 8 to 11 and 13, the compressed gas (a gas having monopropellant proprieties) is stowed in a pair of pressurized tanks 202. When the electronic control unit 201 activates the system, the control means, usually valves 203, are activated in order to allow the gas flowing into a pair of combustion chambers 212. The electronic control unit 201 is connected to the valves 203 by means of the electrical wirings 210 of the valves. Inside each combustion chamber 212 there is a catalyst element (whose selection essentially depends on the gas type being used), which, when contacting the monopropellant gas, divides the latter in a combustible gas and oxidant gas mixture. Those gases react producing a gas mixture having high temperature and high pressure, which by flowing through the pneumatic connection pipes 205 reaches the nozzles 204 where it expands and is accelerated to allow generating the desired correction thrust.

The actuation control of the system follows the logic being previously described in the diagram for both the actuation cases.

It should be underlined how the electronic control unit 101, 201 can not only take advantage of the information received by the on-board systems 111, 211 of the aircraft 1 such as the FCS (Flight Control System) but also take advantage of data coming from additional sensors 108, 208 of different kinds (e.g. pressure, temperature) located in suitable positions and communicating with the control unit through dedicated electrical wirings 109, 209.

Further sensors 106, 107, 206, 207 cooperate with this kind of sensors, which account for the proper system operation, for example, by monitoring the operation of some elements of the system itself such as the control members (the valves 103, 203) or the tanks 102, 202.

The system activation logic is the one shown in the block diagram of figure 14. The detailed description of the blocks follows. The block A represents the acquisition of the quantities characterizing the flight of the aircraft 1 at every time the aircraft 1 is flying (e.g.: flight velocity, altitudes, angles, angular velocities, angular roll, pitch and jaw accelerations) from the Flight Control System FCS (and/or from other preexisting systems on the aircraft 1) or from dedicated units, and the acquisition of any other quantity being useful to decide the danger condition: (e.g.: pressure distribution along the wing/fuselage 100, 200 etc.).

In particular, the electronic control unit 101, 201 contains in a memory all the information related to the inertial and geometrical design parameters of the aircraft 1 on which it is placed. For example, the moments of inertia with respect to the body- axes, the mass, the center of gravity position, the nozzles 104, 204 position, the selected aerodynamic profile characteristics and its configuration, the geometrical wing characteristics such as taper, twist, pitch setting and dihedral angle.

Thanks to the listed parameters, the electronic control unit 101, 201 is able to calculate the incidence angle value for all the points of the wing span of the aircraft 1 ; moreover the electronic control unit 101, 201 is able to calculate the sideslip angle, the rolling one and the first and second derivatives of all three the above-described parameters. Therefore, by adding the flight velocity, the parameters of interest useful to discern the system activation are ten. The electronic control unit 101, 201, by having data related to the inertial and geometrical quantities of the aircraft 1 in a memory, is able to estimate the values of the roll, pitch and jaw correction momenta. Those values are needed so that the system is able to modulate the thrust being provided by the auxiliary propulsion means. In other words, it is found the proper thrust value that the system must provide in the specific time where the described analyses are performed and the system is activated. To the following analyses of the parameters will potentially correspond following activations of the system. Since the flight parameters will be probably changed with respect the previous situation, the following activations of the system will produce different thrust values.

The block B shows the comparison of the acquired values with the minimum activation conditions of the system, and thus the parameters showing which the critical flight conditions for the given aircraft 1

The block C represents the analysis result of block B and decides if the acquired flight condition is a critical flight condition.

If the answer is negative then the control unit 101 , 201 restarts the cycle from block A. If the answer is positive then the control unit 101, 201 continues to the block D.

The block D consists of activating the system and thus generating the correction thrusts. That action is exerted until the desired flight conditions are reached or the propellant is finished.

The block E is the same as block A. The block F is the same as block B.

The block G is the same as block C.

If the acquired flight conditions do not satisfy the desired flight conditions then the control unit 101, 201 goes back to block D, thus to command the system activation. If, instead, the reached flight conditions satisfy the desired ones, the electronic control unit 101, 201 ends the system activation (block H) and goes back to block A.

The described elements, in both the operation modes, correspond only to functional descriptions, thus they can be present in a variable number (e.g.: one or more tanks, one or more nozzles, etc.) and/or can be present in different types (e.g.: different types of valves: control valves, safety valves, etc.).

Moreover, different activation modes can be considered, for example, bipropellant, solid and hybrid rocket engines.

The system can be placed on board of any type of aircraft. Moreover, it should be underlined how the placement of the system on board of any aircraft is variable as a function of the action efficiency needs of the system itself (for example advantageous placement in terms of correction forces "distances"). The positioning also follows system considerations (for example, relationship with other apparatuses existing on the aircraft, space availability to house the system itself) which depend on the particular aircraft on which the proposed system must be installed. In particular, the position of the propulsion means can be any position where those propulsion means are able to generate continuous forces having such a distance with respect to the center of gravity of the aircraft to directly create one or more correction momenta bringing the aircraft back in the proper flight attitude.

Therefore, the attached images are suggesting housing solutions (in the tail or near the wing) which are only solution examples for the system housing problem, and therefore they are not definitive and/or binding solutions for each aircraft on which the system can be installed.

During its activation, the system presents characteristics differentiating it from the common control systems of the aircraft such as the flight commands (ailerons, stabilizer and rudder): actually, those devices depend on the flight conditions and in particular on the aerodynamic conditions existing on the mobile surfaces.

In normal flight conditions, the flow is "not separated" from the mobile surfaces.

In a stall condition ("symmetrical stall" or "spin") the flow is "separated" and thus the mobile surfaces efficiency (e.g. ailerons and stabilizer) is strongly reduced or even zero. In those conditions, the aircraft control is strongly compromised, with effects being often catastrophic.

The proposed system efficiency instead essentially depends on the ratio between the pressure existing in the tank (in the combustion chamber, if the activation by means of monopropellant rocket engine is considered) and the external one at the nozzle exit. The external pressure at the nozzle exit is decided by the flight altitude and by the local conditions of the flow skimming the nozzle exit itself. However, it is not possible to assert that the system, during its operation, depends on the flight conditions. Actually, for any altitude and condition of the flow skimming the nozzle exit, the tank pressure is much higher than the external one. This implies that the system operating characteristics are depending on the tank pressure (or combustion chamber pressure in case of rocket engine activation) and on the characteristics of the nozzle where the gas expansion and thus the gas acceleration generating the correction force to bring the aircraft back in the proper flight attitude occurs.

Therefore, the propulsion means action is independent from the airflow hitting the aircraft surfaces during flight. In particular, the gas flows exiting the nozzles are independent from the airflow hitting the airfoils of the aircraft during flight. The system is automatic and independent, during its operation, from the pilot control, the correction maneuver (short, developing in few seconds) allows recovering the proper flight attitude without the delays due to the pilot decisional inertia.

It is possible to provide a manual intervention system by the pilot.

Claims

1. Auxiliary safety system for controlling the attitude of an aircraft (1), including sensor means located on the aircraft that control the general attitude of the aircraft, in particular the aircraft orientation with respect to the horizon and to the movement direction, and identify stalling and/or other critical flying conditions that are imminent and/or have occurred, characterized in that it comprises an electronic control unit (101, 201) connected to said sensor means to detect said stalling and/or other critical flying conditions that are imminent and/or have occurred, auxiliary propulsion means located on the aircraft being able to generate one or more continuous forces acting directly on the aircraft in order to create one or more correction momenta with respect to the aircraft center of gravity in order to bring the aircraft back in the proper flight attitude, said control unit (101, 201) activating said propulsion means in the event of said stalling and/or other critical flying conditions that are imminent and/or have occurred in order to generate a continuous thrust for the entire duration of the correction of the attitude and recover the proper flight attitude of the aircraft.

2. Auxiliary safety system for controlling the attitude of an aircraft (1), including sensor means located on the aircraft that control the general attitude of the aircraft, in particular the aircraft orientation with respect to the horizon and to the movement direction, and identify stalling and/or other critical flying conditions that are imminent and/or have occurred, characterized in that it comprises an electronic control unit (101, 201) connected to said sensor means to detect said stalling and/or other critical flying conditions that are imminent and/or have occurred, auxiliary propulsion means being able to generate one or more continuous forces acting directly on the aircraft in order to create one or more correction momenta with respect to the aircraft center of gravity in order to bring the aircraft back in the proper flight attitude, said control unit (101, 201) activating said propulsion means in the event of said stalling and/or other critical flying conditions that are imminent and/or have occurred in order to generate a continuous thrust for the entire duration of the correction of the attitude and recover the proper flight attitude of the aircraft, wherein the action of said propulsion means is independent of the airflow that hits the aircraft surfaces during flight.

3. Auxiliary safety system for controlling the attitude of an aircraft according to claim 1 or 2 wherein said propulsion means generate a continuous thrust by expansion of gas in one or more nozzles (104, 204).

4. Auxiliary safety system for controlling the attitude of an aircraft according to any of claims 1, 2 and 3, wherein said propulsion means comprise one or more compressed gas rocket engines.

5. Auxiliary safety system for controlling the attitude of an aircraft according to any of claims 1, 2 and 3 wherein said propulsion means comprise one or more monopropellant-gas rocket engines.

6. Auxiliary safety system for controlling the attitude of an aircraft according to claim 4, wherein said compressed gas propulsion means comprise one or more control valves (103), managed by said electronic control unit (101), to convey the gases from one or more tanks (102) to one or more nozzles (104) for the expansion of said gases, and further comprise sensor means (106, 107, 108) and on-board systems (H I) for detecting different physical and flight parameters and supplying the values thereof to said electronic control unit (101).

7. Auxiliary safety system for controlling the attitude of an aircraft according to claim 6, further including one or more compressors (112) for maintaining said one or more tanks (102) pressurized, said one or more control valves (103) closing the connection between compressors (112) and tanks (102) when the compressors (112) are not operating.

8. Auxiliary safety system for controlling the attitude of an aircraft according to claim 5, wherein said monopropellant-gas rocket engine propulsion means comprise one or more tanks (202) and one or more control valves (203), managed by said electronic control unit (201), to convey the gases to one or more combustion chambers (212) and then to one or more nozzles (204) for the expansion of said gases, and further comprise sensor means (206, 207, 208) and on-board systems (211) for detecting different physical and flight parameters and supplying the values thereof to said electronic control unit (201).

9. Auxiliary safety system for controlling the attitude of an aircraft according to claim 3 wherein the flows of gas exiting the nozzles (104, 204) are independent of the airflow that hits the aircraft airfoils during flight.

10. Aircraft (1) including an auxiliary safety system according to any preceding claim.