EP2289060B1 - Adaptation of selective terrain alerts, as a function of the instantaneous manoeuverability of an aircraft - Google Patents

Description

The present invention relates to the general technical field of aids to piloting rotary-wing aircraft, and in particular to automatic alerts for terrain avoidance.

To clarify the description, the existing technologies as well as the technical problems encountered with them will first be discussed in general. It is only then that the citation of various documents illustrating these technologies is gathered.

In the technical field mentioned, the invention relates to the so-called “on-board” piloting aid, that is to say at least partly on board manned aircraft, such as helicopters or convertible rotary-wing aircraft.

Also, the invention relates to so-called “remote” help. In this case, it is aimed at rotary wing drones, that is, unmanned rotorcraft. Then, this piloting aid according to the invention cannot be addressed to a pilot, since there is no one on board the aircraft. It is then addressed to a human operator who operates the remote control of said drone.

More specifically, the invention relates to piloting aids by warning for terrain avoidance, known by the acronym in English “TAWS” (Terrain Avoidance Warning System).

These TAWS systems must make it possible to indicate, as they approach, the dangerous obstacles located in front of the trajectory of the aircraft foreseen at a given moment, in a danger zone.

In other words, such a system makes it possible to automatically produce alerts, based on a mapping, if in a danger zone in front of the aircraft, an obstacle interferes with the planned trajectory of an aircraft at a given moment. given.

In fact, starting from known coordinates of the instantaneous position of the aircraft, as well as its flight plan and the cartography of the terrain overflown, an alert is issued in the event that an obstacle interferes with the planned avoidance trajectory while at the risk of making avoidance impossible.

The triggering of an alert is conventionally determined as a function of an avoidance trajectory considered as possible for the aircraft, its trajectory initially planned and its instantaneous speed.

In practice, it appeared that the warning systems for terrain avoidance (whether it be TAWS or other Ground Proximity Warning Systems called "GPWS") provided for airplanes, are not satisfactory for aircraft. rotary wing aircraft.

For example, the document EP0750238 (equivalent to document US5892462 ) which was deprived for lack of novelty, describes this type of collision avoidance system on the ground. This system is said to be adaptive.

Although this system seems generally dedicated to aircraft of whatever type, it is only suitable for airplanes. In particular, this system is not intended for a rotary wing aircraft or a helicopter. In addition, this document does not describe a conical curve, or a curve of its own type such as a parabola, ellipse or hyperbola. This document evokes a data update logic integrating parameters specific to the device and a notion of "possible maneuver". This document refers to an avoidance maneuver with a maximum value of the acceleration “G”, which does not explicitly require nor the taking into account of a power margin or outputs of an IPL. In fact, the avoidance maneuver at the maximum of G is mainly done by a resource using the pitch command so that the use of collective is not compulsory.

On the other hand, the teaching of this document EP0750238 does not take into account the instantaneous maneuverability of a rotary wing aircraft (cf. Column 13, lines 39-54). Such a calculation, as a function of up-to-date data (eg possible vertical acceleration and / or instantaneous mass) produced by flight equipment, is not described by this document. In a separate approach, this document provides that the terrain and input altitudes are taken from active terrain sensors, an inertial navigation system and a radar altimeter.

Due to the specific features of the structure and operation of rotary wing aircraft, the influence of certain parameters is more decisive on the effective possibilities of obstacle avoidance, for rotorcraft than for airplanes.

Indeed, a rotorcraft is capable of performing many different types of flight, in comparison with a fixed-wing aircraft. Apart from take-offs and touchdowns, for rotorcraft, only point-to-point transport flights are comparable with aircraft flights, in particular civil ones.

Thus, the same helicopter can notably perform close observation flights, tactical missions, rescues, disaster response, etc.

During such flights, the parameters taken into consideration and the alerts provided by a terrain avoidance system provided for an airplane are inappropriate, even undesirable or even dangerous.

The same applies to the take-off and touchdown phases, during which the piloting aids provided for airplanes prove to be inappropriate.

Faced with this observation, recommendations specific to helicopters were recently prepared by a major advisory authority in the field of aviation, namely the "RTCA (Radio Technical Commission Aeronautics)", relating to warning systems for terrain avoidance. . These recommendations dedicated to helicopters recommend systems called are called HTAWS.

Indeed, with conventional terrain warning technologies for airplanes, the anticipation distance of an obstacle involving a modification of the trajectory is almost exclusively calculated as a function of the absolute value of the forward speed of the airplane. .

Schematically, the higher this speed value, the longer the anticipation distance. In other words, the faster the flight, the farther ahead the terrain warning system monitors.

Thus, said anticipation distance is a value expressed in units of length (meters or kilometers for example). Since it is within this distance that the alert system checks whether or not there is an obstacle on the ground, this distance in front of the device is also called the danger zone.

Conventionally, the anticipation distance is most often evaluated by the multiplicative product between the instantaneous speed of the plane and a constant time which corresponds to a whole family of planes.

This anticipation distance includes a transfer time, that is to say the estimated duration of the pilot's reaction, which elapses between the alert and the initiation by the pilot of an avoidance path.

However, no other flight parameter (e.g. tactics, transfer, rescue, etc.) being taken into account, it then happens too often that in practice, alerts are triggered unexpectedly or too numerous. This hinders the pilot instead of assisting him. In fact, to overcome this inconvenience, the pilot sometimes comes to completely interrupt the operation of this piloting aid.

This is particularly common when an aircraft terrain warning system is adapted to a rotorcraft.

With such systems, the calculated avoidance trajectory also takes the form of a succession between a rectilinear section corresponding to the transfer time, to which is attached an arc of a circle oriented in the direction of a distance from the obstacle. This is then referred to as a “ski tip” trajectory.

In other words, most current systems provide in practice, a rectilinear transfer time based on the speed and an avoidance curve in an arc of a circle whose radius corresponds to a maximum safety margin, without really taking into account the intrinsic capacities. actual or instantaneous situation of the device.

Of course, the “ski tip” avoidance trajectory is calculated so that the pilot can act on the airplane so as to avoid the obstacle in a danger zone.

But as we have seen, due to its method of calculation, it is frequent, for example in tactical flight, for alerts to be triggered in the absence of real danger or to be erroneous or even to be almost permanent.

From the foregoing, it is understood that it would be appropriate to propose a terrain warning system for rotary wing aircraft, which only generates alerts that are truly useful to its pilot, at the most opportune time possible, that is to say - say neither too early nor too late. This selective exclusion of superfluous alerts is called “reliability”.

In addition, a terrain alert system offering increased safety for rotary wing aircraft would be desirable, so that an alert does not require recourse to the optimal or even maximum and instantaneous avoidance possibilities (ie maneuverability) of the aircraft. affected device, either inhibited or postponed to a later time.

This would make it possible to preserve the flight path as close as possible to the terrain, without increasing the risks specific to the obstacles in this terrain. Such increased security would be highly desirable, for example in the case of tactical military flight.

It is nevertheless understood that the security requirements on the one hand, and the flight constraints on the other hand, are contradictory when it comes to developing, in practice, a terrain warning system for winged aircraft. rotating that would generate alerts both at the right time, while remaining reliable and safe in terms of avoidance possibilities.

In order to alleviate these drawbacks in particular, it was chosen with the invention not to base the origin of the avoidance almost exclusively on the measured absolute value of the instantaneous speed.

However, three additional technical problems condition this approach in practice.

First, the logical integration of maneuverability parameters is complex, in particular compared to terrain warning systems for airplanes which in practice only use the integration of a single and absolute value (without physical unit) of speed.

Secondly, it is undesirable that obtaining significant maneuverability parameters requires additional dedicated equipment, such as sensors, cabling and on-board PLCs. This would make the apparatus unacceptably heavy.

Thirdly, since the piloting assistance methods are executed by computers programmed using computer codes, it is not possible to design and write a complete and specific code or algorithm, for each model, each type and each configuration of rotary wing aircraft.

More specifically, concerning the logical integration of maneuverability parameters, it is understood that the instantaneous maneuverability of an aircraft is correlated with a large number of parameters, which should therefore be sorted and qualified. and make them compatible with each other as well as with their integration into the field warning system.

In particular, these parameters include the aircraft model in question, in the sense that a light, powerful and modern aircraft model has greater maneuverability than another aircraft model, less light, powerful and modern.

However, for a given model of rotary wing aircraft, the maneuverability varies significantly due to various situations.

• son environnement de vol (température et pression atmosphérique ambiantes, altitude, hygrométrie, poussières, etc.) ;
• sa phase de vol (décollage, croisière, approche, posé, etc.) ;
• son état fonctionnel initial lors du vol donné (i.e. les états de maintenance, de vieillissement, de remplissage réservoirs, de chargement à bord, les équipements embarqués, etc.) ;
• son état instantané (i.e. les paramètres de fonctionnement à un instant donné, tels que température / pression des fluides et flux, charge électrique restante, masse totale de l'appareil, puissance motrice disponible, mode de pilotage à vue ou aux instruments, etc.) ;
• sa feuille de route (mission civile ou militaire, tactique ou de simple transport, sauvetage, etc.).

In particular, the maneuverability of a device changes according to parameters such as:

• its flight environment (ambient temperature and atmospheric pressure, altitude, hygrometry, dust, etc.);
• its phase of flight (take-off, cruise, approach, touchdown, etc.);
• its initial functional state during the given flight (ie the states of maintenance, aging, filling tanks, loading on board, on-board equipment, etc.);
• its instantaneous state (ie the operating parameters at a given instant, such as temperature / pressure of fluids and flows, remaining electric charge, total mass of the device, available motive power, visual or instrument piloting mode, etc. );
• its roadmap (civil or military mission, tactical or simple transport, rescue, etc.).

It would therefore be advantageous to be able to efficiently integrate such parameters into the method for determining the avoidance trajectory, without increasing and slowing down the calculations for triggering an alert.

It will be understood that this would amount to adapting in real time the response of a terrain warning system for rotary wing aircraft, as a function of the actual performance of this aircraft at a given instant, and in particular of its maneuverability.

In addition, it would be advantageous for the avoidance trajectory to be more in line with the terrain, compared to a “ski tip”.

For this purpose, the invention proposes an avoidance trajectory with a substantially rectilinear section, and proximal to the device. This proximal section would reflect the transfer time, without major recourse to the speed of the aircraft.

The avoidance trajectory proposed by the invention comprises a section contiguous to the preceding one, of conical curvilinear shape.

The frame of reference in which such an avoidance trajectory would fit would include an axis comparable to an abscissa, linked to the speed of the rotary wing aircraft at a given instant, which it is desired to slow down. The other axis of this reference, comparable to an ordinate, would correspond to the vertical acceleration capacity.

We understand that in this way false alarms could be limited, and often avoided.

But failing to be able to obtain these parameters without complicating or weighing down the rotary wing aircraft, the advantages obtained by taking them into account (that of these parameters) would be greatly reduced, or even annihilated.

The question therefore arises of obtaining significant, consistent parameters that qualify as maneuverability, without requiring significant additional dedicated equipment.

This dilemma is unexpectedly resolved. In summary, the useful parameters are obtained by judicious approximations, from data produced by standard flight equipment in modern rotary wing aircraft.

In particular, these approximations are made possible by the logical coupling, that is to say the puncture for piloting assistance, of data already available on board. This goes against the usual existing prejudices.

On the other hand, the invention provides for choices contrary to the obvious as regards the data resulting from this puncture, allowing them to be both significant and compatible with the approximations to be made, which is advantageous.

• une distance de transfert sous forme de temps minimisé en fonction d'au moins un paramètre disponible à bord, tel que la feuille de route (e.g. vol tactique ou de croisière) ; et
• une trajectoire proposée d'évitement qui est optimisée, comportant au moins un tronçon conique (non circulaire), calculé notamment en temps réel en fonction de données à jour produites par des équipements de vol, telles que l'accélération verticale possible à partir du pas collectif du ou des rotors de propulsion et sustentation et / ou de la masse instantanée de l'aéronef à voilure tournante.

Thus, one embodiment of the invention provides in particular:

• a transfer distance in the form of minimized time as a function of at least one parameter available on board, such as the route map (eg tactical or cruising flight); and
• a proposed avoidance trajectory which is optimized, comprising at least one conical section (non-circular), calculated in particular in real time as a function of up-to-date data produced by flight equipment, such as the vertical acceleration possible from the pitch collective of the propulsion and lift rotor (s) and / or of the instantaneous mass of the rotary wing aircraft.

Said data produced by flight equipment has been produced by one or more existing or conventional flight equipment, for example by a First Limitation Indicator or “IPL”, as explained above.

Indeed, many rotary wing aircraft already have flight equipment such as an IPL, which continuously calculates an available power margin, here in the form of the collective pitch value of its or its so-called rotors. "Main".

This collective pitch value is then available on board without requiring additional equipment. This collective pitch value corresponds to the product of the vertical acceleration possible at a given instant, multiplied by a coefficient proportional to the mass (either at take-off or estimated at the chosen instant) of the aircraft.

The integration of this significant collective step value of the power margin makes it possible to obtain, in a simple way, a terrain warning system which would then be said to be "adaptive", for the calculation of the danger zone, of the distance. transfer and the avoidance curve, without influencing the specific algorithm of this piloting aid.

Thus, depending on the situations with which a rotary wing aircraft is confronted, the avoidance trajectory is optimized.

This avoidance trajectory approaches a tangent to the horizontal, in the event that little or no power margin is available.

In contrast, the avoidance trajectory approaches a tangent to the vertical, the lower the longitudinal speed of the device and / or the greater the available power. This The latter situation is particularly useful in the case of tactical flight, since it reduces the danger zone to a minimum, while allowing safe flight at low altitude in relation to the terrain.

This integration of values obtained by existing flight equipment also makes it possible to overcome the third additional technical problem mentioned, namely the writing of a unique algorithm, complete and compatible with many models, types and configurations of winged aircraft. rotating.

Indeed, the parameters obtained can be logically considered as simple variables which only have to be injected as data, in a single algorithm, that is to say compatible with a wide range of winged aircraft. rotating.

Let us now cite various documents relating to piloting aids, in addition to the document EP0750238 equivalent to document US5892462 .

In general, we find in particular in Circular N ° 0236-2005.07.29, of Transport Aviation Civile du Canada, definitions and some brief explanations of the various on-board impact warning and alarm systems (TAWS). , other collision avoidance systems and forward looking terrain avoidance systems (FLTA).

In the same vein, Recommendation RTCA-309 relating to future HTAWS systems, proposes functionalities to be provided for these systems dedicated to helicopters.

The document XP010668914 (BARNHART et. Al.) IEEE, US, Vol. 2 of 12 October 2003 pages 9A109-9A113, describes a TAWS for helicopter . This document indicates that it is classic either maneuver the roll control and then command a maximum G to right the aircraft, or a second trajectory aims to eliminate these wing maneuvers and maintain an angle in memory while commanding a maximum G, these two trajectories being identical (p9A106, col 2). On his figure 4 showing avoidance trajectories, we see a straight line before obstacle separated from an almost straight ascending avoidance trajectory. According to page 9A107 col. 2, in a navigation improvement by terrain referencing (TRNE), the data existing in the aircraft (INS, RADALT and GPS) are merged using a multi-level Kalman filter, in an attempt to position the aircraft horizontally with respect to the terrain database.

The document US2002036574 describes a computer program for a helicopter improved terrain proximity alarm system. in FIG. 6, an observed reaction time of 5 seconds and a stopping distance required by employing 10 degrees of angle in altitude gain or by maintaining the initial altitude, and juxtaposed to a straight trajectory.

The document XP002526709 of September 15, 2006 (Section 8 - Appendix / Index; page 8-23) "Emergency / anormal and normal procedures from the Airplane Flight Manual Supplement ", is a manual for piloting assistance by warning for terrain avoidance, known by the acronym in English" TAWS "(Terrain Avoidance Warning System). This aid should make it possible to indicate, as and when measurement of their approach, the dangerous obstacles located in front of the trajectory of the aircraft foreseen at a given moment, in a danger zone This indication is provided by the display of bubbles (pop-up) on the screen.

The document XP002526710, FAA “Advisory Circular AC-27-1B-MG-18” of June 24, 2005: “CHAPTER 3 AIRWORTHINESS STANDARDS TRANSPORT CATEGORY ROTORCRAFT (page 1 ) ”, Is an extract from a regulation relating to warning systems for terrain avoidance for helicopters (HTAWS) with a view to improving the so-called“ Ground Proximity Warning Systems ”or“ GPWS ”.

The document FR2756256 describes an IPL power margin indicator for a rotary wing aircraft. This indicator is intended to provide information on the power margin available on at least one engine and one gearbox of the aircraft as a function of the flight conditions. It is typically from the units also serving as a source of real-time data for this type of indicator that the invention proceeds to generate an alert. But in a manner distinct from what this document provides, the invention defines an avoidance trajectory with a proximal section significant of a transfer time and an avoidance curve with a distal section in the form of a conical curve.

The document EP1726973 describes a method and a device for assisting the piloting of an aircraft at low altitude. According to one embodiment, a guard height is predetermined by the pilot or may be a function of parameters of the aircraft, which in the latter case gives it a dynamic aspect.

The document FR1374954 offers an automatic pilot for very low altitude aircraft flights, in which maneuvers are limited in their effects to a determined minimum.

The document FR2813963 describes a visual display of ground collision avoidance information in a aircraft and especially an airplane. A control factor comprises the distance to the obstacle, but also the variation of said distance, and the direction of the speed vector, depending on whether it is ascending, horizontal or descending.

To avoid the overload of information and alarms during the take-off and landing phases, certain information is inhibited insofar as the lowest point is below a chosen altitude and the proximity of the aircraft with its landing zone corresponds to a validated criterion.

To this end, static and dynamic parameters are taken into account, including components of the speed vector, and if applicable of the acceleration vector. According to this document, in the approach phase, the predicted axis can be curvilinear and the vertical plane is not necessarily flat.

The document FR2749545 describes the foundations of an Indication of First Limitation (IPL) system. This system determines the power margin available on one or more engines of an aircraft, depending on its flight conditions. The goal is to allow a pilot to “remove” the relevant information for piloting.

Furthermore, this document indicates that the information provided by the indicator (IPL), in addition to the display display, can serve as basic information for the development of a force law, making it possible to warn the pilot when it approaches a limitation by physical means: hardening of a spring or jack, vibrations for example.

Besides the document FR2749545 , the documents FR2749546 , FR2755945 , FR2756256 , FR2772718 , FR2809082 , FR2902407 and FR2902408 describe characteristics specific to "IPL" First Limitation Indicators, and have all been filed by the Applicant. Their teaching is integrated into the present application, to avoid unnecessary repetitions.

The document FR2712251 describes a low-altitude piloting aid. To determine the dangerous obstacles and help to avoid them, the position of an optimal avoidance point is calculated in particular from the speed vector of the helicopter. A nose-up limit load factor depends on the mass of the helicopter in particular. An audible alarm is possible, in addition to the visual display. A search zone per angular sector is limited to a distance L from the helicopter.

The document FR2853978 describes a method and a device for securing the flight of an aircraft in instrument flight conditions outside of instrument flight infrastructures. A theoretical trajectory is corrected which takes into account the forecast possibilities of the aircraft so as to obtain a secure route which the aircraft must follow. This route takes into account the current possibilities of the aircraft, through current parameters (weight, performance, configuration) relating to the aircraft; the current flight point (altitude, position relative to the theoretical trajectory, speed) of the aircraft; and also current parameters (wind, temperature, pressure) relating to the external environment.

The document FR2886439 describes a low-altitude piloting aid, for performing contour or tactical flight. For this aid, an optimum curve is determined as a function of the speed of the aircraft.

The document US3245076 aims to make optimum use of the maneuvering capabilities of the aircraft in an automatic pilot.

The document US3396391 mentions the use of representations of acceleration as well as aircraft load factors to calculate a flight path. The speed of an aircraft is taken into account in determining a desired height to the ground.

The document US6347263 describes an aircraft terrain alarm generator which has an alert envelope with a lower limit formed from the lesser value between a flight direction angle and a possible ascent gradient.

The alert envelope has a first segment between two points, and the aircraft's climb prediction is calculated based on various parameters, such as the predictable pitch-up, lift, drag, estimated weight of the aircraft. aircraft.

The document US6380870 describes a determination of the forward distance value, for high speed flight, typically for an airplane. The objective is to make the flight as constant as possible, by switching between a value with variable reaction to a value with constant reaction at high speed. Also, this limits spurious alerts at low speed.

The document US6583733 describes a ground proximity warning system for a helicopter, with a first mode and a second mode of operation. These modes are selected by the pilot. A display for the pilot is shown. This system is described as a TAWS or GPWS, which integrates the flight specifics of a rotorcraft compared to a fixed-wing aircraft. In addition, the objective is to adapt the system to the type of flight in progress, while taking into account the instantaneous possibilities of the aircraft and limiting parasitic alerts. For this purpose, information is collected from a GPS.

The document US7064680 describes a frontal sweeping obstacle avoidance (FLTA) for airliner, which conventionally provides audible alerts in the form of a warning (eg “terrain”) and a warning (eg “Cabrer”). In addition, once the avoidance maneuver is completed and depending on a horizontal projection of the aircraft previously said end, an audible alert is emitted with both a warning (eg "terrain") and a end of danger notice (eg "clear", ie "clear").

As for the present invention, it aims to provide adaptive, safe and reliable piloting assistance, by integrating data compatible with useful approximations, as well as significant for the instantaneous maneuverability of a rotary-wing aircraft, such as a helicopter. , a convertible aircraft or a drone.

For example, such an aid proposes an HTAWS logically coupled to an IPL, which executes algorithms ensuring the integration of instantaneous maneuverability data, to issue selective alerts that are sufficiently safe and reliable, in particular which are not overabundant.

The alerts thus made selective can include a dedicated audible alarm, while remaining effective, comfortable, that is to say not intrusive.

For example, such an audible alarm takes the form of an explicit and contextual voice message, audible by the person at the controls and relieving his attention to allow him to concentrate on the piloting instruments to be operated.

To this end, the invention is defined by the claims.

Various embodiments of a method according to claim 1, of terrain warning device according to claim 7 and of a rotary wing aircraft according to claim 9 and in accordance with the invention are described.

An example of the invention relates to a method for generating an alert for the avoidance of terrain by an aircraft with a rotating car.

This method provides for the development of an avoidance trajectory which comprises a proximal section significant of a transfer time and an avoidance curve.

Said proximal section is extended in the continuation of a planned trajectory, over a distance which reflects an applicable duration minimized as a function of a route map of the aircraft.

Said avoidance curve comprises at least one distal section of conical profile, attached to the proximal section and calculated as a function of the instantaneous maneuverability of the aircraft.

In one implementation of the method, the proximal section is substantially rectilinear.

In one implementation of the method, the minimized applicable duration is a function of a route map and a parameter reflecting the model of the aircraft.

According to one implementation of the method, the applicable duration is minimized as a function of a route map, then by division by at least one limiting ratio reflecting a flight parameter of the aircraft.

In one embodiment, the conical curve is of its own type, such as a parabola, ellipse or hyperbola.

According to one implementation of the method, the conical curve is calculated in real time, as a function of up-to-date data produced by flight equipment, including a possible vertical acceleration value and / or a value of the instantaneous mass of the aircraft. rotary wing aircraft.

Another example of the invention relates to a terrain warning device.

This device is logically coupled to a maneuverability indicator system, for example an IPL.

According to one embodiment, this device is at least partially on board, and comprises flight equipment with a flight computer capable of executing a code which allows the implementation of the above method.

Yet another example of the invention relates to a rotary wing aircraft, which is either a helicopter, or a convertible aircraft, or a rotary wing drone.

According to one embodiment, this aircraft is able to implement the method mentioned above and / or comprises a terrain warning device as mentioned.

This aircraft has, in one embodiment, an audible alarm intended to be triggered selectively by the terrain alert.

• la figure 1 est une vue partielle schématique en perspective d'élévation longitudinale, qui illustre une réalisation d'aéronef à voilure tournante, ici un hélicoptère, équipé et apte à mettre en œuvre l'adaptation d'alertes de terrain conforme à l'invention, notamment en fonction de la manœuvrabilité de cet appareil ; sur cette figure sont représentés à des fins de comparaisons, les temps de transfert (TT) et les courbes d'évitement (CE) selon les techniques connues en partie supérieure (traits en tirets), ainsi que selon l'invention en partie inférieure (traits alternés discontinus) ;
• la figure 2 est une vue partielle schématique en perspective d'élévation longitudinale, qui illustre une réalisation d'aéronef à voilure tournante, ici un aéronef convertible, équipé et apte à mettre en œuvre l'adaptation d'alertes de terrain conforme à l'invention ;
• la figure 3 est une vue partielle schématique en perspective d'élévation longitudinale, qui illustre une réalisation d'aéronef à voilure tournante, ici un drône et son poste de radiocommande à distance, équipé et apte à mettre en œuvre l'adaptation d'alertes de terrain conforme à l'invention ;
• la figure 4 est une vue partielle schématique qui illustre une réalisation de dispositif d'alerte de terrain pour d'aéronef à voilure tournante conforme à l'invention ; et
• la figure 5 est un graphe logique qui illustre les principales étapes et phases conformes à l'invention, d'une réalisation de procédé d'alerte de terrain pour d'aéronef à voilure tournante, notamment en fonction de la manœuvrabilité de cet appareil.

The invention is now described with reference to the exemplary embodiments, given without limitation and illustrated by the appended drawings, in which:

• the figure 1 is a schematic partial perspective view of longitudinal elevation, which illustrates an embodiment of a rotary wing aircraft, here a helicopter, equipped and able to implement implement the adaptation of field alerts in accordance with the invention, in particular as a function of the maneuverability of this device; in this figure are shown for comparison purposes, the transfer times (TT) and the avoidance curves (CE) according to known techniques in the upper part (dashed lines), as well as according to the invention in the lower part ( alternating broken lines);
• the figure 2 is a schematic partial perspective view of longitudinal elevation, which illustrates an embodiment of a rotary wing aircraft, here a convertible aircraft, equipped and able to implement the adaptation of terrain alerts according to the invention;
• the figure 3 is a schematic partial perspective view of longitudinal elevation, which illustrates an embodiment of a rotary wing aircraft, here a drone and its remote radio control station, equipped and able to implement the adaptation of terrain alerts in accordance with to invention;
• the figure 4 is a schematic partial view which illustrates an embodiment of a terrain warning device for a rotary wing aircraft according to the invention; and
• the figure 5 is a logic graph which illustrates the main steps and phases in accordance with the invention of an embodiment of a terrain warning method for a rotary-wing aircraft, in particular as a function of the maneuverability of this device.

Overall figures 1 to 5 , like elements are designated by the same reference numbers.

In the figures are shown three directions X, Y and Z orthogonal to each other, which form a reference orthogonal X, Y, Z. When this is necessary, this reference X, Y, Z is orthonormal, for example to lighten the calculations

A so-called longitudinal direction X corresponds to the main lengths or dimensions of the structures described. Thus, the longitudinal direction X defines the main axis of advance of the aircraft described, and the tangent to their instantaneous trajectory at their center of gravity.

Another so-called transverse direction Y corresponds to the trajectories or lateral coordinates of the structures described; these longitudinal X and transverse Y directions are sometimes called horizontal, for simplicity.

A third direction Z is said to be elevation and corresponds to the height and altitude dimensions of the structures described: the terms up / down or nose up / nose up refer to them; for simplicity, this direction Z is sometimes called vertical.

For example, the term "nose-up" designates an action on the trajectory, which brings it closer upwardly to the tangent of said direction of elevation, while the term "to pitch" indicates a downward approach of the trajectory towards the longitudinal direction. .

The X and Y directions jointly define a so-called main X, Y plane within which the support polygon of an aircraft described is inscribed.

In the figures, the reference 1 generally designates a rotary wing aircraft or rotorcraft, in the sense that it has at least one rotor 2 for propulsion and lift.

In other words, the aircraft 1 according to the invention are capable of vertical takeoffs and lift flights. Some aircraft 1 according to the invention have several propulsion and lift rotors 2, for example two rotors 2 in tandem or on top of each other. An engine 44 is of course provided on each aircraft 1.

On the figure 1 , the aircraft 1 is a rotorcraft, and more particularly a helicopter 3 according to the invention, with a single propulsion and lift rotor 2, as well as an anti-torque rotor 4 on its tail.

On the figure 2 , the aircraft 1 is a convertible apparatus 5 according to the invention, provided with two propulsion and lift rotors 2, which are orientable.

The figure 3 shows an unmanned rotary-wing aircraft 1, here a drone 6 and its remote radio control station 7, in accordance with the invention. This drone 6 has a single rotor 2 for propulsion and lift. Certain drones 6 according to the invention have at least two rotors 2, sometimes superimposed and integrated into a fuselage 8, for example in the shape of a saucer.

All the aircraft 1, 3, 5 and 6 in accordance with the invention have at least one electronic flight device 9, such as that which is shown diagrammatically in dotted lines on the diagram. figure 4 .

Also, each flight equipment 9 has at least one piloting aid such as the terrain warning devices 10 shown in the figures. figures 1 to 5 .

These devices 10 are impact warning and alarm systems, and typically but not exclusively TAWS.

Each impact warning and alarm device 10 makes it possible to produce and supply the person (pilot or remote operator) who drives the aircraft 1, an avoidance trajectory designated in TA on the figures 1 to 3 .

In the examples, each aircraft 1 has an alarm 45, able to be triggered by the device 10. The alarm 45 is audible and / or displayed.

The avoidance trajectory TA is here formed by two contiguous sections, one of which is proximal to the aircraft 1 which is substantially rectilinear and whose projection on a transverse and longitudinal plane (X, Y) reflects the transfer time (TT ).

The other section of the avoidance trajectory TA, draws at least in part and / or transiently a curve and is distant from the aircraft 1. We speak of a distal or curvilinear section.

On the figure 1 , this curvilinear section is extended continuously with respect to the proximal section and has a projection on said transverse and longitudinal plane (X, Y) which expresses the travel time by the aircraft 1 of its avoidance curve (CE).

According to the invention, the distal section comprises, or even fits entirely in, a conical curve, whereas in the ski-tip trajectories, the latter is in the arc of a circle.

At this stage, it therefore seems useful to recall the concept of a conical curve.

Conics form a family of curves resulting from the intersection of a plane with a cone of revolution. Conics are said to be clean, when the secant plane is not perpendicular to the axis of the cone, and does not pass through its vertex. We will see that the curvilinear sections of the avoidance trajectory TA according to the invention are frequently of their own conical type.

There are also three kinds of proper conics depending on the angle of inclination of the plane secant with the axis of the cone: ellipses, parabolas and hyperbolas. All these clean conical curves can give the trace of the avoidance curve CE of the trajectory TA, according to the invention.

If the two angles are equal, the conical curve is a parabola. A mono focal definition of conics implies a focus and a directrix.

More commonly, a conic curve is put into a second order algebraic equation, in affine analytical geometry, by considering these conics as plane curves, that is to say curves whose Cartesian coordinates x and y of the points along X and Y respectively, are the solutions of a quadratic polynomial equation, of the form: $$\mathrm{<figure-callout\; id="2"\; label="Ax"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">Ax</figure-callout>}2+\mathrm{Bxy}+\mathrm{<figure-callout\; id="2"\; label="Cy"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">Cy</figure-callout>}2+\mathrm{Dx}+\mathrm{Ey}+\mathrm{F}=0;$$

Where A, B, C, D, E and F are the coefficients of the conical curve.

The frame of reference used in the examples, is the frame of reference formed by the three orthogonal directions X, Y, Z, where x, y, z are the variables of the points of the curve respectively on one of said axes or directions X, Y, Z .

If E is not zero, a translation along the X axis of the variables y cancels F (F being the focus of the parabola). Then, by setting p = - A / E, it is possible to obtain the reduced Cartesian equation of a written parabola: y = px2. If D is not zero, the reduced equation of a parabola is written: x = qy2.

In the case of parabolas, these conical curves are obtained by the intersection of a cone of revolution with a plane, said parabola appearing when this plane is parallel to one of the generatrices of said cone.

We consider that the parabola is given by its focus F and its directrix D. Then, a projection O is obtained by orthogonal projection of the focus F on the directrix D. A parameter of the parabola is called "p", and corresponds to the distance OF, which forms a segment [FO]. This segment [FO] has a midpoint S. Then, in the reference X, Y, Z (which we consider orthonormal), where Z is in the same direction and sense as the vector OF , we write the equation of the parabola in the form: y = x2 / 2p.

These geometric details having been provided, let us return to the invention.

In general, the terrain warning device 10 is at least partly on board, in the sense that it is essentially located on board the aircraft 1. However, in embodiments, components of such a device 10 according to invention are on board while others are at a distance from the aircraft 1.

For example, in the particular case of a drone 6 such as on the figure 3 , this warning device 10 is physically partly on board the aircraft 1, and partly integrated into its radio control station 7, or even deported by a data transfer link 11 to a dedicated computing center 12. This link 11 is a telecommunication link on the figure 3 .

In addition to the warning device 10, the flight equipment 9 includes various other functionalities such as navigation aids, automatic piloting, a ground proximity warning system, front-sweeping obstacle avoidance, an algorithm. descent, an on-board collision avoidance system, a traffic warning and collision avoidance system, a global positioning system, etc.

It should already be noted that the flight equipment 9 as well as its sub-assemblies such as the device 10, operate iteratively and in real time.

According to the invention, the flight equipment 9 has what will be called here a maneuverability indicator system 13. It too operates iteratively and in real time.

Such a maneuverability indicator system 13 is able to produce and / or provide various significant and contextual data or parameters, from which it is possible, thanks to the characteristics of the invention, to have aircraft maneuverability indicators. 1.

Subsequently, and to clarify the description without limiting it, only one so-called main rotor 2 is considered, knowing that a person skilled in the art is able to implement the invention on the basis of this exposed, in the various cases where the aircraft 1 has several propulsion and lift rotors 2.

In fact, the system 13 therefore takes into consideration all the rotors 2 of the aircraft 1, and therefore supplies data reflecting the situation of the whole of this aircraft.

In the examples illustrated, the maneuverability indicator system 13 comprises a first Limitation indicator or IPL. Obviously, other systems 13 are compatible with the invention, as long as they provide the necessary data, the content of which will be seen below.

In one embodiment, the maneuverability indicator system 13 takes up the teaching of the document FR2756256 so as to provide information on the available power margin as a function of the flight conditions. From control parameters and values of engine use limitations, a power margin indicator is developed, expressed in particular as a collective pitch value.

As indicated previously, the IPL system 13 continuously calculates an available power margin, in the form of the collective pitch value of the so-called “main” rotor 2 of the aircraft 1, whether it is the helicopter 3. , convertible 5 or drone 6.

This collective pitch value is therefore available for the flight equipment 9 and in particular for the piloting aid device 10.

This value of the available collective pitch corresponds to the product of the vertical acceleration noted here "Gz", possible at a given instant, multiplied by a coefficient K proportional to the mass of the aircraft 1. The coefficient K is initialized on take-off, and estimated in real time at the chosen time.

Insofar as this value Gz is comparable to a constant during the calculation of the conical curve 24 of the avoidance CE, then this curve takes the form of a parabola.

Note here that the equipment 9, just like the device 10 and the system 13 comprise at least one computer 14, programmed according to computer codes 15 ( Figures 4 and 5 ).

In particular in the device 10 according to the invention, the programming code 15 or complete algorithm has been designed and written to be compatible without substantial modification, with the largest possible number of aircraft models 1. Only the models. data or parameters injected into this code 15 therefore allow adaptation of the invention to each type and / or each configuration of rotary wing aircraft 1.

An example of a terrain alert functionality which takes into account in particular the room for maneuver of the aircraft 1, is now exposed with reference to the figures 1 , 4 and 5 .

On the figure 1 , we see a terrain 16 above which the aircraft 1 performs a flight. In the aircraft 1, and more especially in its equipment 9, a map 17 is recorded which reflects this overflight terrain 16.

However, some of the recordings useful for the equipment 9 are sometimes deported outside the aircraft 1, in particular in the case of the drone 6.

On this terrain 16, there is an obstacle 18. For its part, the aircraft follows a flight plan 19 recorded in the equipment 9, according to which is defined a trajectory 20 provided for the flight, which is shown diagrammatically on the diagram. figure 1 by a straight line, for the sake of simplicity.

It can be seen that the obstacle 18 is placed on the planned trajectory 20, at a certain distance 21 in front of the aircraft 1, so that there is a risk of collision.

As we have seen, the invention proposes to emit at the most opportune moment an alert expressing this risk, while allowing the aircraft 1 to fly as close as possible to the terrain 16.

We also see on the figure 1 , an anticipation distance 22, that is to say the distance between the obstacle 18 and the position of the aircraft 1 at the moment T0 when the alert is issued. We call back that usually, this distance 22 is calculated as a function of the flight speed of the aircraft 1.

The alert only intervenes if the obstacle 18 is on the intended trajectory 20 of the aircraft 1, and within the distance 22 also called the danger zone.

On the figure 1 , the classic TA avoidance trajectory in dashes which corresponds to the transfer time TT abutting the circular arc CE, clearly shows the drawbacks of technologies based on flight speed (often flight speed multiplied by a given transfer time , most often constant).

In summary, the anticipation distance is excessive in relation to the effective resources of the aircraft 1, which furthermore avoids the obstacle 18 by flying over it at a height markedly greater than what it really requires to comply with the flight instructions, real context and security.

To improve the terrain warning systems, in particular by reducing the anticipation distance 22, also by eliminating unnecessary alerts and by making it possible to avoid obstacles as closely as possible, the maneuverability margin of the aircraft 1 is in particular taken into account according to the invention.

• un temps de transfert TT limité à l'instar du temps de réaction 23 sur la figure 1, par exemple de l'ordre de 0,5 à 2 secondes, ceci correspondant à un tronçon proximal 25, sensiblement rectiligne, derrière lequel on accole ;
• une courbe d'évitement CE en forme de tronçon distal 24 et comportant une courbe conique, par exemple parabolique, qui est fonction de la manœuvrabilité instantanée de l'aéronef 1.

We have seen that according to the invention, the value of the flight speed is not preponderant in the determination of the avoidance trajectory TA, because we take to form this trajectory TA:

• a limited TT transfer time like the reaction time 23 on the figure 1 , for example of the order of 0.5 to 2 seconds, this corresponding to a proximal section 25, substantially rectilinear, behind which one adjoins;
• an avoidance curve CE in the form of a distal section 24 and comprising a conical curve, for example parabolic, which is a function of the instantaneous maneuverability of the aircraft 1.

Therefore, for a highly maneuverable aircraft 1 and / or in a demanding flight context, the avoidance trajectory TA = section 25 + section 24, is short, that is to say included in an anticipation distance 22 minus extended in the longitudinal direction X, than the distance 22 calculated for an aircraft 1 less maneuverable and / or in a less demanding flight context.

Consequently, above, the terrain alert is emitted by the device 10 at a less distance 22 from the obstacle 18, in the first case than in the second.

As mentioned, the reaction time 23 to which a proximal section 25 corresponds is evaluated in particular as a function of the route map 41 of the flight performed by the aircraft 1.

Thus, if the flight is a tactical military mission operated by a modern, light and powerful aircraft 1, this reaction time 23 will for example be of the order of 0.5 to 1 second. If the flight is a simple transport operated by a basic aircraft 1 of high tonnage, this reaction time 23 will for example be of the order of 1 to 2 seconds.

Of course, the finally fixed value of this reaction time 23 can first be evaluated as a function in particular of the flight sheet 41, then adjusted as a function of contextual values, such as variable parameters significant of the state of the air. aircraft 1 on the verge of avoidance, for example obtained in real time.

In one example, the TT / 23 transfer time that defines the proximal section 25 is calculated by device 10, as follows.

First, an initially applicable time or duration value of between approximately 0.5 seconds and 2 seconds is determined based on the model of aircraft 1.

Then, a limitation weighting is carried out over this initially applicable period, as a function of the roadmap 41. This weighting is for example a first limitation, such as a division by a first limiting ratio.

In one implementation, this first ratio is of the order of 1 for a cruising flight, and of the order of 1.1 to 2 for a tactical military flight. This provides a TT / 23 transfer time which is taken into account for the terrain alert determination, by device 10.

In another implementation, further adjustment is made to result in a doubly reduced TT / 23 transfer time. Thus, the TT time provided by dividing the duration initially applicable by the first ratio is still limited by division, this time as a function of a parameter reflecting an acceptable increase in risk, here the experience of the person in charge. of the aircraft 1.

One implementation provides that the second limitation parameter reflects the degree of piloting expertise. If the pilot or the ground operator is experienced, this second limitation parameter is of the order of 1.1 to 1.3, for example 1.25. If the pilot or the ground operator is normally qualified, this second limitation parameter is of the order of 1.

Another implementation provides that the second limitation parameter reflects the priority factor of the mission. Yes the mission is a high priority, and includes an intrinsic risk as in wartime, the second limitation parameter is of the order of 1.1 to 1.2, for example 1.15. If the mission is of more common importance, this second limitation parameter is of the order of 1.

Note that according to the invention, the proximal section 25 is not systematically rectilinear. In fact, it is obtained in some embodiments by continuation of the trajectory 20 provided, whether straight or curvilinear, during the time TT obtained and noted at 23.

Thus, with the invention, it is possible to further improve the safety and reliability of the avoidance alert, by taking into consideration data which translates at a given moment, the true structural state of the aircraft 1 , that is to say to further shorten, if possible, the distances 21 and 22.

Indeed, for a given aircraft 1, in particular according to the instantaneous operating conditions (including the temperatures and pressures influencing the engine 44) and according to its instantaneous state of mass, and with an identical roadmap 41, facing an obstacle 18 similar, having distinct avoidance resources, in terms of response time and vertical acceleration margin Gz in particular.

In other words, we want to take into account the real performance of the aircraft 1, at a given moment, in order to avoid all false alarms and to reduce as much as possible the distances from the planned trajectory 20.

To this end, the conical section 24 defines according to the invention, on the abscissa in the longitudinal direction X, at least part of the avoidance curve CE as a function of a value of attainable speed (by permissible deceleration) by the aircraft 1 at the end of the transfer time 23, and on the y-axis as a function of the vertical acceleration capacity Gz of this aircraft 1, at the end of the transfer time 23.

Consequently, the conical curve 24 is relatively close to the planned trajectory 25 and therefore to the longitudinal direction X (called “horizontal”), if the maneuverability of the aircraft 1 is low. This necessarily results in an extension of the anticipation distance 22.

Conversely, the conical curve 24 is momentarily capable of strongly deviating from the planned trajectory 25 and therefore of approaching the direction of elevation Z (called "vertical") going upwards, if the maneuverability of the aircraft 1 is high. This necessarily results in a shortening of the anticipation distance 22.

This ability of the aircraft 1 to momentarily diverge in nose-up from the planned trajectory 25, is translated significantly, by an increase in the angle 26 of the collective pitch of the lift and propulsion rotor 2.

In particular, this value of increasing the angle 26 is called the collective pitch margin 27. These are shown schematically on figure 1 .

Such maneuverability parameters, namely the collective pitch 26 and the collective pitch margin 27, are judiciously obtained by virtue of a preferred embodiment of the invention.

To obtain the available vertical acceleration margin Gz, the piloting aid device 10, for example a TAWS, is coupled from a logical point of view with the avoidance warning system 13, for example an IPL. This is represented by the arrow 28 on the figure 4 and requires little or no additional wiring, and the additional processing means to be provided in this case are most often limited to code 15 for programming the computer 14.

In the equations for calculating the avoidance curve 24, this available margin of vertical acceleration Gz is denoted by ΔGz (delta of Gz). It turns out that from the collective pitch 26, ΔGz defines an increase 27 of this pitch angle, on the blades of the rotor 2.

This represents a value which is both approximate but admissible, such that the angle 27 corresponds to: ΔGz = (K x Gz), where is a coefficient K which represents the instantaneous mass of the aircraft 1 at the time of calculation, c ' that is to say at time T0.

In fact, with the calculation of the coefficient K as a function of the mass of the aircraft 1, and since the margin 27 or ΔGz available for vertical acceleration Gz reflects the force Fz (see figure 3 ) according to the direction of elevation Z that the rotor 2 is capable of developing, it is possible to obtain a significant value of the vertical acceleration Gz, from the parameters introduced from the system 13 to the device 10.

From this instantaneous value of Gz, the latter is introduced into a conical function 24, in order to provide the avoidance curve CE according to the invention.

In one embodiment, the instantaneous value of Gz is fed into a conical function 24 and provides a draft avoidance curve CE, which is then adjusted as a function of maneuverability parameters or additional contextual data.

Note that the conical function 24 of the avoidance curve CE is thus a function of the power margin, or at least of vertical acceleration Gz, of the aircraft 1, at the instant T0.

From this instant T0, the proximal transfer section 25 and the conical function 24 of the avoidance curve CE can be deduced according to the invention. The sum of the projections 23 of the section 25, and of a projection 29 of the conical avoidance curve 24 on the longitudinal direction X is visibly less than the distances TT and 21 obtained with conventional technologies.

• son environnement de vol (température et pression atmosphérique ambiantes, altitude, conditions atmosphériques, visibilité, etc.) ;
• sa phase de vol (décollage, croisière, approche, posé, etc.) ;
• son état fonctionnel initial lors du vol donné (état de maintenance, vieillissement, remplissage réservoirs, chargement à bord, équipements embarqués, etc.) ;
• son état instantané (paramètres de fonctionnement à un instant donné, tels que température / pression des fluides et flux, masse totale de l'appareil, puissance motrice disponible, mode de pilotage à vue ou aux instruments, etc.) ;
• sa feuille de route 41 (mission civile ou militaire, tactique ou de simple transport, sauvetage, etc.).

In embodiments, various parameters designated at 30 on the figure 4 , are taken into account and integrated into the evaluation of the avoidance trajectory TA specific to the invention, because they influence the maneuverability of the aircraft 1, including:

• its flight environment (ambient temperature and atmospheric pressure, altitude, atmospheric conditions, visibility, etc.);
• its phase of flight (take-off, cruise, approach, touchdown, etc.);
• its initial functional state during the given flight (state of maintenance, aging, filling tanks, loading on board, on-board equipment, etc.);
• its instantaneous state (operating parameters at a given instant, such as temperature / pressure of fluids and flow, total mass of the device, available motive power, visual or instrument piloting mode, etc.);
• its roadmap 41 (civil or military mission, tactical or simple transport, rescue, etc.).

The integration of parameters 41, 17, 19 and 30 is shown schematically by arrow 31 on the figure 4 . From a logical point of view, this is a coupling between the piloting aid device 10, for example a TAWS, with the warning system 13 maneuverability indicator.

Referring to the figure 5 , an implementation of the method according to the invention is shown schematically and summarized below.

In this example, instantaneous parameters such as the temperature 32 at the engine 44 of the aircraft 1, the pressure 33 at this engine 44 as well as the torque 34 effective at the rotor 2, are logically injected into the system 13 maneuverability indicator 13 , for example an IPL.

This method is iterative if necessary, and the injection of parameters 32 to 34 is the step of starting a logic loop, at time T0.

On the basis of these parameters 32, 33 and 34 in particular, the maneuverability indicator system 13 calculates an instantaneous value of the available power margin, designated at 35. We have seen above how this is operated according to the invention.

In a step 36 (illustrated by an integrating arrangement also designated as 36), so-called static parameters 37 are integrated, and in particular parameters 37 significantly reflecting the model of the aircraft 1 (stored within the equipment 9 , for example via the computer 14 or a link 11).

It is also at this step 36 that, as illustrated by arrow 31, other significant parameters are integrated, including the flight plan 19.

Step 36 also allows the production of the transfer time TT = 23, and therefore of the proximal section 25.

At a subsequent step 38, a collective pitch margin 27 is deduced, attainable at instant T0, by the aircraft 1. We have seen that by satisfactory approximation, it could be said that the vertical ascent force Fz can be developed by the rotor 2 is significant, or even equal, to K times Gz, this being a function of the collective pitch margin 27. This corresponds to the equation: Fz = (K. Gz).

Then, in a step 39, the proximal 25 and curvilinear 24 sections of the avoidance curve CE are defined, that is to say the avoidance trajectory TA (which may optionally be adjusted subsequently).

This trajectory TA is developed so as to correspond to the following equation: $$\mathrm{YOUR}=\left(\mathrm{TT}\right)+1/2\mathrm{<figure-callout\; id="2"\; label="Gz\; TT"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">Gz\; </figure-callout>}\left(\mathrm{TT}\right)\left(\right)2.$$

Where the time TT is equal to the calculated duration 23.

At a subsequent step 40, the results of this equation are estimated on the assumption of a transient avoidance curve (not applied other than as a transient calculation value) which is circular, in order to deduce a value R which defines a radius of this transient avoidance curve.

Où : W1 (Oméga) est l'accélération de l'aéronef 1, et V1 sa vitesse.We then read: $$\left(\mathrm{W}1\right)2\mathrm{time}\mathrm{R}=\mathrm{Gz}=\left(\mathrm{V}1\right)2/\mathrm{R};$$

Where: W1 (Omega) is the acceleration of aircraft 1, and V1 its speed.

There is then an additional approximation, following this first calculation, to say that since: (W1) = (V1) / R;
We obtain a conical curve 24 such that: $$\mathrm{R}=\left(\mathrm{V}1\right)2/\mathrm{Gz}=\left(\mathrm{V}1\right)2/\mathrm{value}\mathrm{of}27\left(\mathrm{margin}\mathrm{not}\mathrm{collective}\right).$$

Note that if R is infinite, the margin 27 is non-existent, that is to say zero.

As we have seen, to avoid false alarms produced in rotorcraft by current TAWS, especially in tactical flight, it is useful for the distance that defines the danger zone to be as short as possible, while maintaining maximum safety. .

Let's bring some additional information. The mechanical power P (Vz), necessary for the aircraft 1 to produce the climb force Fz, is equal to the sum of the forward power (in the direction X) with the climb capacity, noted:
P (Vz) = P (Vx) + (Fn. Vz / 2), knowing that Fn is the normal force equal to the product of the mass by gravity, that is (Fn = Mg).

où :

• NR0 est la vitesse de rotation du rotor 2 au temps T0 et NR à l'obtention de la force Fz visée ;
• (Col.P0) est le pas collectif du rotor 2 au temps T0;
• (Col.P) est le pas collectif du rotor 2 à l'obtention de la force Fz visée ;
• A, B et C sont des constantes dépendant de la vitesse d'avancement Vx de l'aéronef 1.

In addition, for the selection between step and power, it is possible to start from the equation: $$\mathrm{W}=\mathrm{AT}+\mathrm{B}\left[\left(\mathrm{Collar}.\mathrm{P}\right)\mathrm{-}\left(\mathrm{Collar}.\mathrm{P}0\right)\right]2.\left[\mathrm{NR}/\mathrm{NR}0\right],$$

or :

• NR0 is the speed of rotation of rotor 2 at time T0 and NR when the target force Fz is obtained;
• (Col.P0) is the collective pitch of rotor 2 at time T0;
• (Col. P) is the collective pitch of rotor 2 when the target force Fz is obtained;
• A, B and C are constants depending on the forward speed Vx of the aircraft 1.

As a first approximation, we can say that the current collective pitch (Col. P0) applied corresponds to developing the necessary power P (Vx) for forward flight, and therefore that the power margin will result in an ascent speed equal to: (Fn. Vz) / 2.

According to the formula for the power P (Vz), the power margin is related to the collective pitch margin [(Col. P) - (Col. P0)] in the form of a proportionality squared of the margin of not collective. However, this collective pitch margin is provided according to the invention, by the maneuverability indicator system 13.

où (M0) est le couple à l'instant (T0).Note that if only values in percent (%) of the collective pitch margin are available, for example at the output of an IPL, and that values in the form of an angular value or other physical unit are desired, it is possible to relate the power to the torque margin, according to the equation: $$\mathrm{W}=\mathrm{K}\left(\mathrm{NR}\right).\left(\mathrm{M}0\right),$$

where (M0) is the torque at time (T0).

In one implementation, the system 13 comprises a logical link with a redundant full authority automatic regulator of the engine 44 (such as a “Full Authority Digital Engine control” or FADEC) of the aircraft 1, the latter delivering a value of available torque margin, after having transformed the available margins (temperature 32 or pressure 33 for example) into instantaneous torque value via the mathematical model of said motorization 44.

In such an embodiment, the engines 44 are controlled and regulated by this control and regulation device provided with an electronic regulation computer, called FADEC, which determines in particular the position of the fuel metering device as a function of a loop. of regulation comprising a primary loop based on maintaining the speed of rotation of rotor 2 of rotorcraft 1, and on the other hand a secondary loop based on a set value of the piloting parameter.

A FADEC also receives signals relating on the one hand to the monitoring parameters of the motorization 44 that it controls, and, on the other hand to the monitoring parameters of important components of the rotorcraft 1 such as the speed of rotation of the rotorcraft 1. main rotor 2 for advancement and lift, for example.

So here, the FADEC is part or even constitutes the maneuverability indicator system 13, in order to participate in the supply to the device 10 of the parameters and data which are necessary for it. In particular, the FADEC is integrated into the computer 14 and therefore into the on-board equipment 9.

Consequently, the system 13 then transmits the values of the monitoring parameters to a control and regulation display, arranged in the cockpit of the rotorcraft 2, via a digital link.

With reference to the document FR2749545 , this display may include a first limitation instrument which identifies and displays a limiting parameter, namely the monitoring parameter closest to its limit.

It should be noted that the FADEC can possibly determine this limiting parameter, the first limitation instrument then contenting itself with the display.

Finally, the FADEC is able to trigger a plurality of alarms if incidents occur, a minor failure or total fuel regulation of the engine 44 for example.

In addition, the FADEC sends information to the display system via a digital link when a turbine engine monitoring parameter exceeds a limit predetermined by the engine manufacturer.

où (Δpas) est ladite variation de pas.Furthermore, we know that any increase in pitch results in a vertical force of the rotor 2, which instantly corresponds to an acceleration oriented in the Z direction, and according to the formula: $$\mathrm{Gz}=\mathrm{K}.\left(\mathrm{\Delta }\mathrm{not}\right);$$

where (Δpas) is said variation in pitch.

où (ΔS) est la marge de pas fournie.Therefore, if we use the maximum step margin calculated by the system 13, for example an IPL, we obtain: $$\mathrm{Gz}=\mathrm{K}\prime .\left(\mathrm{\Delta S}\right);$$

where (ΔS) is the step margin provided.

• une phase égale au tronçon proximal 25, correspondant à un vol en pallier ;
• sur la courbe conique, une phase de prise d'altitude avec une accélération sensiblement de l'ordre de la valeur de Gz ; et
• une phase pseudo-rectiligne, avec une vitesse Vz sensiblement constante et lors de laquelle l'aéronef utilise la puissance motrice maximale disponible.

This makes it possible to identify three distinct phases, successive and contiguous within an approximation of the avoidance trajectory, with a view to the calculation according to the invention of the final trajectory TA, namely:

• a phase equal to the proximal section 25, corresponding to a level flight;
• on the conical curve, an altitude gain phase with an acceleration substantially of the order of the value of Gz; and
• a pseudo-rectilinear phase, with a substantially constant speed Vz and during which the aircraft uses the maximum engine power available.

This approximation more faithfully represents the real avoidance capacity of the aircraft 1. Since the margin delivered by the system 13 is used, this approximation reflects the instantaneous reality by including all the parameters of mass and environment, as well as aging of the engine 44.

In addition, if the aircraft 1 has a large amount of margin, this makes it possible to avoid a terrain alert, for example in a tactical flight. On the other hand, if the aircraft 1 is already at the power limit at the instant T0, the terrain analysis will be automatically carried out over a distance 22 further towards the front of the aircraft 1.

The switching from a so-called short distance 22 chosen when the available power was high, to a distance 22 further ahead of the aircraft 1 when the available power has a value less than a predetermined limit, and vice versa, is carried out in real time by implementations of the invention. This constitutes a step of the method of the invention, in this implementation.

In practice, with the invention, this anticipation distance switching step 22 is carried out as a function of the mapping 17 within which the interactions of obstacles 18 are sought in two calculation sectors within d 'a maximum anticipation distance value, with a large margin of application of a calculation.

During a first phase, we consider a trajectory TA, with the margin (ΔS) = K "(Gz). The available pitch is evaluated at the junction between the proximal section 25 and the conical curve 24.

• suivant la direction X, le mouvement Mx est : (Vx) fois la durée (Dx) prévue pour être écoulée entre T0 et le temps à la jointure entre le tronçon proximal 25 et courbe conique 24, soit : $$\left(\mathrm{Mx}\right)=\left(\mathrm{Vx}\right).\left(\mathrm{Dx}\right);$$
  
• suivant la direction Z, le mouvement (Mz) est : 1 / 2 (Gz) fois le carré de la durée prévue pour être écoulée entre T0 et le temps à la jointure entre le tronçon proximal 25 et courbe conique 24, soit : $$\left(\mathrm{Mz}\right)=\left(1/2\right).\left(\mathrm{Gz}\right).\left(\mathrm{<figure-callout\; id="2"\; label="Dx"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">Dx</figure-callout>}\right)2.$$
  

The movement is put into an equation:

• following the direction X, the movement Mx is: (Vx) times the time (Dx) expected to have elapsed between T0 and the time at the junction between the proximal section 25 and the conical curve 24, that is: $$\left(\mathrm{Mx}\right)=\left(\mathrm{Vx}\right).\left(\mathrm{Dx}\right);$$
  
• following the Z direction, the movement (Mz) is: 1/2 (Gz) times the square of the time expected to have elapsed between T0 and the time at the junction between the proximal section 25 and the conical curve 24, i.e.: $$\left(\mathrm{Mz}\right)=\left(1/2\right).\left(\mathrm{Gz}\right).\left(\mathrm{<figure-callout\; id="2"\; label="Dx"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">Dx</figure-callout>}\right)2.$$
  

So, we can write that the movement following the direction Z, is equal to: 1/2 (Gz) times (1 / Vx2), multiplied by the value obtained from the movement following X, that is to say: $$\left(\mathrm{Mz}\right)=\left(1/2\right).\left(1/\mathrm{<figure-callout\; id="2"\; label="Vx"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">Vx</figure-callout>}2\right).\left[1/\left(\mathrm{K}″.\mathrm{Mx2}\right)\right].\left(\mathrm{\Delta S}\right).$$

Such an equation defines a sector of conical curve, here a parabola, the characteristic of which is linked to the margin expressed in collective pitch.

It is deduced from this that at the end of a period (T), the aircraft 1 has reached an ascent speed (Vz), such that: $$\left(\mathrm{Vz}\right)=\left(\mathrm{Gz}\right).\left(\mathrm{T}\right)\mathrm{is}:\left(\mathrm{T}\right)=\left(\mathrm{Vz}\right):\left(\mathrm{Gz}\right)=\left[\left(\mathrm{K}″\right).\left(\mathrm{\Delta S}\right)\right].\left(\mathrm{Vz}\right).$$

Therefore, at this time (T), it is considered that a secondary phase has been reached, the power available during this secondary phase being low or even non-existent. Therefore, the speed (Vz) is balanced, which corresponds to the ascent at constant speed mentioned above.

où : (MT) est le couple utilisé à l'instant (T).By referring then to the preceding equations, we can pose: $$\left(\mathrm{K}\right).\left(\mathrm{NR}\right).\left(\mathrm{MT}\right)=\left[\left(\mathrm{Mg}\right).\left(\mathrm{Vz}\right)\right];$$

where: (MT) is the torque used at time (T).

et utiliser ces données pour le secteur d'ascension stabilisée.Knowing that the system 13 is able to provide the noted available torque margin (ΔMT), we can obtain: $$\left(\mathrm{Vz}\right)=\left[\left(\mathrm{K}.\mathrm{NR}\right).\left(\mathrm{\Delta }\mathrm{MT}\right)\right].2/\left(\mathrm{Mg}\right)=\mathrm{K}\left(\mathrm{\Delta }\mathrm{MT}\right),$$

and use this data for the stabilized ascension sector.

tel que $$\left(\mathrm{KT}\right)=\left[\left(\mathrm{K}"\right)/2.\left(\mathrm{Mg}\right)\right].\left(\mathrm{K}\right).\left(\mathrm{NR}\right).$$

In other words, we draw a conic section, here in parabola, for the first phase, up to time T during which: $$\mathrm{T}=\left[\left(\mathrm{K}"\right)/\left(\mathrm{\Delta S}\right)\right]=\left[\left(\mathrm{KT}\right).\left(\mathrm{\Delta }\mathrm{S}/\mathrm{\Delta }\mathrm{MT}\right)\right],$$

such as $$\left(\mathrm{KT}\right)=\left[\left(\mathrm{K}"\right)/2.\left(\mathrm{Mg}\right)\right].\left(\mathrm{K}\right).\left(\mathrm{NR}\right).$$

$$\left(\mathrm{Mz}\right)=\left(\mathrm{K}\prime \right).\left(\mathrm{\Delta S}\right).\left(\mathrm{Dx}\right)2.$$

In the end, we obtain until time T, a phase of the first sector in conic, here parabolic, with: $$\left(\mathrm{Mx}\right)=\left(\mathrm{Vx}\right).\left(\mathrm{Dx}\right)$$

$$\left(\mathrm{Mz}\right)=\left(\mathrm{K}\prime \right).\left(\mathrm{\Delta S}\right).\left(\mathrm{<figure-callout\; id="2"\; label="Dx"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">Dx</figure-callout>}\right)2.$$

et $$\left(\mathrm{Mz}\right)=\left[\left(\mathrm{K}\prime \right).\left(\mathrm{\Delta }\mathrm{S}.\mathrm{T}\right)\right]2+\left[\left(\mathrm{K}"\right).\mathrm{\Delta }\mathrm{MT}.\left(\mathrm{T}0-\mathrm{T}\right)\right]2.$$

For the successive phase of the second sector, the ascent takes place at the speed (Vz) = (K "). (ΔMT);
is $$\left(\mathrm{Mx}\right)=\left[\left(\mathrm{Vx}\right).\left(\mathrm{Dx}\right)\right]$$

and $$\left(\mathrm{Mz}\right)=\left[\left(\mathrm{K}\prime \right).\left(\mathrm{\Delta }\mathrm{S}.\mathrm{<figure-callout\; id="2"\; label="T"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">T</figure-callout>}\right)\right]2+\left[\left(\mathrm{K}"\right).\mathrm{\Delta }\mathrm{MT}.\left(\mathrm{T}0-\mathrm{<figure-callout\; id="2"\; label="T"\; filenames="imgf0001.png,imgf0002.png"\; state="\{\{state\}\}">T</figure-callout>}\right)\right]2.$$

An additional possibility of improvement is provided, in an implementation of the invention. To facilitate the calculations of interaction with the terrain 16 via the mapping 17, it is possible to use a protection zone in the form of a linear torso zone.

Such a linear torso zone rests on a parallel to the longitudinal direction and to the planned trajectory 20, but draws a broken line instead of the curvilinear path obtained with the preceding calculations.

From an initial time T0, the intersection P is sought, between said parallel to the longitudinal direction and to the planned path 20 and a tangent 43 to said curvilinear path. This is illustrated schematically on the figure 3 .

d'où $$\mathrm{T}1=\mathrm{T}0-\left[\left(\mathrm{K}\prime .\mathrm{\Delta }\mathrm{S}.\mathrm{T}0\right)\right]2/\left[\left(\mathrm{K}"\right).\mathrm{\Delta }\mathrm{MT}\right];$$

de sorte que : $$\mathrm{xP}=\left(\mathrm{Vx}\right).\mathrm{T}1.$$

This point P has a position (xP; zP) such that: $$\mathrm{zP}=0;$$

from where $$\mathrm{T}1=\mathrm{T}0-\left[\left(\mathrm{K}\prime .\mathrm{\Delta }\mathrm{S}.\mathrm{T}0\right)\right]2/\left[\left(\mathrm{K}"\right).\mathrm{\Delta }\mathrm{MT}\right];$$

so that : $$\mathrm{xP}=\left(\mathrm{Vx}\right).\mathrm{T}1.$$

In other words, on parallel to the longitudinal direction and to the predicted path 20, the distance from the origin to xP is the margin along the direction X.

The formula T1 = T0 - [(K '. ΔS. T0)] 2 / [(K "). ΔMT] involving the ratio between the output margin of the system 13 and the torque margin of the motorization 44, the choice of the linear torso zone instead of the curvilinear path initially calculated, remains perfectly linked to the instantaneous maneuver margins of the aircraft 1. In fact, such a rectilinear torso zone forms a significant and coherent calculation relief.

Indeed, this formula can easily be reduced either in terms of margin of the system 13, or in terms of torque margin, depending on the type of system 13 used by means of the aircraft models 1 used by this system 13.

We have seen that the invention assigns a pilot reaction time duration, for example determined as a function of the type of flight in progress (eg military or civilian / cruise or flight phase attentive). This results in a delimitation segment proximal to the rotorcraft 1, of the danger zone, which is not exclusively proportional to the speed.

Furthermore, the direction (nose up / nose down) of the speed vector of the rotorcraft 1, as well as the rotorcraft resources available at a given instant, are incorporated into the calculations for delimiting the danger zone by the invention.

To do this, a proposed solution consists of coupling, from a logical point of view, the TAWS to the IPL of rotorcraft 1.

In fact, the IPL translates the resources of the rotorcraft 1, available at a given moment, in particular in terms of power, in the form of a collective pitch. In fact, it is possible to deduce at the given instant, the vertical acceleration, the mass and the direction of the speed vector of rotorcraft 1.

In particular, the IPL which can be used can correspond to the teaching of the document FR2756256 which describes a power margin indicator where, on the basis of control parameters and values of the use limitations of the motorization 44, a power margin indicator is produced, expressed in particular as a collective pitch value.

From these deductions from the IPL and / or the FADEC, the adaptive TAWS calculates a shortened danger zone, delimited by a conical curve, while maintaining maximum safety.

• de produire une valeur limitée (temps de réaction du pilote court, de l'ordre par exemple de moins d'une seconde pour un vol attentif, à moins de deux secondes pour un vol de croisière, caractérisé par un segment sensiblement proportionnel à la vitesse de l'appareil) de transfert homogène à une durée, c'est-à-dire un temps, qui serait la plus limitée que possible (par exemple en fonction du type de vol, de la phase de vol, de données historiques et d'aptitudes personnelles du pilote dans un pareil contexte) ; et
• d'en déduire une courbe dite pseudo-conique (c'est-à-dire dont la projection dans un plan sensiblement parallèle à une direction longitudinale de l'aéronef et sécante à sa trajectoire à son origine, dessine au moins un secteur de courbe conique, telle que de parabole) d'évitement, qui serait quant à elle, liée notamment à la manœuvrabilité de l'aéronef 1 à voilure tournante, en temps réel.

One approach would include:

• produce a limited value (short pilot reaction time, for example of the order of less than one second for attentive flight, less than two seconds for a flight of cruise, characterized by a segment substantially proportional to the speed of the aircraft) of homogeneous transfer at a duration, that is to say a time, which would be as limited as possible (for example according to the type of flight, the phase of flight, historical data and personal skills of the pilot in such a context); and
• to deduce a so-called pseudo-conical curve (that is to say whose projection in a plane substantially parallel to a longitudinal direction of the aircraft and secant to its trajectory at its origin, draws at least one curve sector conical, such as a parabola) avoidance, which would be linked in particular to the maneuverability of the rotary-wing aircraft 1, in real time.

The invention is however not limited to the embodiments described.

Claims (8)

1. Method for elaborating a warning for terrain avoidance by a rotary wing aircraft (1);said method comprising elaborating an avoidance trajectory (TA) that comprises a proximal segment (25) and an avoidance curve (CE), and issuing a warning at the start of the proximal segment (25);
   the proximal segment (25) extending in the continuation of a planned trajectory (20) over a distance which translates into an applicable reaction time (23), the applicable reaction time (23) being calculated by determining an initially applicable time value which is assigned to the model of the aircraft (1), and by carrying out limitation weighting, which is predefined by the roadmap (41) of the aircraft (1), on the initially applicable time value in order to minimise the applicable reaction time (23); and
   the avoidance curve (CE) comprising at least one distal segment (24) that connects the proximal segment (25) and that is calculated by means of a conic function using at least one possible vertical acceleration value (Gz) determined from an increase (27) in the collective pitch angle of one or more propulsion and lift rotors (2) of the aircraft (1).
2. Method according to claim 1, characterized in that the proximal segment (25) is substantially straight.
3. Method according to either claim 1 or claim 2, characterized in that the limitation weighting is operated by dividing by at least one limiting ratio predefined by the roadmap (41) of the aircraft (1).
4. Method according to any of claims 1 to 3, characterized in that the distal segment (24) is a conic curve, the conic curve (24) being of a proper type, such as a parabola, ellipse or hyperbola.
5. Terrain warning device (10) for implementing the method according to any of claims 1 to 4, characterized in that said device (10) is logically coupled to a system (13) that indicates manoeuvrability.
6. Device (10) according to claim 5, characterized in that said device (10) is at least partly on-board, and comprises a flight apparatus (9) that has a flight computer (14) capable of executing a code (15) which allows the method according to any of claims 1 to 4 to be implemented.
7. Rotary wing aircraft (1) for implementing the method according to any of claims 1 to 4, characterized in that it (1) is a helicopter (3) or a convertible rotary wing aircraft (5) or a drone.
8. Rotary wing aircraft (1), characterized in that it (1) is capable of implementing the method according to any of claims 1 to 4 and/or in that it (1) comprises a terrain warning device (10) according to either one of claim 5 or claim 6; said aircraft (1) having an audible alarm (45) intended for being selectively triggered by the terrain warning.

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