METHOD AND DEVICE FOR DETERMINING AND OPTIMIZING CHARACTERISTIC PARAMETERS OF THE OPERATION OF A ROT

专利摘要:
The present invention relates to a method for determining characteristic parameters of the operation of an aircraft (10) comprising a powerplant (20) provided with at least one motor (21,22) and a mechanical transmission means (23). sensors and display means (4). During this process, various information relating to said aircraft (10), its state and / or its operation and / or its environment is measured, and then, for at least one parameter Pi relative to the state and to the operation of said aircraft, is determined. (10), a first limit value Pi_lim of said parameter Pi, a second value Pi_X of each parameter Pi so that said aircraft (10) can perform a predetermined operation X, and a third instantaneous value Pi_inst of each parameter Pi. simultaneously each first, second and third value Pi_lim, Pi_X, Pi_inst to highlight the relative position of a second and a third values Pi_X, Pi_inst vis-à-vis a first value Pi_lim for each parameter Pi.
公开号:FR3033316A1
申请号:FR1500411
申请日:2015-03-04
公开日:2016-09-09
发明作者:Serge Germanetti
申请人:Airbus Helicopters SAS；
IPC主号:

专利说明:

[0001] 1 Method and device for determining and optimizing characteristic parameters of the operation of a rotary wing aircraft. The present invention is in the field of piloting assistance for vehicles, and rotary wing aircraft in particular. The present invention relates to a method and a device for determining and optimizing parameters characteristic of the operation of a vehicle and in particular of a rotary wing aircraft.
[0002] The operation of a vehicle is generally carried out under the supervision of several characteristic parameters by means of several instruments located on a dashboard of the vehicle. These characteristic parameters are representative of the current operation of the vehicle and in particular of its engine or of its powerplant. For physical reasons, there are many limitations on these characteristic parameters which must be taken into account at each moment of the operation of the vehicle. These different limitations may depend on external conditions as well as the mode of operation of the vehicle. For example, for a motor vehicle, these characteristic parameters may be the speed of rotation of its engine, the temperature of the cooling water of the engine or the temperature of the lubricating oil of the engine.
[0003] According to another example, the vehicle may be a rotary wing aircraft comprising a powerplant provided with two turbine engines and a main power transmission gearbox, the powerplant driving in rotation at least one main rotor and possibly one main rotor. anti-torque rotor, like a rear rotor. In this case, these characteristic parameters may be, inter alia, the rotation speed Ng of the gas generator of each turbine engine, the gas discharge temperature T4 at the inlet of the free turbine of each turbine engine and the engine torque. Cm of each turbine engine. Thus, the pilot of a vehicle must monitor, during the operation of this vehicle and permanently, the current values of these characteristic parameters on several instruments located on the dashboard of the vehicle and compare the current values of these characteristic parameters. with their respective limitations. In the case of a rotary wing aircraft comprising a powerplant equipped with two turbine engines, the pilot must monitor, during the operation of the aircraft and permanently, at least three instruments per turbine engine, ie at least six instruments. This pilot is then supposed to detect, for each turbine engine, an inconsistency between the current values of these characteristic parameters and their respective limitations on these at least six instruments. This requires the pilot's particular attention, which must also be concentrated on the actual flight of the aircraft. In addition, these limitations are generally different according to the flight phase of the aircraft and / or the external conditions such as altitude for example. Indeed, according to each phase of flight and / or external conditions and the mode of operation of the power plant, the maximum power that can provide the power plant is different and its duration of availability can also be limited.
[0004] For example, for the take-off phase of the aircraft, a maximum take-off power PMD can be used for a limited period of the order of five to ten minutes, corresponding to a torque level for the main transmission gearbox. 5 power and heating of each turbine engine eligible without significant degradation of the power plant. Similarly, after the take-off phase, a continuous maximum power PMC can be used continuously, without limitation of duration. In addition, there are also emergency power-up regimes used on a power plant with at least two turboshaft engines when one of the turboshaft engines is down. The valid turbine engine (s) can then provide emergency powers for limited periods of time, these emergency powers being greater than the maximum continuous power PMC in order to compensate for this failure. However, the use of these emergency powers then generally require specific maintenance operations. As a result, the limitations of the various parameters characteristic of the operation of the aircraft can be different, in particular according to the power available at each turbine engine. Today, some help to the driver can limit the parameters that must monitor the driver. Document FR2749545 and FR2756256 which describe a First Limitation Instrument often referred to by the acronym "IPL" are particularly known. This First Limiting Instrument identifies, among various characteristic parameters, the critical characteristic parameter as being the one that is closest to its limit value. The current value of this critical characteristic parameter and its limit value are then grouped together on a single display, respectively for each turbine engine, if necessary, making it possible to limit the number of instruments necessary to monitor the operation of the aircraft in order to simplify the pilot's task.
[0005] These IPLs thus make it possible to display, by means of the current value of the critical characteristic parameter and its limit value, a power margin available to the aircraft or each turbine engine. For example, the current value of the critical characteristic parameter and its limit value can be displayed according to a graduated scale in engine torque for each turbine engine thus characterizing the available power margin of each turbine engine of the aircraft, as described in document FR2749545. The current value of the critical characteristic parameter and its limit value can also be displayed according to a graduated scale in collective pitch, the collective pitch indicating the incidence of the blades of the main rotor of the aircraft relative to the incident wind induced by the control of collective pitch of these blades, thus characterizing the available power margin of the aircraft 20 as a whole, as described in the document FR2756256. However, the limitations used by the instruments of the dashboard and this IPL in particular do not represent the actual limitations of each turbine engine, but predetermined limitations corresponding for example to a minimum guaranteed power of a turbine engine over its life. life. Indeed, the engine maker establishes, by calculations or by tests, the available power curves of a turbine engine depending in particular on the altitude of the aircraft and the outside temperature, and this for each of the power regimes 30 usable by each turbine engine. In addition, the motorist 3033316 5 determines these available power curves according to different levels of aging of each turbine engine between a new turbine engine and a turbine engine arrived at the end of its life. As a result, a minimum guaranteed power over the entire life of a turbine engine is defined. The value of this guaranteed minimum power is variable, in particular according to the altitude of the aircraft and the outside temperature, and corresponds to the power supplied by an aged turbine engine, namely a turbine engine having reached its maximum life time. Thus, any turbine engine in normal operation, that is to say not undergoing any failure, can always deliver a higher power and at least equal to the guaranteed minimum power over its entire life. In this way, the instruments of the dashboard and the IPL in particular which use limitations corresponding to this guaranteed minimum power are favorable in terms of safety, the pilot always having a power actually available at each turbine engine which is generally greater and at least equal to the maximum power indicated by the instruments of the dashboard or the IPL. Conversely, the use of each turbine engine is not optimized, the power used being the guaranteed minimum power and not the maximum power actually available. In fact, each turbine engine is underutilized. The use of limitations corresponding to the maximum power actually available would make it possible to improve the performance of the aircraft such as the total mass transported or the range, for example. In addition, the dashboard instruments of a vehicle as well as the IPL for the aircraft indicate the current values and the limits of the characteristic parameters. In fact, when a pilot is planning to perform a maneuver, he must rely on his experience and the difference between these current values and their limits in order to estimate whether he has a sufficient margin. on the characteristic parameters to achieve this maneuver. Then, the pilot will have confirmation only during this maneuver that no characteristic parameter exceeds his limit and that he can thus realize this maneuver in full safety. In the opposite case and according to the maneuver performed, the pilot can stop this maneuver to return to a safe flight phase, each characteristic parameter then remaining below its limit. This is typically the case when the pilot of an aircraft engages a descent and can use the aircraft's inertia and / or the total power available at the power plant to perform a maneuver. avoidance of an obstacle. However, for some maneuvers, backtracking is impossible once the maneuver engaged and an accident can occur then, for example during a landing and transition off ground effect and ground effect.
[0006] This poor estimation of the pilot of the available margin, especially in terms of power, is at the origin of many accidents with rotary wing aircraft and in particular during the landing phases, of stationary flights, particularly near the ground. and takeoff in pure vertical mode.
[0007] The object of the present invention is therefore to propose a method and a device making it possible to overcome the abovementioned limitations by allowing the pilot to simultaneously display the current values and the actual limits of at least one characteristic parameter of the operation of this vehicle, as well as the value of each characteristic parameter necessary for the realization of a predetermined maneuver. An object according to the present invention is a method for determining characteristic parameters of a vehicle, the vehicle comprising a powerplant provided with at least one motor and a mechanical power transmission means, displacement control means. of the vehicle, a plurality of sensors, at least one calculation means, at least one memory and at least one display means.
[0008] The vehicle is for example a rotary wing aircraft comprising a main rotor provided with main blades, an anti-torque rotor provided with secondary blades and a powerplant. The mechanical power transmission means is a main power transmission gearbox rotating the main rotor and the anti-torque rotor through at least one engine which may be a turbine engine. The control means are control means for the variation of the collective pitch and the cyclic pitch of the main rotor blades of the main rotor as well as the collective pitch of the secondary blades of the anti-torque rotor.
[0009] During this method of determining characteristic parameters of a vehicle, various information relating to the environment of the vehicle, and / or to the state and operation of the vehicle and / or to the state and operation of the power plant 25 and / or the state of the control means, and / or the position and the movements of the vehicle relative to its environment, -on determines at least a first value P, Iim equal to one limit value that a parameter P must not exceed, e being a characteristic parameter relating to the state or the operation of the vehicle, or to the state or to the operation of the powerplant, or to the state control means, or at the position or displacements of the vehicle relative to its environment, at least one second value PLX is determined at which the parameter P1 must be equal so that the vehicle can perform a predetermined maneuver X and simultaneously displaying on the same graphical representation each first value and each second value PLX in order to clearly highlight the relative position of a first value Pi lim with respect to each second value Pi_X for each parameter P1. The parameter P1 may be a characteristic parameter relating to the state or operation of the power plant such as a total power or a total torque supplied by the power plant or a torque at the output of the mechanical transmission means. . The parameter Pi can also be a parameter relating to the state or the operation of a motor of the power plant, such as a power or a torque supplied by this engine, a temperature inside this engine or a rotational speed of an element of this engine such as a transmission shaft. The parameter Pi may also be a parameter relating to the position or displacements of the vehicle relative to its environment, for example a speed of the vehicle relative to the air or a speed of the vehicle relative to the ground, or a vehicle height relative to a reference such as the ground. When the vehicle is a rotary wing aircraft, the parameter P may also be, inter alia, a torque of the main rotor or a torque of the anti-torque rotor, a steering position of the aircraft or a step value. collective or cyclic main rotor blades of the main rotor or secondary blades of the anti-torque rotor, a position of the collective pitch or cyclic pitch of the main rotor blades of the main rotor or a position of the collective pitch control of the secondary blades of the rotor tail. The parameter Pi may also be a logical or arithmetic combination of at least two previously cited features. In this way, each parameter Pi makes it possible to characterize the operation or the state of the vehicle and / or its powerplant or a position or speed of the vehicle. Advantageously, each first value PL / im is equal to a current limit value of the parameter Pi and not to a predefined limit. Each first value Pi lim is determined from one or more of the different measured information. This first value Pi lim thus takes into account the current characteristics of the vehicle, its power plant and its environment. This first value can also take into account a state of aging of the vehicle and / or its power plant. Each first value Pi lim can be determined using vehicle operating curves, and more particularly vehicle performance curves, correlated with the power curves of the engines of that vehicle. Advantageously, the vehicle performance curves are defined by taking into account the location of each engine in the vehicle.
[0010] Each second PLX value is equal to a value that the parameter P must take so that the vehicle can perform a predetermined operation X. Each second PLX value can be determined from one or more of the different measured information and characteristics necessary to perform this predetermined operation X. Thus, the difference between the first value Pi lim and the second value PLX is visible on the graphical representation. The driver of the vehicle directly displays if he has a margin on the parameter Pi to perform this predetermined maneuver X vis-à-vis the first value Pi lim. The second values P; _X to which the parameter P1 must be equal so that the vehicle can perform the predetermined operation X can form a range of values of this parameter Pi. This range thus consists of second acceptable values PLX so that the vehicle can perform the maneuver X with margins of safety. These safety margins make it possible, for example, to take into account approximations in the measurement of the various pieces of information, the characteristics necessary for carrying out this predetermined maneuver X or in the knowledge of the mass of the vehicle. Moreover, during this process for determining characteristic parameters of a vehicle, it is possible to determine a third value P, inst equal to an instantaneous value of each parameter P 'and then to display on the graphical representation each third value P , _inst. Thus, the driver of the vehicle directly displays firstly a difference between the third value P, _inst and the first value P, _fim and secondly the margin available to the pilot between the third value 3033316 11 PLinst and the second value PLX necessary to perform the predetermined operation X. In addition, the graphical representation can take several forms. For example, the graphical representation 5 may be a substantially circular shaped dial. A pointer indicates the third value P, _inst and marks outside the dial indicate the first value and the second value PLX. The first value can also be represented by a variable color zone, for example from green to orange 10 and then red, on this dial. The graphical representation may also be in a vertical scale, with the first, second and third values P, lim, PLX, P, _inst being represented by moving horizontal markers or variable color scrolling bands. Each second value P can be displayed; _X and each third value PLinst where appropriate on the vehicle display means according to a percentage of the first value P, _Hm corresponding to the parameter P ,, this first value P; Hm corresponding to a 100% indication on the vehicle display means. This type of display makes it possible to ignore the numerical values of the first, second and third values P; lim, PLX, P, _inst. Thus, the vehicle driver can easily check the consistency of each second value P; However, the first values P, Hm, the second values PLX and the third values P1 inst can be displayed as a percentage of a reference value of this parameter 30 P 'this reference value of this parameter P, corresponding then to a 100% indication on the vehicle display means. For example, since the parameter Pi is the output torque of the vehicle's mechanical transmission means, the reference value of the parameter P may be the limit torque that can be used by this mechanical transmission means of the vehicle in continuous operation. Similarly, since the parameter Pi is the rotation speed Ng of the gas generator of a turbine engine of the vehicle, the reference value of the parameter Pi may be the maximum rotational speed permitted by this gas generator in continuous operation. In addition, when the vehicle is a rotary wing aircraft comprising a main rotor provided with main blades, a control means being a means for controlling the variation of the collective pitch of the main blades, the display means are displayed on the display means. the aircraft the collective pitch values of the main blades respectively corresponding to each first value Pi lim, each second value Pi_Xet and each third value P, _inst if appropriate. The collective pitch setting of the main blades being one of the essential elements of the piloting of a rotary wing aircraft on which the pilot acts directly and permanently, the pilot can thus directly visualize the interaction between the various values related to the parameter Pi. and these values of the collective pitch of the main rotor blades. The correspondence between the collective pitch values and the values P, _lim, P, _X, P; inst can be performed through the power at the main rotor of the aircraft. Furthermore, in order to limit the number of information provided to the driver of the vehicle, it is possible to display information concerning a reduced number of parameters P, these parameters P being the most critical. For example, each first value P, _lim, 3033316 13 is displayed each second value PLXet each third value P, inst for the parameter P1 among at least two parameters Pi, Pj for which a difference between the first value Pi and the third value P , _inst associated respectively with each parameter Pi, P; is the 5 weakest. For example, it is possible to display each first value P.sub.i / im, each second value P.sub.X and each third value P.sub.inst if necessary on a first limitation instrument of the aircraft, whether this display is in percentage of the first value Pi or well according to the collective pitch values of the main rotor blades. One of the determining criteria in the operation of a vehicle is the total power available at its power plant for the operation of the vehicle and its movement in particular. Thus, in order to monitor this total power of the power plant, the parameter P1 can be this total power supplied by the power plant of the vehicle. The parameter Pi can also be the power supplied by each motor of this power plant making it possible to monitor more precisely the power distribution between the motors of the power plant. Moreover, the total power supplied by the power plant or the power supplied by an engine of this power plant is a function of various parameters of this power plant and / or each engine of this power plant. For example, the instantaneous total power Winst supplied by a power plant is substantially equal to the instantaneous power supplied by a motor, when the power plant has only one motor. This total instantaneous power Winst is for example defined according to the formula Winst = Cm-Nnt, where Nm is an instantaneous speed of rotation of the engine transmission shaft and Cm is the engine torque supplied by this engine on this transmission shaft. .
[0011] The instantaneous total power Winst supplied by the power plant when it comprises several motors is substantially equal to the sum of the instantaneous power supplied respectively by each motor of the power plant such as Winst = En [Cm. Generally, a power plant 10 comprises identical motors and the total power supplied by the power plant is provided substantially equally by each motor. In addition, when the vehicle is a rotary wing aircraft equipped with a main rotor and an anti-torque rotor, the total power supplied by the power plant of the aircraft is distributed firstly in a power of flight of the aircraft for the actual flight of the aircraft and secondly in an accessory power to supply equipment of the aircraft. This accessory power is used for example to supply the air conditioning of the cabin of the aircraft, the electrical equipment of the aircraft such as avionics or the hydraulic equipment of the aircraft. This accessory power, composed mainly of electric power and hydraulic power, can be determined in a known manner. The flight power of the aircraft is in turn distributed between the main rotor and the anti-torque rotor to provide lift and movement of the aircraft. This flight power of the aircraft is therefore the sum of the power supplied to the main rotor level and the power supplied to the anti-torque rotor level. The power at the main rotor can be defined in a known manner according to the formula WRp = CR. NR, NR being an instantaneous rotational speed of the main rotor of the aircraft and CR being the rotor torque at the main rotational driving mast of the rotor. The instantaneous power at the anti-torque rotor can also be determined in a known manner, this instantaneous power at the anti-torque rotor being used essentially to oppose a torque due to the reaction of the main rotor of the aircraft to the engine torque used. to rotate this main rotor. For example, this instantaneous power at the level of the anti-torque rotor can be determined according to the speed of advance of the aircraft: when this forward speed of the aircraft is zero, the anti-torque rotor opposes this pair alone, whereas, when this forward speed of the aircraft is non-zero, a transverse aerodynamic force, proportional to the square of this forward speed, is generally generated by a substantially vertical empennage located near the rear rotor, making it possible to reduce this instantaneous power at the anti-torque rotor. In addition, when each engine of the power plant is a turbine engine comprising a gas generator and a free turbine, the power supplied by each turbine engine is a function of the output torque of the turbine engine, an internal temperature T4 gas at the inlet of the free turbine of this turbine engine and the speed of rotation Ng of the gas generator of this turbine engine. The instantaneous power Wminst supplied by each motor can then be defined such that Wminst = Cm.Ng, Cm being the engine torque relative to the gas generator. Consequently, the parameter Pi can be, when the vehicle is a rotary wing aircraft, the rotor torque CR at the level of the main rotor drive mast or the engine torque Cm supplied by a motor of the power plant as well as when this engine is a turbine engine, the internal temperature T4 of the gas at the inlet of the free turbine of a turbine engine or the speed of rotation Ng of the gas generator of a turbine engine. The values of this parameter Pi can then be determined, for example, from the power supplied by each engine of the power plant or from the power of flight of the aircraft. Thus, in order to monitor this parameter Pi according to a variant of the method for determining parameters characteristic of the operation of a vehicle, an available power margin of each engine of the vehicle power plant is determined relative to a guaranteed minimum power, this available power margin characterizing a state of aging of each motor of the power plant, an instantaneous power Wminst supplied by each engine of the power plant is determined, a maximum power available at the power plant is determined; the level of each engine of the power plant taking into account the available power margin, - at least one characteristic power of each motor of the power plant corresponding to the achievement of a predetermined maneuver X of the vehicle is determined, determines at least a first value Pjim corresponding to the maximum power available for each motor r, 3033316 17 - it is determined at least a second value Pi_Xcorresponding to the characteristic power of each motor for the realization of a predetermined operation X of the vehicle, - one determines at least a third value P, inst 5 corresponding to the instantaneous power Wminst provided by each motor, and - simultaneously displaying on the same graphical representation each first value P, lim and each second value PLXet each third value P, inst.
[0012] The available power margin of each engine of the power plant of the vehicle is, for example, determined during a "motor health check", also designated by the acronym EPC for the English designation "Engine Power Check". Such an engine health check is generally recommended by the engine manufacturer and must be carried out regularly, especially when the engine is a turbine engine. An engine health check verifies the operating status of an engine and determines the level of deterioration and / or aging of the engine. A motor health check thus makes it possible periodically to check the maximum and actual performance of the engine in relation to the guaranteed minimum performance. The engine health check is thus performed by comparing the current performance of the engine with engine performance obtained on a test bench and declared by the manufacturer. The engine health check makes it possible to determine a margin of one or more parameters for monitoring this engine with respect to a limit value of each monitoring parameter and, consequently, to determine the available power margin of an engine 30 which is the difference between the maximum power it can provide and its guaranteed minimum power for the current surrounding conditions. Current environmental conditions are, for example, information relating to the environment of the vehicle and more particularly the atmospheric conditions. Moreover, it can be deduced from these monitoring parameters whether the motor has suffered damage and whether it must undergo maintenance operations, in particular in order to be able to supply the mechanical powers for which it is adapted again. For example, if this engine is a turbine engine, a monitoring parameter may be the internal temperature T4 of the gases at the inlet of the free turbine, the rotation speed Ng of the gas generator or the engine torque Cn, delivered by the turbine engine. In addition, each motor health check must be carried out according to a procedure predetermined by the manufacturer. This engine health check can be performed during flight or between two flights. The instantaneous power Wminst supplied by each motor of the power plant can be determined in a known manner, for example by means of the instantaneous torque delivered by the motor. In addition, when this engine is a turbine engine, this instantaneous power can be determined by means of the instantaneous internal temperature T4 of the gases at the inlet of the free turbine of this turbine engine or the instantaneous speed of rotation Ng of its gas generator. Similarly, a characteristic power for each motor 25 of the power plant corresponding to the achievement of a predetermined operation X of the vehicle can be determined. Each characteristic power is necessary for the realization of a predetermined operation X which can in particular be consuming a large mechanical power supplied by each engine.
[0013] This predetermined maneuver X is for example a hover or a landing in the case where the vehicle is a rotary wing aircraft. The maximum power available at each motor of the power plant is determined using vehicle performance curves and taking into account the available power margin with respect to the guaranteed minimum power. The maximum power available is a function of the current environmental conditions and the flight phase of the vehicle when the vehicle is an aircraft. When a power plant comprises a single motor, the power that must provide the engine to perform the predetermined operation X is equal to the sum of the characteristic power of the engine and the accessory power of the vehicle. When a power plant comprises several identical motors, the total power that the power plant must provide to carry out the predetermined maneuver X is equal to the sum of the characteristic power of each motor and the accessory power of the vehicle. This total power supplied by the power plant is generally provided equally by each motor. However, a particular distribution can be defined between each motor of the power plant, especially when these motors are different. As previously, in order to limit the number of information provided to the driver of the vehicle, it is possible to display each first value each second value PLX and each third value P, inst for parameter P, among at least two parameters Pi, P, for which a difference between the instantaneous power and the maximum available power respectively associated with each parameter Pi, P; is the weakest, possibly for each motor of the power plant 5 when it comprises several motors. Moreover, it is possible to use an estimated mass M of the vehicle in order to determine each characteristic power of each motor of the power plant for carrying out a predetermined operation X of the vehicle. This is particularly useful in the case where the vehicle is a rotary wing aircraft. Indeed, in this case, the power plant of the aircraft must provide sufficient power so that the lift of the main rotor is at least opposed to the estimated mass M of the aircraft. This estimated mass M of the vehicle can be determined before the aircraft takes off. During the flight, the instantaneous mass of the aircraft decreases following the particular fuel consumption and is then less than this estimated mass M. As a result, this estimated mass M of the vehicle can be used to determine each characteristic power for the realization a predetermined operation X of the vehicle. However, this characteristic power is then overvalued with respect to the mechanical power actually necessary for the predetermined operation of the aircraft X, the estimated mass M of the aircraft being greater than its instantaneous mass.
[0014] This overvalued characteristic power goes in the direction of safety, the mechanical power actually required being less than this overvalued characteristic power. On the other hand, for certain maneuvers requiring a large mechanical power, this overvalued characteristic power may be greater than the maximum available power, whereas the mechanical power actually necessary for the predetermined operation of the aircraft is less than this maximum available power. for each engine. In fact, this maneuver is not performed by the pilot who mistakenly thinks that the power plant can not provide a sufficient total power.
[0015] Preferably, an estimated instantaneous ground Minst of the vehicle is defined and used in order to more precisely determine each characteristic power for carrying out a predetermined operation X of the vehicle. This estimated instant mass Minst of the aircraft is for example determined by calculating the fuel consumption during the flight, the mass of fuel consumed then being deducted from the estimated mass M of the aircraft before take-off. In addition, from the maximum total power available at the power plant and the estimated instantaneous ground Minst of the vehicle, a maximum mass transportable by the vehicle can be determined. Indeed, it is possible by using the vehicle performance curves and taking into account the available power margin to determine the total mass of the vehicle for which the total power supplied by the power plant is equal to the maximum total power. available. It is then possible to deduce the maximum transportable mass which is the difference between this total mass and the estimated instantaneous mass Minst of the vehicle. Furthermore, by determining at least a second value 13_, X 25 and / or a characteristic power corresponding to a predetermined operation X, the method for determining parameters characteristic of the operation of a vehicle makes it possible to assist the driver of the vehicle in the vehicle. anticipation and in the possible realization of this predetermined maneuver X. The risks associated with this predetermined maneuver X are then reduced significantly, especially for maneuvers for which the accident rate is high, the pilot knowing, for example, whether he has a margin of sufficient power to perform this maneuver. X predetermined. Likewise, the realization of this predetermined operation X can also be optimized, for example in terms of fuel consumption. This is particularly the case where the vehicle is a rotary wing aircraft comprising a main rotor provided with main blades, an anti-torque rotor, the mechanical power transmission means being a main gearbox driving the main rotor and the anti-torque rotor. A predetermined maneuver X is, for example, for a rotary wing aircraft, a slow descent flight for a landing during which many accidents can occur, in particular during transitions between an area where the aircraft is subject to a ground effect and an area where the aircraft is not subject to that ground effect. Indeed, near the ground, the main rotor blast is returned by the ground on the fuselage and the main blades of the main rotor modifying the behavior of the aircraft. It is said that the aircraft is in the ground effect, defined by the acronym DES or IGE for the expression in English "In Ground Effect". This DES zone covers a height from the ground to about four times the diameter of the main rotor of the aircraft.
[0016] Above this DES area, that is to say for a height above the ground greater than about four times the rotor diameter of the aircraft, the aircraft no longer undergoes ground effect. It is said that the aircraft is out of ground effect, defined by the acronyms HES or OGE for the expression in English "Out 30 Ground Effect".
[0017] On the other hand, when the aircraft is hovering, that is to say that its horizontal speed Vh and its vertical speed Vz are substantially zero, the functional characteristics of the aircraft are defined in particular by a series of firsts. 5 performance curves of the aircraft according to a first formula (NRNR0) 3 = k. i [ftL1 NR NR0) 21. In this case, W is the flight power of the aircraft, a is a reduction coefficient, k is a coefficient of influence of the ground effect on the behavior of the aircraft, M is the estimated mass of the aircraft, the aircraft, NRo is a rotation speed setpoint of the main rotor, NR is the main rotational speed of the main rotor and fi is a first function represented by a series of first aircraft performance curves having NR) for abscissa a first value Al = (NR0J2 and for ordinate (a second value A2 = 17 NR0 3. The estimated mass M is preferably UIR), preferably the estimated instant mass Muist of the aircraft, This series of first performance curves of the aircraft is specific to each aircraft and provided with the flight manual of the aircraft Po The reduction coefficient a is equal to the ratio - Po and To To 20 being respectively the atmospheric pressure and the temperature around the aircraft Atmospheric pressure Po is expressed in millibar (mb) and the temperature To is expressed in kelvin (K). The coefficient of influence k is equal to unity when the aircraft is in an HES zone. This influence coefficient k varies according to the height of the aircraft from the ground relative to the ground from a minimum value to a maximum value when the aircraft is in a DES zone. The maximum value is generally equal to 1.1 for all aircraft while the minimum value, which corresponds to a ground-level aircraft 3033316 24, varies according to the aircraft and can be between 0.6 and 0.9. The coefficient of influence k is for example between 0.6 and 1.1 for a given aircraft. Thus, in order to determine a first characteristic power Wk of the aircraft for carrying out a predetermined maneuver X in the case of a hover regardless of the height Hz of the aircraft relative to the ground, the coefficient of reduction 0- such that u = To 'V - the first value Ai is calculated such that Al (NRO 1 o -NR 10 - the height Hz of the aircraft is measured with respect to the ground, for example using an altimeter radio present in the aircraft, the influence coefficient k corresponding to the height Hz of the aircraft with respect to the ground is determined, it is determined, by means of a first curve of performance of the aircraft, according to the first function fi, corresponding to the flight conditions of the aircraft and according to the first value Al, the second value Az such that A2 = fit-Ma (NRR) 211 'is calculated from the second value Az the first characteristic power Wk of the aircraft for the realization of a man X predetermined work in the case of a hover out of ground effect or in the ground effect according to the coefficient of influence k, such that Wk = k. A2. CT. (-NR) 3. NRo The first performance curves according to the first function fi are dependent on the aircraft and make it possible to characterize this aircraft. The first performance curve of the aircraft, according to the first function fi used to determine the second value A2, is defined according to the flight conditions of the aircraft and in particular the current ambient conditions such as atmospheric conditions. These first performance curves make it possible to determine the flight power W of the aircraft, which is the sum of the power required at the main rotor and the power required at the level of the anti-torque rotor to hover the aircraft. The total power of the power plant of the aircraft to achieve this hover is then equal to the sum of this flight power W and the accessory power of the power plant. Then, the first characteristic power Wk and the maximum total power available for the flight of the aircraft as well as the total instantaneous power consumed for the flight of the aircraft can be displayed simultaneously on the graphical representation. The maximum total power available for the flight of the aircraft and the instantaneous total power consumed for the flight 15 of the aircraft are respectively equal to the total instantaneous power Winst supplied by the power plant and its maximum total available power deduced from the accessory power. The total instantaneous power WInst supplied by the power plant and its maximum total available power are respectively equal to the sum of the instantaneous powers Wm, nst supplied by each engine of the power plant and to the sum of their maximum powers available when the power plant has several motors. The pilot can then check the difference between the first characteristic power Wk of the aircraft for carrying out the predetermined maneuver X and the maximum total power available for the flight. Similarly, it is possible to display on the graphical representation a second value PLX corresponding to the first characteristic power Wk and a third value P, inst 3033316 corresponding to the instantaneous total power consumed for the flight of the aircraft as well as first value P, / im corresponding to the maximum total power available for the flight of the aircraft. The pilot can then check the difference between the second value P, X and the first value PLIim. In addition, it is then possible to hover the aircraft at this height Hz automatically by applying this first characteristic power Wk to the power plant by adding the accessory power. For this purpose, it is possible to control the engine torque Cmk of each engine of the powerplant for the coefficient of influence k corresponding to this height Hz or the collective pitch of the main rotor blades of the main rotor of the aircraft so that the The power plant provides a total power equal to the sum of the first characteristic power Wk and the accessory power. For example, it is possible to determine a motor torque Cmk that each engine of the power plant corresponding to this total power must supply, then this motor torque Cmk is applied to each motor of the power plant, thus automatically achieving a hovering at the height Hz. It is also possible to determine, the control means of the aircraft being means for controlling the variation of the collective pitch and the cyclic pitch of the main rotor blades of the main rotor, a collective pitch value of these main blades corresponding to this first characteristic power Wk, then this collective pitch is applied to the main rotor blades of the main rotor thus automatically achieving this hovering flight at the height Hz. Moreover, a second characteristic power Wk, of a hovering flight of the aircraft can be determined. in an HES area. For this, the coefficient of influence k is defined equal to 1, then from the second value A2 a second characteristic power Wk = 1 of the aircraft is calculated for carrying out a predetermined maneuver X of the aircraft in the case of a ground hover off ground effect of the aircraft such that 5 Wk = i = A2. (-Ni NRo) - Similarly, in order to determine a third characteristic power Wkrnin, of a hovering flight of the aircraft at ground level, the coefficient of influence k is defined equal to a minimum value km, nor corresponding to a stationary flight of the aircraft at ground level, then from the second value A2 the third characteristic power Wkmin of the aircraft is calculated for carrying out a predetermined maneuver X of the aircraft in the case of a ground hover with ground effect of the aircraft at ground level NR such that W k kmini = min. A2. As a result, the second characteristic power Wk = 1 and the third Winnini characteristic power can be displayed simultaneously on the graphical representation as well as the maximum total power available for the flight. to verify, before engaging the predetermined maneuver X 20, that the first and second characteristic powers Wk = /, Wkmini are well below the maximum total power available for the flight thus guaranteeing the realization of a slow descent of the aircraft by in view of a landing and in particular the transition from an HES zone to a DES zone, without the risk that the total power supplied by the power plant is insufficient This slow descent must be carried out without the appearance of a vortex effect to correspond to the aerodynamic modeling of the flight of the aircraft used to define the first 3033316 28 performance curves of the aircraft. vortex occurs mainly during a rapid descent flight of the aircraft when the main rotor of the aircraft is in a vortex flow that has itself generated then causing a sudden drop in its lift. Similarly, two second PLX values corresponding respectively to the second characteristic power Wk _, / and to the third characteristic power Wkmini and to a first value PL / im corresponding to the maximum total power can be displayed simultaneously on the graphical representation. available for the flight. Thus, the pilot can check before engaging the predetermined operation X that the two second PLX values are well below the first value Pi_lim.
[0018] In addition, the method for determining characteristic parameters of the operation of an aircraft according to the invention makes it possible to automatically achieve a slow descent of the aircraft from a height Hz of the aircraft relative to the ground to the ground, then a landing. For this purpose, a total torque to be provided by the power plant corresponding to a power of descent between the third characteristic power Wkmini and the first characteristic power Wk is determined. The total torque is applied to the power plant. thus achieving automatically a descent of the aircraft, - the height Hz of the aircraft is measured with respect to the ground, - the coefficient of influence k is adjusted during the descent of the aircraft as a function of the reduction of the the height Hz and thereafter the first characteristic power Wk, and consequently the total torque that the motor installation 3033316 must provide to achieve a descent flight of the aircraft to the ground level, then a landing. It is also possible to determine a collective pitch of these main blades corresponding to the total torque to be supplied by the powerplant, then this collective pitch is applied to the main rotor blades of the main rotor thus automatically achieving a slow descent of the aircraft to its landing. In addition, in order to avoid an excessive descent of the aircraft and a hard landing, or even a crash on the ground, it is possible to apply a factor of safety increasing the third characteristic power Wkmini. The value of this safety coefficient is preferably proportional to the difference between the third characteristic power Wkmini and the first characteristic power Wk so that the descent power converges towards the third characteristic power Wkmini during the descent of the aircraft. A predetermined maneuver X may also be, for a rotary wing aircraft, a level cruise flight, corresponding to a vertical speed Vz of the aircraft substantially zero and a horizontal speed Vh nonzero. During this cruising flight, the functional characteristics of the aircraft are defined by a series of second performance curves according to a second formula W = f2 (Vh). In this case, W is the flight power of the aircraft, Vh is the horizontal speed of the aircraft and f2 is a second function represented by a series of second aircraft performance curves.
[0019] This series of second f2 performance curves of the aircraft is specific to each aircraft and provided with the aircraft flight manual. From this series of second curves f2, a horizontal line tangent to the second curve corresponding to the flight conditions of the aircraft makes it possible to determine a fourth characteristic power Wend and a first characteristic horizontal speed Vend to obtain a fuel consumption of minimum fuel and, consequently, a maximum flight time of the aircraft. From this series of second curves f2, a line tangent to the second curve corresponding to the flight conditions of the aircraft and passing through the origin point of the reference of the second curve makes it possible to determine a fifth power characteristic Wrange 15 and a second Vrange characteristic horizontal speed allowing a maximum range to be obtained by the aircraft. In the case where the aircraft is subjected to a longitudinal wind, the speed of this longitudinal wind can be taken into account to determine this fifth Wrange range characteristic power and this second characteristic Vrange speed. The line tangent to the second curve does not then pass through the origin point of the reference of the second curve but by a point offset on the abscissa axis of the origin point of the mark according to the value of this longitudinal wind speed. It is then possible to display simultaneously on the graphical representation the fourth characteristic power Wend end and the fifth characteristic power Wrange range as well as the instantaneous total power consumed for the flight of the aircraft and the maximum total power available for the flight. The pilot can then adapt the instantaneous power according to these objectives. Similarly, two second PAX values corresponding respectively to the fourth characteristic power Wend and to the fifth power characteristic Wrange and to a first value Pilim corresponding to the maximum total power available for the flight and to the first value Pilim can be displayed on the graphical representation. a third value P _inst corresponding to the instantaneous total power consumed for the flight of the aircraft. The pilot can then check the difference between each second PLX value and the first Pd / m value. A predetermined maneuver X may also be, for a rotary wing aircraft, an ascending flight, corresponding then to a vertical speed Vz of the non-zero aircraft. This upward flight is carried out after a take-off phase for example and in an area where the aircraft is not subject to the ground effect. During this ascending flight, the functional characteristics of the aircraft are defined by a series of third performance curves according to a third formula (17) Vy = f3 (e). In this case, W is the flight power of the aircraft, a is the reduction coefficient, M is the estimated mass of the aircraft, Vy is an optimum vertical climb speed of the aircraft, f3 is a third function. represented by a series of third performance curves of the aircraft. The ratio (11) Vy is obtained for a vertical speed Vz of the aircraft equal to the optimum climb speed Vy of the aircraft, so that the flight power W then corresponds to the optimum climb speed Vy.
[0020] This series of third performance curves f3 of the aircraft is specific to each aircraft and provided with the aircraft flight manual. Thus, in order to determine a sixth characteristic power Wvy of an ascending flight of the aircraft in a zone HES, a third value A3 such that A3 = Cr is calculated, it is determined by a third performance curve of the the aircraft, according to the third function f3, corresponding to the flight conditions of the aircraft and, as a function of the third value A3, a fourth value A4, such that A4 = 0-) Vy - is calculated from the fourth value A4 the sixth characteristic power Wvy corresponding to the optimum rise speed Vy, such that Wvy = A4.6. It is then possible to display on the graphical representation 15 the sixth characteristic power Wvy and the maximum total power available for the flight of the aircraft as well as its instantaneous total power consumed for the flight. The pilot can then verify that this sixth characteristic power Wvy is less than the maximum total power available for the flight and that he can realize this ascension flight at the optimum speed Vy. Similarly, a second PLX value corresponding to the sixth characteristic power Wvy and a first value Pjim corresponding to the maximum total power available for the flight can be displayed on the graphical representation. Moreover, the display of the instantaneous, maximum available and characteristic total powers of the aircraft as well as the PLinst, P, Jim, P, X values corresponding to these total powers can be replaced by the display of the instantaneous, maximum powers. available and characteristics for each motor of the power plant as well as values PEinst, PENm, PLX corresponding to these powers at these engines. In fact, the distribution of the total power of the power plant on each of its engines is known. We then have a graphical representation for each engine of the power plant. The present invention also relates to a device for determining characteristic parameters of a vehicle, the vehicle comprising a power plant provided with at least one motor and a mechanical power transmission means and control means, the device comprising a plurality of sensors, at least one calculating means, at least one memory and at least one visualization means. The sensors provide measurements of various information relating to the vehicle environment and / or the condition and operation of the vehicle and its equipment and / or the position and movements of the vehicle relative to its environment, each computing means receiving the sensor measurements and processing this information. The vehicle may also include an altimeter radio to determine a height Hz of the vehicle relative to the ground. The memory stores vehicle performance curves and calculation instructions, the calculation means using the calculation instructions to implement the method of determining characteristic parameters of the vehicle described above. The invention and its advantages will appear in more detail with reference to the following description with examples given by way of illustration with reference to the appended figures which show: FIG. 1, a rotary wing aircraft provided with a device for determining parameters characteristic of the operation of the aircraft, - FIGS. 2 and 3, two block diagrams of a method for determining parameters characteristic of the operation of the aircraft, FIGS. 4 to 6, representations graphical type of the 10 values of each parameter Pi, and - Figures 8 to 10, the aircraft performance curves. The elements present in several separate figures are assigned a single reference. FIG. 1 shows a rotary wing aircraft 10 comprising a main rotor 11 provided with main blades 12, a rear rotor 13 having in particular an anti-torque function, this rear rotor 13 being provided with secondary blades 14. The aircraft 10 comprises also a power plant 20 provided with two turbine engines 21,22, a main gearbox 20 of power 23 driving in rotation the main rotor 11 and the rear rotor 13 and control means 24,25,26 composed of a lever of collective step 24 and a cyclic stick 25 which are respectively means for controlling the variation of the collective pitch and the cyclic pitch of the main blades 12 as well as a spreader 26 which is a means for controlling the variation of the pitch collective collective of the secondary blades 14. These control means 24,25,26 allow the pilot to control movements of the aircraft 10. The two turbine engines 21,22 are identical and comprise a generator of gas and a free turbine. The two turbine engines 21, 22 provide substantially the same power to the main gearbox 2333316 when both of these turboshaft engines 21, 22 are operating properly. However, the two turbine engines 21,22 may be different and then provide different powers to the main power transmission box 23. A particular distribution of these powers is then defined between these turboshaft engines 21,22. Finally, the aircraft 10 comprises a device 1 for determining parameters characteristic of the operation of the aircraft 10 10 comprising a plurality of sensors 5-8, a calculation means 2, a memory 3 and at least one visualization means 4. sensors 5-8 make it possible to measure various information relating to the environment of the aircraft 10, to the state and to the operation of the aircraft 10 and the powerplant 20, in the state of the control means 24 , 25,26, as well as the position and movements of the aircraft. The sensors 5, 6 are for example a means for measuring the atmospheric pressure and a means for measuring the temperature outside the aircraft 10. The sensors 7, 8 are for example a means for determining the attitude and the heading of the aircraft 10 such as a device AHRS for the English expression "Attitude and Heading Reference System" and a means of control of the power plant 20. The device 1 also comprises a radio altimeter 9 determining the height Hz of the aircraft 10 from the ground. The device 1 can implement a method for determining parameters characteristic of the operation of the aircraft 10 whose block diagram is shown in FIG. 2. The memory 3 stores aircraft performance curves 3033316 36 10 and instructions calculation and calculation means 2 uses these calculation instructions to implement this method. The process according to Figure 2 comprises five steps. During a first step 110, different information relating to the environment of the aircraft 10, and / or the state and operation of the aircraft 10 and the power plant 20 and / or in the state of the control means 24,25,26, and / or the position and movements of the aircraft 10 relative to its environment. These different pieces of information are notably obtained via the sensors 5-8. During a second step 131, at least a first value Pi_lim corresponding to a limit value that a parameter Pi must not be exceeded, Pi being a parameter relative to the state or to the operation of the aircraft 10, is determined. of the power plant 20, or the state of the control means 24,25,26, or the position or movements of the aircraft 10 relative to its environment. Each parameter Pi can be chosen from the following list or can be a logical or arithmetic combination of at least two elements of this list: a total torque supplied by the power plant 20 or a torque supplied by a turbine engine 21,22 or a rotor torque at a driving mast in rotation of the main rotor 11 or a torque at said rear rotor 13, 25 - a total power supplied by the power plant 20 or a power provided by one of the turbine engines 21,22 - a speed of the aircraft 10 relative to the air or a speed of the aircraft 10 relative to the ground, - a height of the aircraft 10 relative to a reference such as 30 that the soil, a temperature inside a turbine engine 21,22 such as the temperature of the gas at the inlet of the free turbine, a rotational speed of a member of a turbine engine 21 , 22 such as the gas generator, 5 - a position of a e of the aircraft 10 or a pitch value of the main rotor blades 12 of the main rotor 11 or of the secondary blades 14 of the rear rotor 13, - a position of the control means 24 of the collective pitch of the main rotor blades 12 of the main rotor 11, 10 - a position of the control means 25 of the cyclic pitch of the main blades 12 of the main rotor 11, - a position of the collective control means 26 of the secondary blades 14 of the rear rotor 13. During a third step 132, at least a second value PLX is determined at which the parameter Pi must be equal so that the aircraft 10 can perform a predetermined operation X. During a fourth optional step 133, a third value PLinst equal to an instantaneous value of each parameter Pi is determined. During a fifth step 140, each first value P is displayed simultaneously on the same graphical representation. , _lim, each second value PLX and each third value P, inst if necessary to clearly highlight the relative position of a first value P, hm with respect to each second value PLX and each third value P, _inst if necessary for each parameter P ,. Thus, the pilot easily and quickly displays on the one hand a first margin which he has between the third value P instantaneous _inst 30 and the first value P _limlim limit of each parameter 3033316 38 P, and secondly a second margin it possesses between the second value PLX and the first value Pi_lim limit of each parameter P. Consequently, the pilot can adapt the use of each parameter P, according to the first margin and anticipate the achievement of the maneuver X according to the second margin. The graphic type representation on the display means 4 may take several forms such as a dial 30 of substantially circular shape shown in Figure 4 or a vertical scale 40 shown in Figures 5 and 6.
[0021] In FIG. 4, two needles 31, 31 'are rotatable on the dial 30 and respectively indicate the third value P, inst relating to each turbine engine 21, 22. Marks 32,33,34,35 positioned outside the dial 30 indicate second values PLX respectively corresponding to predetermined maneuvers X. The position of each marker 32,33,34,35 may change according to the flight conditions of the aircraft 10 and according to the environmental conditions surrounding the aircraft 10. Each second PLX value and the third value PI inst are represented on the dial As a percentage of the first value Pi Hm, the first value Pi / im corresponding to the mark 39 representing a value of 100% on the dial 30. However, each first value Pi lim, each second value PLX and each third value P Inst can be represented on each dial 30 as a percentage of a reference value of this parameter P 1. The device 1 for determining parameters characteristic of the operation of the aircraft 10 may also comprise two dials 30 provided with a single needle 31, a dial 30 being dedicated to a single turbine engine 21,22.
[0022] In addition, the graduations of the dial 30 go beyond 100%, since each first value Pi lim generally corresponds to normal and continuous operation of the aircraft and the turboshaft engines in particular, but may be temporarily exceeded. by a third value Pi inst, for example following a failure of a turbine engine. In addition, second values PLX may exceed each first value signifying that the predetermined operation X requires exceeding this first value Pi lim for the parameter Pi.
[0023] In FIGS. 5 and 6, a first vertical scale 40 has two subscales 55,56, a first subscale 55 indicating the collective pitch value of the main rotor blades 12 of the main rotor 11 and a second subscalig 56 indicating the longitudinal advancement speed of the aircraft 10. In this case, there is no indication 15 specific to each turbine engine 21,22 of the values relating to the parameter P 1. The first value P, lim and the second values PLX are represented by horizontal marks 49,42,43,44,45 as well as the third value Pi inst represented by a horizontal mark 41. The first value Pi lim, every second value PLX and the third value pi inst are represented on the subscale 55 according to collective pitch values of the main blades 12. The collective pitch of the main blades 12 of the main rotor 11 of the aircraft 10 being an essential element for the pilot of the rotary wing aircraft 10, the pilot thus clearly and quickly displays the collective pitch values of the main blades 12 corresponding respectively to the first value P, hm, to each second value PLX and to the third value Pi_inst as well as the margins between these values of collective pitch.
[0024] In FIG. 5, a second scale 60 indicates the height Hz of the aircraft 10 relative to the ground. The current value of this pitch Hz is determined by means of the radio altimeter 9 and displayed by the mark 61. The display means 4 is for example an instrument of first limitation of the aircraft 10.
[0025] Moreover, the parameters Pi represented in FIGS. 4 to 6 are frequently determined in correlation with the flight power of the aircraft 10 or with the power of flight at each turbine engine 21, 22. The flight power of the aircraft 10 is the sum of the power required at the main rotor 11 and the power required at the rear rotor 13 in order to achieve the flight of the aircraft 10. The total power supplied by the power plant 20 is the sum of this flight power and an accessory power required to power the equipment of the aircraft 10. In addition, the total power supplied by the power plant 20 is substantially equal. to the sum of the power supplied respectively by each turbine engine 21,22 of the power plant 20, this total power being supplied substantially equal by each turbine engine 21,22 when these two identical 21,22 turbine engines operate properly. As a result, the flight power of the aircraft 10 is equal to the total power provided by the power plant 20 deduced from the accessory power. Similarly, the power of flight at each turbine engine 21,22 is equal to half of this flight power of the aircraft 10.
[0026] The method for determining characteristic parameters of the operation can then comprise intermediate substeps between the first and the second steps 110, 131. A synoptic diagram of such a method, then comprising nine steps and sub-steps, is shown in FIG. 3. During a first substep 121, an available power margin of each turbine engine 21, 22 is determined. with respect to a guaranteed minimum power, the available power margin characterizing a state of aging of each turbine engine 21,22. During a second substep 122, an instantaneous power Wminst supplied by each turbine engine 21, 22 is determined. During a third substep 123, a maximum available power is determined at each turbine engine 21, 22 taking into account the available power margin. During a fourth substep 124, at least one characteristic power of each turbine engine 21,22 corresponding to a predetermined operation X of the aircraft 10 is determined.
[0027] Then, during the second step 131, two first values P. lim are determined, respectively corresponding to the maximum available power of each turbine engine 21,22. During the third step 132, two second values PLX corresponding to a characteristic power for each turbine engine 21, 22 are determined to perform the predetermined operation X of the aircraft 10. During the fourth step 133, two thirds are determined. P values, _inst respectively corresponding to the instantaneous power Wminst 3033316 provided by each turbine engine 21,22. Finally, during the fifth step 140, the first value & Hm, each second value PL_X and the third value P inst for each turbine engine 21, 22 are simultaneously displayed on the same graphical representation. The parameter Pi is in this case relative to the operation of a turbine engine, such as the engine torque Cm supplied by a turbine engine 21,22, the internal temperature T4 of the gases at the inlet of the free turbine of a turbine engine 21, 22 or the speed of rotation Ng of the gas generator of a turbine engine 21,22. In addition, in order to limit the number of parameters Pi that the pilot must monitor, only the parameter P, the most critical, is displayed on the visualization means 4. This most critical parameter P1 for the piloting of the aircraft 10 is the one for which the difference between the third value P, _inst and the first value is the lowest. Thus, the pilot can clearly and quickly visualize the margin he currently has on the current flight conditions and for the predetermined maneuver X that he plans to achieve. The available power margin of each turbine engine 21,22 of the power plant 20 can be determined during a "motor health" control, which must be carried out regularly. The instantaneous power Wminst supplied by each turbine engine 21,22 can be defined according to the formula Wminst = Cm.Ng, C being the engine torque supplied by the turbine engine 21,22 and Ng being the speed of rotation of the gas generator of this turbine engine 21,22.
[0028] The instantaneous power Wminst supplied by each turbine engine 21, 22 can also be determined from the flight power of the aircraft 10. The power at the main rotor 11 is known according to the formula Winst = CR.NR, NR being the instantaneous speed of rotation of the main rotor 11 and CR being a rotor torque at a main rotor drive mast 11. Likewise, the power at the rear rotor 13 as well as the accessory power can also be determined in a known manner. Finally, this instantaneous total power WInst 10 supplied by the power plant is distributed equally by each turbine engine 21,22 when these two turbine engines 21,22 work properly. The characteristic power of each turbine engine 21,22 is then determined according to the predetermined maneuvering X envisaged to achieve the pilot of the aircraft 10. This predetermined maneuver X may be for example a hover, a landing or a flight of cruise. When the aircraft 10 is hovering, the functional characteristics of the aircraft 10 are characterized by a series of first performance curves shown in FIG. 7. These first performance curves are N (defined by the first formula 1 ± (-NR0NR) 21, where W is 0- NR120 V = k .f the flight power of the aircraft 10, a being a reduction coefficient, k being a coefficient of influence of the ground on the behavior of the aircraft; the aircraft 10 as a function of the height Hz of the aircraft 10 with respect to the ground, M being the estimated mass of the aircraft 10, NR0 being a rotational target speed of the main rotor 11, NR being the actual rotational speed the main rotor 11 and fi being a first function represented by a series of first 30 performance curves of the aircraft 10. In FIG. 3, the abscissae of these first performance curves correspond to the first value A1 = 1- '1.M / 2 and their ordinates at the second value A2 = W - 3 to NR NR The influence coefficient k is specific to each aircraft 5 and defined by a ground influence curve shown in FIG. 8 having the abscissa the height Hz of the aircraft 10 with respect to the ground and for ordinate this coefficient of influence k. In the curve shown in FIG. 8, this influence coefficient k is between 0.9 and 1.1.
[0029] In order to determine the characteristic flight power of the aircraft 10 for a hovering flight, the reduction coefficient a is first calculated such that a = (M, then the first value Ai such that Al = Al (NR0) 2. The mass M can be an estimated mass NR of the aircraft 10 determined before its take-off or preferably an estimated instantaneous mass Minst of the aircraft 10. Then, the height Hz of the aircraft 10 is measured with respect to on the ground, for example by means of the radio altimeter 9 of the aircraft 10, and the influence coefficient k corresponding to the height Hz is determined using the ground influence curve of FIG. 8.
[0030] Then, using a first performance curve of the aircraft 10 according to the first function fi corresponding to the flight conditions of the aircraft and according to the first value A1, a second value A2 such that A2 = o is determined. - (V! NRO) 2] as shown in Figure 7.
[0031] Finally, from the second value A2, a first characteristic power Wk of the aircraft 10 is calculated in the case of hovering out of ground effect or in the ground effect in accordance with the height Hz of the aircraft. aircraft 10 with respect to the ground and according to the influence coefficient k, such that Wk = k.A2.6. (NR) 3. The characteristic power of each turbine engine 21,22 in the case of a hover out of ground effect or in the ground effect is then equal to half of this first power characteristic Wk of the power plant. In order to determine a second characteristic power Wk = 1 of the aircraft 10 during a hovering out of ground effect, the influence coefficient k is defined equal to 1 and the second value A2 is calculated from the second value A2. characteristic power Wk = 1 such that Wk-, i = NR 4Rol In order to determine a third characteristic power Wkmini of the aircraft 10 during a stationary hover at ground level, the minimum influence coefficient k is defined, such as 15 kmini = 0.9 according to the soil influence curve of FIG. 8 and the third characteristic power Wkmini is calculated from the second value A2 such that Wkmini = (0.9) .A2.CT. (4NR) 3 1R J The second values PLX corresponding respectively to the second characteristic power Wk, / and the third characteristic power Wkmini can then be determined and then displayed on the display means 4 via the marks 34 and 35 of the dial 30. the fig 4 and markers 44 and 45 of the scale 55 in FIG. 5. In addition, on the scale 60 of FIG. 5, a mark 61 indicates the height Hzk, the lowest corresponds to the coefficient k equal to 1. This lowest Hzk height is defined by the curve shown in FIG. 8, providing the variation of the coefficient k according to the height Hz. Thus, the pilot of the aircraft 10 displays the height Hzk that he will reach if he modifies progressively. the current collective step value 3033316 46 up to the mark 44 corresponding to the second value PLX which corresponds to the second characteristic power Wk.i. However, this change in the value of the collective pitch must be progressive in order to limit the dynamic and inertial effects of this descent of the aircraft 10. The aircraft 10 can also automatically make a slow descent to a landing. following the calculation of the first characteristic power Wk and the third characteristic power Wkmini.
[0032] For this purpose, the torque to be provided by each turbine engine 21,22 corresponding to a power of descent between the third characteristic power Wkmini and the first characteristic power Wk is determined, taking into account the accessory power and possibly a factor of safety. , 15 then this torque is applied to each turbine engine 21,22. The coefficient of influence k is then adjusted during the descent of the aircraft 10 as a function of the reduction of the height Hz and subsequently the first characteristic power Wk, and consequently the torque that each turbine engine must provide 21,22 until landing of the aircraft 10. It is also possible to determine the collective pitch of the main rotor blades 12 of the main rotor 11 corresponding to this power of descent between the third characteristic power Wkmini and the first characteristic power Wk.
[0033] The collective pitch is applied to the main blades 12 of the main rotor 11, for example via an automatic pilot of the aircraft 10, thus automatically achieving a slow descent towards a landing of the aircraft 10. The value of this collective pitch is then adjusted following the descent of the aircraft 10 and the variation 3033316 47 of the influence coefficient k is adjusted according to the reduction of the height Hz. In addition, from the maximum total power available at the facility 20 and thanks to the first 5 performance curves according to the first function fi shown in Figure 7, it is possible to determine the maximum mass transportable by the aircraft 10. Indeed, the maximum transportable mass is the difference between the mass of the 10 for which the total power supplied by the power plant 20 is equal to this maximum total available power, taking into account the power margin available. and the estimated instantaneous mass Minst of the aircraft 10. This maximum transportable mass can then be indicated to the pilot of the aircraft 10 on the display means 4 by means of a visual indication 38,48.
[0034] When the aircraft 10 is in cruising flight, the functional characteristics of the aircraft 10 are characterized by a series of second performance curves shown in FIG. 9. These second performance curves are defined by a second formula W = f2 (Vh), Vh being the horizontal speed of the aircraft 10 and f2 being a second function represented by a series of second performance curves of the aircraft 10. In FIG. 9, the abscissae of these first performance curves are constituted by the horizontal velocity Vh of the aircraft 10 and their ordinates by the flight power W of the aircraft 10. These second performance curves pass through a minimum power consumed by the power plant 20 which is a fourth characteristic power. Wend to obtain a minimum consumption for the power plant 20 and a maximum flight time of the aircraft 10 under the conditions of this 303331 6 48 cruise flight. A first characteristic point Ci of this fourth characteristic power Wend is determined in FIG. 13 by the point of tangency between a horizontal straight line and the second performance curve corresponding to the flight conditions of the aircraft 10. The first characteristic horizontal speed Vend corresponding to this fourth characteristic power Wend is also defined by this first characteristic point Ci. From these second performance curves, a fifth characteristic power Wrange range and a second characteristic horizontal speed Vrange corresponding to the best compromise between the speed of progress of the aircraft and its power consumed to obtain a maximum range by the aircraft 10 under the conditions of this cruising flight can be determined. A second characteristic point C2 of this fifth Wrange characteristic power is determined in FIG. 13 by the intersection between the second performance curve which corresponds to the flight conditions of the aircraft 10 and a line tangent to this second curve passing through the origin point of the reference of this FIG. 9. The second characteristic horizontal speed Vrange corresponding to this fifth characteristic power Wrange range is also defined by this second characteristic point C2. In the case of a longitudinal wind experienced by the aircraft 10, this straight line tangent to the second curve does not pass through the origin point of the reference point of FIG. 9, but by a point shifted on the abscissa axis from the reference point of the reference point. according to the value of the longitudinal wind speed. Thus, in the event of a headwind suffered by the aircraft 10, the fifth characteristic power Wrange range and the second characteristic horizontal speed Vrange are defined by the third characteristic point C3 of FIG. 9, the face wind speed Vvf being positive. Similarly, in the event of a tailwind 3033316 49 experienced by the aircraft 10, the fifth power characteristic Wrange range and the second characteristic horizontal speed Vran, ge are defined by the fourth characteristic point C4 of Figure 9, the speed of the tailwind Vva being negative.
[0035] The second PAX values corresponding respectively to the fourth characteristic power Wend and to the fifth characteristic power Wrange respectively can then be determined and then displayed on the display means 4 via the marks 32 and 33 in FIG. 4 and reference marks. 42 and 43 in FIG. 6. As indicated in FIG. 6, it is also possible to display on the second sub-scale 56 indicating the longitudinal advancement speed of the aircraft 10, the first characteristic horizontal speed Vend and the second speed horizontal characteristic Vrange.
[0036] When the aircraft 10 is in ascending flight, functional characteristics of the aircraft 10 are characterized by a series of third performance curves shown in FIG. 10. These third performance curves are defined by a third formula (i4-- T-) = f3 () W being the Vy 20 flight power of the aircraft 10, Vy being the optimum vertical climb speed of the aircraft 10 and f3 being a third function represented by a series of third performance curves of the aircraft 10; the aircraft 10. The ratio (11) is thus obtained Vy for a vertical speed Vz of the aircraft equal to the optimum climb speed Vy and makes it possible to determine a sixth characteristic power Wvy corresponding to an ascending flight of the aircraft 10 at the optimum rate of climb Vy. A third value A3 is calculated such that A3 = CL10.) Knowing the mass M of the aircraft 10 and the coefficient of reduction 3033316 a. A third performance curve of the aircraft 10 corresponding to the flight conditions of the aircraft according to the third function f3 is then determined and, according to the third value A3, a fourth value A4, such that A4 = f3 C).
[0037] Then, from the fourth value A4, the sixth characteristic power Wvy corresponding to the optimum rise speed Vy, such as Wv3, is calculated. Naturally, the present invention is subject to numerous variations as to its implementation. Although several embodiments have been described, it is well understood that it is not conceivable to exhaustively identify all possible modes. It is of course conceivable to replace a means described by equivalent means without departing from the scope of the present invention. 15

权利要求:
Claims (25)
[0001]
REVENDICATIONS1. A method for determining characteristic parameters of the operation of a vehicle (10), said vehicle (10) comprising a power plant (20) provided with at least one motor (21,22) and a mechanical power transmission means (23), control means (24,25,26) for displacing said vehicle (10), a plurality of sensors (5-8), at least one calculating means (2), at least one memory (3) and at least one visualization means (4), during which various information relating to the environment of said vehicle (10) and / or the state and operation of said vehicle (10) and said power plant are measured. (20) and / or in the state of said control means (24,25,26), and / or at the position and displacements of said vehicle (10) relative to its environment, - at least one first value is determined P, lim corresponding to a limit value that a parameter P1 must not exceed, Pi being a parameter relating to the state or the operation of said vehicle (10), or the state or operation of said power plant (20) or the state of said control means (24,25,26) or the position or displacements of said vehicle ( 10) relative to its environment, - at least one second value P is determined; _X to which said parameter P, must be equal so that said vehicle (10) can perform a predetermined operation X, and - is displayed simultaneously on the same graphical representation each first value P, lim and each second value P, _X so to clearly demonstrate the relative position of a first value PLIim vis-à-vis each second value Pi X for each parameter P ,.
[0002]
2. Method according to claim 1, characterized in that: - at least a third value Pi inst is determined equal to an instantaneous value of each parameter Pi and - on said graphical representation each third value Pi inst is displayed.
[0003]
3. Method according to any one of claims 1 to 2, characterized in that one displays each second value P; _X and each third value Pi instates if necessary on said display means (4) in a percentage vis-à-vis said first value P, // m corresponding to said parameter Pi.
[0004]
4. Method according to any one of claims 1 to 2, characterized in that one displays each first value P, lim, each second value P; _X and each third value PLinst if necessary on said display means (4) in a percentage vis-à-vis a reference value of said parameter
[0005]
5. Method according to any one of claims 1 to 2, characterized in that, said vehicle (10) being a rotary wing aircraft comprising a main rotor (11) provided with main blades (12), a control means (24,25,26) being a means for controlling the variation of the collective pitch of said main blades (12), collective pitch values of the main blades (12) corresponding to each first value P, lim, are displayed at each second value P, Xet at each third value P, inst, if appropriate, respectively at each first value P, lim, at each second value P, _X and at each third value Pi inst, if appropriate, on said display means ( 4).
[0006]
6. Method according to any one of claims 2 to 5, characterized in that one displays each first value PLIim, each second value P; _X and each third value PLinst on an Instrument of First Limitation IPL of said vehicle (10).
[0007]
7. Method according to any one of claims 1 to 6, characterized in that: - an available power margin of each engine (21,22) of said power plant (20) vis-à-vis a guaranteed minimum power, said available power margin characterizing a state of aging of each motor (21,22) of said power plant (20), - a maximum power available is determined at each motor (21,22) of said power plant (20) taking into account said available power margin, determining at least one characteristic power of each motor (21,22) of said power plant (20) corresponding to the achievement of a predetermined operation X 20 said vehicle (10), - at least a first value Pi_lim corresponding to said maximum available power of each engine (21,22) of said power plant (20) is determined, and at least one second value P is determined; _X corresponding to a characteristic power of each motor (21,22) for performing a predetermined operation X said vehicle (10).
[0008]
8. Method according to any one of claims 1 to 6, characterized in that: - one determines an available power margin of each motor (21,22) of said power plant (20) vis-à-vis a guaranteed minimum power, said available power margin 5 characterizing a state of aging of each motor (21,22) of said power plant (20), an instantaneous power Wminst supplied by each motor (21,22) of said power plant (20) determines a maximum power available at each engine (21,22) of said power plant (20) taking into account said available power margin; at least one characteristic power is determined of each motor (21,22) of said power plant (20) corresponding to the achievement of a predetermined operation X of said vehicle (10), - at least a first value P i _lim corresponding to said power m is determined; maximum available from each motor (21,22) of said power plant (20), - at least one second value P is determined; X corresponding to a characteristic power of each motor (21,22) for performing a predetermined operation X of said vehicle (10) and determining at least a third value Pi_inst corresponding to said instantaneous power Wminst supplied by each motor (21,22).
[0009]
9. Method according to any one of claims 7 to 8, characterized in that it determines an estimated instantaneous mass Mittst said vehicle (10) to determine each characteristic power. 30
[0010]
10. The method of claim 9, characterized in that it is determined from a maximum total power available at said power plant (20) equal to the sum of said maximum power available at each engine (21,22), a maximum mass transportable by said vehicle (10), said maximum transportable mass being the difference between the mass of said vehicle (10) for which a total power supplied by said power plant (20) is equal to said the maximum total power available according to performance curves of said vehicle (10) taking into account said available power margin and said estimated instantaneous mass Minst of said vehicle (10).
[0011]
11.A method according to any one of claims 7 to 9, characterized in that, said vehicle (10) being a rotary wing aircraft comprising a main rotor (11) provided with main blades (12), an anti-torque rotor ( 13) provided with secondary blades (14), said mechanical power transmission means (23) being a main power transmission gearbox rotating said main rotor (11) and said anti-torque rotor (13), when said aircraft (10) ) is in stationary flight, its horizontal speed Vh and its vertical speed Vz being substantially zero, the functional characteristics of said aircraft (10) being characterized in particular by a series of first performance curves according to a first formula NR (NR013 =. 1, W being a flight power of said aircraft J Jf1 NR J 25 (10), u being a reduction coefficient, k being a coefficient of influence of the ground on the behavior of said aircraft (10) as a function of the height Hz of said aircraft (10) relative to said ground, M being said estimated mass of said aircraft (10), NR0 being a rotational target speed of said main rotor (11), NR being the actual rotational speed of said rotor main (11) and fi being a first function represented by a series of first performance curves of said aircraft (10), said flight power W of said aircraft (10) being equal to the sum of said powers supplied by each engine (21). , 22) of said power plant 5 (20) deduced from an accessory power necessary to power equipment of said aircraft (10), - one calculates said coefficient of reduction a such that a = P) To 111) - one calculates a first value Ai such that Al = (cr "NRNR0 2 - said height Hz of said aircraft (10) is measured with respect to said ground, - said coefficient of influence k corresponding to said height Hz is determined, - it is determined, thanks to a first neck rb of performance of said aircraft (10) according to the first function fi corresponding to the flight conditions of said aircraft (10) and according to said first value A1, a second value A2 such that ciRo 2 A2 = f1 F. , a NR) 1 - a first characteristic power Wk of said aircraft (10) is calculated from said second value A2 in the case of a ground hover out of ground effect or in the ground effect NR3 according to said coefficient d the influence k, such that Wk = k.A2.C7. (- NR0).
[0012]
12. Method according to claim 11, characterized in that said coefficient of influence k equal to 1 corresponding to a hovering ground flight of said aircraft (10) is defined, and from said second value A2 a value of second characteristic power Wk = 1 of said aircraft (10) in the case 3033316 57 of a ground hovering ground effect of said aircraft (10) such that (NNRR03 .Wic = i = A2.
[0013]
13. The method as claimed in claim 12, characterized in that, on a scale 60 representing the height Hz of the aircraft 10 with respect to the ground, a mark 61 indicating the lowest height Hzk corresponds to the coefficient k. equal to 1.
[0014]
14. The method according to claim 11, wherein said coefficient of influence k is equal to a minimum value kmini corresponding to a hovering flight of said aircraft (10) at ground level and is calculated from said second value A2 said third characteristic power Wkmini of said aircraft (10) in the case of a hover of said aircraft (10) at ground level such that NR) 3 Wkmini = kmini. A2. See (- NR0
[0015]
15. Method according to claim 11, characterized in that said coefficient of influence k is defined equal to a minimum value kmini corresponding to a hovering flight of said aircraft (10) at ground level, it is calculated from said second value Az a third characteristic power Wkmini of said aircraft (10) in the case of a hover at ground level such that (NR Wkmini = kmini- A2- NR0 3033316 58 a total torque to be provided by said power plant (20 ) corresponding to a power of said power plant (20) between said third characteristic power Wkmini and said first characteristic power Wk, said total torque is applied to said power plant (20), thus automatically achieving a slow descent to a landing, said coefficient of influence k is adjusted during the descent of the aircraft (10) as a function of the reduction of said height Hz and by the after said first characteristic power Wk, and consequently said total torque that must provide said power plant (20) until the landing of said aircraft (10). 15
[0016]
16. The method of claim 11, characterized in that, control means (24,25,26) being means for controlling the variation of the collective pitch and the cyclic pitch of said main blades (12), -on defines said coefficient of influence k equal to a minimum value k Inini corresponding to a hovering flight of said aircraft (10) at ground level, - is calculated from said second value A2 a third characteristic power Wkmini of said aircraft (10) in the case of a hover ground level such as NR) 3 Wkmini = kmini- AZ. a-Nizo - a collective pitch of said main blades (12) corresponding to a power of said power plant (20) between said third characteristic power Wkmini and said first characteristic power Wk is determined, 3033316 59 - said collective pitch is applied to said blades main (12) thus automatically achieving a slow descent to a landing, - one adjusts said coefficient of influence k during the descent of the aircraft (10) as a function of the reduction of said height Hz and subsequently said third characteristic power Wk, and consequently said collective pitch until the landing of said aircraft (10).
[0017]
17. Method according to any one of claims 7 to 9, characterized in that, when said aircraft (10) is in cruising flight, its vertical speed Vz being substantially zero, a series of second performance curves according to one second formula W = f2 (Vh), W being a flight power of said aircraft (10) and Vh being said horizontal speed of said aircraft (10), a horizontal line tangent to said second curve which corresponds to the flight conditions of said aircraft (10); ) makes it possible to determine a fourth characteristic power Wend of said aircraft (10) and a first characteristic horizontal speed Sends said aircraft (10) making it possible to obtain a maximum flight duration of said aircraft (10), said flight power W of said aircraft ( 10) being equal to the sum of said power supplied by each motor (21,22) of said power plant (20) deduced from an accessory power necessary to power equipment of said eronef (10).
[0018]
18. A method according to any one of claims 7 to 9, characterized in that, when said aircraft (10) is in cruising flight, its vertical speed Vz being substantially zero, a series of second performance curves according to one second formula W = f2 (Vh), W being a flight power of said aircraft (10), Vh being said horizontal speed of said aircraft (10) and f2 being a second function represented by a series of seconds 3033316 60 performance curves of the an aircraft (10), a straight line tangent to said second curve which corresponds to the flight conditions of said aircraft (10) and passing through the origin point of the marker of said second curve makes it possible to determine a fifth w-range characteristic power of said aircraft ( 10) and a second characteristic horizontal speed Vrange said aircraft (10) to obtain a maximum range of maximum distance by said aircraft (10), said flight power W of said aircraft (10) sta nt equal to the sum of said power supplied by each motor (21,22) of said power plant (20) deduced from an accessory power necessary to power equipment of said aircraft (10).
[0019]
19. A method according to claim 18, characterized in that in the case of longitudinal wind experienced by said aircraft (10), said line tangent to said second curve does not pass through said origin point of said mark of said second curve, but by an offset point on the abscissa axis of said origin point of said mark according to the value of the speed of said longitudinal wind.
[0020]
20. A method according to any one of claims 7 to 9, characterized in that, when said aircraft (10) is in ascending flight, its vertical speed Vz being non-zero, functional characteristics of said aircraft (10) being characterized in particular by a series of third performance curves according to a third formula (1 ± ') = wire a), W being a power e Vy a 25 of flight of said aircraft (10), a being a reduction coefficient, M being said estimated mass said aircraft (10), Vy being said optimum vertical climb speed of said aircraft (10), f3 being a third function represented by a series of third performance curves of said aircraft (10), the ratio (±, -) being e Vy Obtained for a vertical speed Vz of said aircraft equal to said optimum climb speed Vy of said aircraft, so that a sixth characteristic power Wvy of said aircraft (10) corresponds to said optimum speed climb Vy, said flight power W of said aircraft (10) being equal to the sum of said powers supplied by each engine (21,22) of said power plant (20) deduced from an accessory power necessary to power equipment of said aircraft (10) Po - one calculates said coefficient of reduction a such that a = = 10 - a third value A3 is calculated such that A3 = - is determined by a third curve of performance of said aircraft (10) according to the third function f3 corresponding to the flight conditions of said aircraft (10) and according to said third value A3 a fourth value A4, such that A4 = f3 (1.% - ier), - is calculated from said fourth value A4, said sixth characteristic power Wvy corresponding to said optimum climb speed Vy, such that Wvy = A4. CT.
[0021]
21. Method according to any one of claims 1 to 20, characterized in that said parameter Pi is chosen from the following list or is a logical or arithmetic combination of at least two elements of said following list: a pair total provided by said power plant (20) or a torque output from the mechanical power transmission means (23) or a torque provided by a motor (21,22), 3033316 62 - a total power supplied by said installation motor (20) or a power supplied by one of said turbine engines (21,22), - a speed of said vehicle (10) with respect to the air or a speed of said vehicle (10) with respect to the ground, - a height of said vehicle (10) relative to a reference, - a temperature inside an engine (21,22), - a rotational speed of an element of a motor (21,22).
[0022]
22. A method according to any one of claims 1 to 21, characterized in that, when said vehicle (10) is a rotary wing aircraft comprising a main rotor (11) provided with main blades (12), an anti-torque rotor (13) provided with secondary blades (14), said mechanical power transmission means (23) being a main power transmission gear driving said main rotor (11) and said anti-torque rotor (13), control means ( 24, 25, 26) being means for controlling the variation of the collective pitch and the cyclic pitch of said main blades (12) and of the collective pitch of said secondary blades (14), said parameter P1 is chosen from the following list or else is a logical or arithmetic combination of at least two elements of said following list: - a total torque supplied by said power plant (20) or a torque provided by a motor (21,22) or a rotor torque at the level of of an entry mast in rotation with said main rotor (11) or a torque at said anti-torque rotor (13), - a total power supplied by said power plant (20) or a power supplied by one of said turbine engines (2, 2, 2 ) - a speed of said aircraft (10) with respect to the air or a speed of said aircraft (10) with respect to the ground, - a height of said aircraft (10) with respect to a reference, - a temperature at the interior of a motor (21,22), 5 - a rotational speed of an element of a motor (21,22), - a steering position of said aircraft (10) or a pitch value of said blades main (12) of said main rotor (11) or said secondary blades (14) of said anti-torque rotor (13), 10 - a control position of the collective pitch of said main blades (12) of said main rotor (11), - a position controlling the cyclic pitch of said main blades (12) of said main rotor (11), - a collective pitch control position said secondary blades (14) of said anti-torque rotor (13).
[0023]
23. A method according to any one of claims 1 to 22, characterized in that said second values P, _X to which said parameter P; must be equal so that said vehicle (10) can perform a predetermined operation X form a range of values of said parameter P ,.
[0024]
24. Method according to any one of claims 2 to 23, characterized in that, at least a third value PLinst equal to an instantaneous value of each parameter P, being determined, each first value P, _lim, each second is displayed. P value, _Xet each third value P, inst for said parameter P, among at least two parameters (P ,, PJ) for which a difference between said first value P, lim and the third value PI inst 3033316 64 associated respectively with each parameter (Pi, Pj) is the weakest.
[0025]
25.Device (1) for determining characteristic parameters of the operation of a vehicle (10), said vehicle (10) comprising a powerplant (20) provided with at least one motor (21, 22) and a mechanical power transmission means (23) as well as control means (24,25,26), said device (1) comprising a plurality of sensors (5-8), at least one calculation means (2), at the at least one memory (3) and at least one display means (4), said sensors (5-8) providing measurements of different information relating to the environment of said vehicle (10) and / or the state and operation of said vehicle (10) and its equipment and / or the position and displacements of said vehicle (10) relative to its environment, each computing means (2) receiving said measurements of said sensors (5-8) and processing said information, characterized in that said memory (3) stores performance curves of said vehicle (10) and calculation instructions, said calculating means (2) using said calculation instructions to implement the method of determining parameters characteristic of the operation of said vehicle (10) according to any one of claims 1 to 24.

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EP3064437A1|2016-09-07|

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US9536358B2|2017-01-03|

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FR3033316B1|2018-04-06|

引用文献:

---

公开号 | 申请日 | 公开日 | 申请人 | 专利标题

---

DE29703902U1|1997-02-24|1997-06-05|Tegethoff Marius|Analog speedometer with additional displays|

---

US20050278084A1|2004-06-15|2005-12-15|Bernard Certain|Pilot indicator for predicting changes in the monitoring parameters of a gas turbine engine|

---

WO2006081334A2|2005-01-28|2006-08-03|Bell Helicopter Textron Inc.|Power situation indicator|

---

EP2505502A1|2011-03-30|2012-10-03|Eurocopter|Method and device to assist piloting of an aircraft, and aircraft|EP3597536A1|2018-07-17|2020-01-22|Airbus Helicopters|Method and device for predictive determination of characteristic parameters of the operation of a rotary-wing aircraft for the performance of a predetermined manoeuvre|

---

FR3110545A1|2020-05-20|2021-11-26|Airbus Helicopters|Method for optimizing the energy consumption of a hybrid helicopter in level flight|US3754440A|1972-08-16|1973-08-28|Us Navy|Helicopter lift margin determining system|

---

FR2749546B1|1996-06-07|1998-08-07|Eurocopter France|PILOTAGE INDICATOR FOR AIRCRAFT|

---

FR2749545B1|1996-06-07|1998-08-07|Eurocopter France|PILOTAGE INDICATOR FOR AIRCRAFT|

---

FR2755945B1|1996-11-19|1999-01-15|Eurocopter France|PILOTAGE INDICATOR FOR AIRCRAFT|

---

FR2756256B1|1996-11-26|1999-01-22|Eurocopter France|POWER MARGIN INDICATOR FOR A TURN WING AIRCRAFT, IN PARTICULAR A HELICOPTER|

---

FR2772718B1|1997-12-22|2000-02-11|Eurocopter France|PILOTAGE INDICATOR FOR AIRCRAFT|

---

FR2809082B1|2000-05-17|2002-09-20|Eurocopter France|POWER MARGIN INDICATOR FOR A TURN WING AIRCRAFT, IN PARTICULAR A HELICOPTER|

---

US7438259B1|2006-08-16|2008-10-21|Piasecki Aircraft Corporation|Compound aircraft control system and method|

---

US9355571B2|2008-01-23|2016-05-31|Sikorsky Aircraft Corporation|Modules and methods for biasing power to a multi-engine power plant suitable for one engine inoperative flight procedure training|

---

FR2946322B1|2009-06-04|2011-06-17|Eurocopter France|HYBRID HELICOPTER ASSISTING DEVICE, HYBRID HELICOPTER PROVIDED WITH SUCH A DEVICE AND METHOD USED THEREBY|

---

FR2997382B1|2012-10-29|2014-11-21|Eurocopter France|METHOD FOR MANAGING AN ENGINE FAILURE ON A MULTI-ENGINE AIRCRAFT PROVIDED WITH A HYBRID POWER PLANT|FR3053314B1|2016-07-01|2018-07-27|Airbus Helicopters|POWER MARGIN MARGIN INDICATOR DEVICE, GIRAVION THEREFOR, AND CORRESPONDING METHOD|

---

US10977880B2|2017-05-31|2021-04-13|General Electric Company|Hover time remaining for an aircraft|

---

US10611496B2|2018-01-22|2020-04-07|Textron Innovations Inc.|Power allocation systems for aircraft|

---

FR3090576B1|2018-12-20|2021-09-10|Airbus Helicopters|Assistance method for single-engine rotary-wing aircraft during engine failure.|

法律状态:
2016-03-21| PLFP| Fee payment|Year of fee payment: 2 |

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2016-09-09| PLSC| Publication of the preliminary search report|Effective date: 20160909 |

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2017-03-22| PLFP| Fee payment|Year of fee payment: 3 |

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2018-03-23| PLFP| Fee payment|Year of fee payment: 4 |

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2019-03-22| PLFP| Fee payment|Year of fee payment: 5 |

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2020-12-18| ST| Notification of lapse|Effective date: 20201110 |

优先权:

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申请号 | 申请日 | 专利标题

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FR1500411|2015-03-04|

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FR1500411A|FR3033316B1|2015-03-04|2015-03-04|METHOD AND DEVICE FOR DETERMINING AND OPTIMIZING CHARACTERISTIC PARAMETERS OF THE OPERATION OF A ROTARY WING AIRCRAFT|FR1500411A| FR3033316B1|2015-03-04|2015-03-04|METHOD AND DEVICE FOR DETERMINING AND OPTIMIZING CHARACTERISTIC PARAMETERS OF THE OPERATION OF A ROTARY WING AIRCRAFT|

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EP16156282.2A| EP3064437B1|2015-03-04|2016-02-18|A method and a device for determining and optimizing parameters that are characteristic of the operation of a rotary wing aircraft|

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US15/059,440| US9536358B2|2015-03-04|2016-03-03|Method and a device for determining and optimizing parameters that are characteristic of the operation of a rotary wing aircraft|