• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The present invention relates to a method and a device (1) for estimating the instantaneous mass of a rotary wing aircraft (10), said aircraft (10) comprising a powerplant (20), rotating at least one rotor main (11) and an anti-torque rotor (13), as well as a plurality of sensors (13-19). This method makes it possible, from functional characteristics and flight of said aircraft (10), atmospheric characteristics and performance curves of said aircraft (10), to determine the measured instantaneous mass Mm of said aircraft (10). In addition, this method makes it possible to consolidate said measured instantaneous mass Mm of said aircraft (10) by comparing it arithmetically or statistically during a flight of said aircraft (10) with a calculated instantaneous mass Mc determined from the variation of the amount of fuel present in said aircraft (10).
    公开号:FR3036789A1
    申请号:FR1501117
    申请日:2015-05-29
    公开日:2016-12-02
    发明作者:Serge Germanetti
    申请人:Airbus Helicopters SAS；
    IPC主号:
    专利说明:

    [0001] 1 Method for estimating the instantaneous mass of a rotary wing aircraft. The present invention is in the field of flying assistance for rotary wing aircraft.
    [0002] The present invention relates to a method and a device for estimating the instantaneous mass of a rotary wing aircraft. A rotary wing aircraft traditionally comprises a powerplant equipped with at least one engine, generally a turbine engine, and a main gearbox, the powerplant driving in rotation, via the gearbox. main power at least one main rotor for the lift or propulsion of the aircraft and possibly an anti-torque rotor. An anti-torque rotor is for example formed by a rear rotor of the aircraft or by two propeller propellers located on either side of a fuselage of the aircraft. The operation of the aircraft is generally carried out under the supervision of various characteristic parameters by means of several instruments located on a dashboard of the aircraft. These characteristic parameters are representative of the current operation of the aircraft and in particular of its powerplant and / or of each turbine engine.
    [0003] These characteristic parameters can be, for example, 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.
    [0004] For physical reasons, there are limitations on these characteristic parameters which must be taken into account at each moment of the operation of the aircraft. These different limitations may depend on external conditions as well as the mode of operation of the aircraft. The pilot of an aircraft must monitor, during the operation of this aircraft and permanently, the current values of these characteristic parameters and compare them with their respective limitations.
    [0005] These limitations are generally different depending on the flight phase of the aircraft and / or external conditions such as altitude and temperature, for example. Indeed, according to each phase of flight and / or external conditions as well as according to 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. Furthermore, the instantaneous mass of the aircraft is an important parameter in the determination of certain limitations and / or of certain current values of these characteristic parameters such as the power that each engine of the power plant must provide for the realization of the flight or a particular maneuver. In particular, the power plant must provide sufficient mechanical power so that the lift of the main rotor is at least opposed to the instantaneous mass of the aircraft and thus ensures the lift of the aircraft. Several methods for estimating the mass of the aircraft exist today. An estimate of the mass of the aircraft is generally determined before the take-off and the flight of the aircraft, by adding for example the mass of the empty aircraft, the mass of the on-board fuel, the mass of the crew and the payload transported. This estimate of the weight of the aircraft before take-off is then used during the flight of the aircraft. This estimate of the mass of the aircraft then does not take into account the fuel consumption of the aircraft. This estimate of the mass of the aircraft is thus constant and identical throughout the flight of the aircraft and the difference vis-à-vis the actual instantaneous mass of the aircraft increases throughout the flight. Consequently, this estimate of the mass of the aircraft is overvalued and then results in non-optimized performance of the aircraft. An estimate of aircraft weight may also be determined during the flight of the aircraft, generally by deducting the amount of fuel consumed from the aircraft's mass estimate prior to take-off. However, this estimate of the mass of the aircraft is generally not very accurate and leads to an understanding of the non-optimized performance of the aircraft. A safety margin is generally used to determine this estimate of the mass of the aircraft and then leads to an estimate of the mass of the overvalued aircraft. This safety margin makes it possible, for example, to take into account approximations in the determination of the mass of the aircraft. This estimate of the overvalued mass goes in the direction of safety, the mechanical power actually necessary for the flight of the aircraft being less than the necessary power determined according to this estimate of the mass of the aircraft. On the other hand, for certain maneuvers requiring a large mechanical power, this determined power which is overvalued may be greater than the maximum power available at the power plant, while the actual mechanical power required is less than this maximum power available. for each engine. As a result, such maneuvers are not performed by the pilot who wrongly believes that the power plant can not provide sufficient total power. This estimation of the mass can thus result in a more or less significant reduction of the flight envelope of the aircraft. However, this estimate of the mass of the aircraft can also be underestimated, in particular following a bad identification of the on-board off-fuel weight, constituted for example by the passengers of the aircraft and their luggage. This undervalued estimate of the weight of the aircraft then poses a risk to the flight safety of the aircraft, as opposed to an overvalued estimate of the accidents that occurred as a result of these types of errors. The object of the present invention is therefore to propose a method and a device making it possible to overcome the abovementioned limitations, in particular by making it possible to determine an accurate and reliable estimate of the instantaneous mass of a rotary wing aircraft underway. flight. According to the invention, a method for estimating the instantaneous mass of a rotary wing aircraft comprises several steps performed during a flight: flight characteristics of the aircraft are measured, such as a speed horizontal Vh and a vertical speed Vz of the aircraft relative to the air and a height Hz of the aircraft relative to the ground, - it measures power characteristics of the aircraft, such as the torque CR and the main rotational speed NR of the main rotor as well as the torque CRAC and the real rotational speed NRAc of the anti-torque rotor, are measures of the atmospheric characteristics relating to the environment of the aircraft, such as the atmospheric pressure Po and the temperature To of the air around the aircraft, the flight power W of the aircraft is determined, an operating point of the aircraft is determined on at least one series of performance curves of the aircraft. aircraft in function of the flight characteristics of the aircraft, the atmospheric characteristics and the flight power W of the aircraft, and the measured instantaneous mass Mm of the aircraft is deduced therefrom. A rotary wing aircraft comprises a plurality of sensors providing measurements of various information relating to the environment of the aircraft and / or to the state and operation of the aircraft and its equipment and / or to the position and the movements of the aircraft relative to its environment.
    [0006] For example, a first type of sensor makes it possible to measure first information relating to the flight of the aircraft such as the speed of the aircraft relative to the air. The speed of the aircraft with respect to the air is for example measured in a decomposed manner according to a horizontal speed Vh and a vertical speed Vz, the horizontal and vertical directions being defined in a terrestrial reference. Moreover, and for the sake of simplifying the description, the term "speed" will be used to designate a speed of the aircraft relative to the air. Indeed, the flight of the aircraft 30 is characterized by aerodynamic effects related to this speed 3036789 6 of the aircraft relative to the air which takes into account the effect of the wind, this wind depending on whether it is face or rear vis-à-vis the aircraft increasing or reducing the lift of the aircraft and therefore conversely reducing or increasing its power requirement.
    [0007] A second type of sensor makes it possible to measure second information relating to the position of the aircraft. A sensor of this second type is for example a radioaltimeter determining the height Hz of the aircraft relative to the ground. These first and second aircraft speed information and height Hz are flight characteristics of the aircraft. A third type of sensor makes it possible to measure third information relating to the environment of the aircraft such as the atmospheric pressure and the temperature outside the aircraft. These third pieces of information constitute atmospheric characteristics relating to the environment of the aircraft, which influence the aerodynamic behavior of the main rotor as well as the operation of each engine of the aircraft.
    [0008] Finally, a fourth type of sensor makes it possible to measure fourth information relating to the performance of the aircraft, which concerns in particular the main rotor, the anti-torque rotor and the powerplant of the aircraft. This fourth information is, for example, the torque CR and the actual rotational speed NR of the main rotor or the torque CRAC and the actual rotational speed NRAc of the anti-torque rotor. These fourth information can also be the torque and rotational speed of a main output shaft of the power plant or each engine. These fourth pieces of information constitute measurements that will make it possible to reconstitute the power characteristics of the aircraft. These power characteristics make it possible in particular to determine the powers consumed respectively by the main rotor and the anti-torque rotor or the power supplied by the power plant or even by each motor. We can then deduce the power W used for the flight of the aircraft itself, which will be more simply designated in the following description by the term "power of flight".
    [0009] This flight power W of the aircraft can be determined according to various calculations. First, this flight power W is distributed between the main rotor and the anti-torque rotor to provide lift and movement of the aircraft. This flight power W is therefore equal to the sum of the powers consumed by the main rotor and the anti-torque rotor. The instantaneous power consumed by the main rotor can be defined in a known manner according to the formula WRp = CR. NR. The instantaneous power consumed by the anti-torque rotor 20 may be determined according to an analogous formula WRAc = - C RAC-N RAC - This instantaneous power consumed by the anti-torque rotor is 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 and can then be estimated in a known manner. For example, this instantaneous power consumed by the anti-torque rotor can be determined according to the speed of advance of the aircraft. When this forward speed is zero, the anti-torque rotor opposes this torque alone, whereas when this forward speed 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, to reduce the instantaneous power consumed by the 5 anti-torque rotor. Moreover, the total power supplied by the power plant of the aircraft is distributed firstly in a flight power W for the actual flight of the aircraft and secondly in an accessory power Wac, to feed different equipment 10 of the aircraft. As a result, the flight power W is equal to the engine power supplied by the power plant from which this accessory power is subtracted. This accessory power is used for example to power the electrical equipment of the aircraft such as the avionics, hydraulic and electrical equipment or air conditioning in the cabin. This accessory power Wacc is thus mainly composed of electric power and hydraulic power and can be determined in a known manner. The total power supplied by the power plant can be determined by the torque C ,,, and the rotational speed Nin, of a main output shaft of the power plant, such as Wim. = Cim.NRim. This total power supplied by the power plant can also be determined by the sum of the powers provided by its motors. 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 of the gases to the engine. inlet of the free turbine of this turbine engine 30 and the speed of rotation Ng of the gas generator of this turbine engine. The instantaneous power Wminst supplied by each engine can then be estimated such that Wminst = Cm.Ng, where Cm is the engine torque relative to the gas generator. Then, depending on the flight characteristics of the aircraft, atmospheric characteristics and flight power W, determine an operating point of the aircraft using one or more performance curves of the aircraft. 'aircraft. The performance curves of an aircraft are provided with the flight manual of the aircraft. They are specific to each aircraft and correlated with the power curves of the engines of this vehicle, taking into account the location of each engine in the aircraft. The performance curves make it possible to characterize each aircraft and in particular the flight power W that the aircraft must supply or may provide depending on atmospheric criteria, such as the atmospheric pressure Po and the temperature To of the air. around the aircraft, only parameters of the aircraft, such as its mass or speed. Several performance curves exist for each aircraft according to the different flight phases of the aircraft, and mainly for stationary flights, level flights and ascents. In order to define which series of performance curves should be used by the method according to the invention, a selection algorithm can be used. This selection algorithm uses, for example, the values of the horizontal speed Vh and the vertical speed Vz of the aircraft in order to determine whether the aircraft is for example hovering, climbing or level flight.
    [0010] Indeed, during a hover, the horizontal speed Vh and the vertical speed Vz of the aircraft are substantially zero. For an ascending flight, the vertical speed Vz and the horizontal speed Vh of the aircraft are non-zero. The term "ascending flight" is understood to mean a rising or falling flight of the aircraft with a non-zero forward speed. Pure vertical flight, that is to say with zero forward speed, is a particular flight meeting specific criteria. During a level flight, the vertical speed Vz of the aircraft 10 is substantially zero and its horizontal speed Vh is non-zero. Advantageously, these three phases of flight, namely hovering, ascending flight and level flight, cover about 90% of the cases of flight of an aircraft. The other flight cases are generally particular and transient phases such as pure vertical flights, flight phases with significant acceleration and turns with a high turn rate, for example greater than 10 ° / s (ten degrees per second). second). Finally, from these aircraft performance curves, the flight characteristics, the atmospheric characteristics and the flight power W are deduced, the measured instantaneous mass Mm of the aircraft in a precise and reliable manner. In this way, the measured instantaneous mass Mm can be accurately and reliably determined mainly during these three particular flight phases of the aircraft. Advantageously, the measured instantaneous mass Mm can be precisely and reliably determined during the majority of the flight time of the aircraft. Indeed, this instantaneous mass measured Mm can not be accurately and reliably determined during the other phases of flight which are transient and therefore short durations. Thus, the absence 3036789 11 of this instantaneous mass measured Mm accurate and reliable is short and does not interfere with the flight of the aircraft. Advantageously, the knowledge of the measured instantaneous mass Mm then makes it possible to optimize the use of the aircraft by using this instantaneous mass measured Mm in place of an estimate of the mass of the aircraft. For example, this instantaneous measured mass Mm of the aircraft makes it possible to determine more precisely the flight power required for the realization of a particular maneuver of the aircraft, such as a flight in slow descent for a landing, or the realization of a level flight with optimized fuel consumption, that is to say with a maximum flight time or a maximum range. Moreover, from a maximum total power available at the power plant and this measured instantaneous mass Mm, it is possible to accurately determine a maximum mass that can be transported by the aircraft. Indeed, it is possible by using the performance curves of the aircraft and by knowing the maximum power available at the power plant to determine the maximum total mass of the aircraft possible in flight. It is then possible to deduce the maximum transportable mass which is the difference between this maximum total mass and the measured instantaneous mass Mm of the aircraft. Finally, the instantaneous mass measured Mm of the aircraft as well as this maximum transportable mass can be displayed on an instrument or a screen present on the dashboard of the aircraft so that the pilot of the aircraft becomes aware of it. In the case where the aircraft is hovering, the functional characteristics of the aircraft are in particular defined by a series of first curves of NpReo dormances according to W (NR0) 3 a first formula - = k. (-.) 1. In these o-NR NR conditions, W is the flight power of the aircraft, where o is a reduction coefficient, k is a ground influence coefficient, Min 5 is the instantaneous measured mass of the aircraft, NR0 is a rotational speed of the main rotor, NR is the actual rotational speed of the main rotor and fi is a first function represented by a series of first performance curves of the aircraft having ordered a first value A1 = 10 and for abscissa a second value A2 = o. - UIR - (NR0) 3 (NR0) 2 .NR) The value of this coefficient of influence k of the ground is defined by a curve as a function of the height Hz of the aircraft relative to the ground. This coefficient of influence k of the ground makes it possible to quantify the effect of the proximity of the ground on the behavior of the aircraft and on the flight power of the aircraft as a function of the height Hz of the aircraft relative to the ground. Indeed, near the ground, the main rotor blast is returned by the ground on the fuselage and the main blades of this main rotor modifying the behavior of the aircraft. The aircraft is said to be in ground effect, defined by the acronyms DES or IGE for the English expression "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. Above this DES area, that is to say, for a height relative to 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 Ground Effect".
    [0011] The coefficient of influence k is equal to unity when the aircraft is in an HES zone. This coefficient of influence k varies according to the height Hz of the aircraft 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, 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.
    [0012] Thus, in order to determine the instantaneous measured mass Min of the aircraft in the case of a hovering flight, the reduction coefficient o- such that o- = (-P12) is calculated, Po and n respectively being the atmospheric pressure. and the temperature around the aircraft, the atmospheric pressure Po 15 being expressed in millibar (mb) and the temperature Ta being expressed in kelvin (K), -on calculates a first value A / such that A1 = 1 '. (NR0 3, cr NR -on determines, by means of a first performance curve of the aircraft according to the first function fi corresponding to the flight conditions of the aircraft and as a function of the first value A1, a second value A2 such that A2 _ = rn) 2] = mm NRO 2 to NR Ro. () O- NR k -on determines the influence coefficient k according to the height Hz of the aircraft relative to the ground, and 25 -on calculates from of the second value A2 the measured instantaneous mass A / 1 'of the aircraft, such that a (NNRR0) 2 Mm = k.A2.
    [0013] In the case where the aircraft is in ascending flight, the functional characteristics of the aircraft are defined in particular by two sets of seconds and third performance curves respectively according to a second formula 5 w, = f2 (Vz) and a third formula (-) = f3mm). In Vy these conditions, W is the flight power of the aircraft, cr is the reduction coefficient, Mm is the instantaneous measured mass of the aircraft, Vz is the vertical speed of the aircraft and Vy, commonly referred to as The term "optimum climb speed" is the horizontal component of the aircraft speed that achieves the best vertical climb speed. The second function f2 relates to an ascending flight while the third function f3 relates to a level flight. These second and third functions f2, f3 are respectively represented by two series of aircraft performance curves. A first ratio (f) is obtained for any vertical speed Vz of the aircraft, the flight power W corresponding to this speed w, vertical Vz, while a second ratio (-) is obtained for a Vy speed Vh of the aircraft only horizontal and equal to the optimum climb speed Vy, the flight power W 'corresponding to this horizontal speed Vh'. Thus, in order to determine the measured instantaneous mass Mm of the aircraft in the case of an ascending flight, a second performance curve of the aircraft is determined according to the second function f2 corresponding to the flight conditions of the aircraft. aircraft and as a function of the vertical speed Vz of the aircraft a third value A3 such that A3 = f2YZ) = e 11471 - the reduction coefficient a is calculated as that a = - is calculated from the third value A3, the flight power 5 W and the reduction coefficient has a fourth value (17) / frw, 114 such that A4 = cr A3 = - is determined by a third performance curve of the aircraft according to the third function f3 and corresponding to the flight conditions of the aircraft and according to the fourth value A4 which is such that A4 = Rn 1 = f3 (& 10.), a fifth Vy value As such that As = (1 ± -ner) , and - from the fifth value As, the measured instantaneous mass Mm of the aircraft such that Mm = A5.u. In the case where the aircraft is in level flight, the 15 functional characteristics of the aircraft are defined in particular by two series of third and fourth performance curves according to the third formula () = f3 (lz-n) and a Vy o- fourth formula, = f4. [() 3 - In these conditions, W Vh mrni (w-cr) vy is the flight power of the aircraft, a- is the reduction coefficient, Mm is the mass measured instantaneous aircraft, Vh is the horizontal speed of the aircraft, V. y is the optimum speed of climb of the aircraft.
    [0014] The third and fourth functions f3, f4 relate to a level flight and are respectively represented by two series of aircraft performance curves. The first ratio (± 70) is obtained for any horizontal speed Vh of the aircraft, the flight power W corresponding to this horizontal speed Vh, while the second ratio () is obtained for a horizontal speed Vy Vh ' of the aircraft only horizontal and equal to the optimum climb speed Vy of an optimum vertical climb speed of the aircraft, the optimum power W 'corresponding to this horizontal speed Vh'. The measured instantaneous mass Min of the aircraft is determined, which makes it possible to simultaneously resolve the third and fourth formulas, according to the values of the flight power W and the values of the horizontal speed Vh of the aircraft.
    [0015] For example, to determine the measured instantaneous mass Mn of the aircraft in the case of a level flight, the sixth value A6 is calculated such that A6 = (Vh) 3 as a function of the horizontal speed Vh and the optimum climb speed Vy of the aircraft, this optimum climb speed Vy being a known characteristic of the aircraft 10 - a series of fourth performance curves of the aircraft is determined according to the fourth function f4 and in FIG. according to this sixth value A6 of the first pairs formed of a third value A3 and a fifth value As PL7 25 such that A3 = "and they = - the reduction coefficient a- such that o- = (e) is calculated , From each third value A3, the flight power W and the reduction coefficient o- a (w) / w, fourth value A4 such that A4 = (7-ji A = [(- -) Vy1, in order to 3 e form second pairs formed of a fourth value A4 5 and a fifth value r As, determining the second pair formed by a fourth value A4 and a fifth value AS solving the third formula by a third performance curve of the aircraft according to the third function f3 and corresponding to the flight conditions of the aircraft, and - from the fifth value As of this second pair solving the third formula, the measured instantaneous mass Min of the aircraft such that Mm = A5.0- is calculated. In order to obtain an estimated instantaneous mass M of the reliable and consolidated aircraft, the measured instantaneous mass Mm of the aircraft thus obtained can be compared with a calculated instantaneous mass iii, which is determined by another method, for example from the variation of the quantity of fuel present in the aircraft. This comparison of the measured instantaneous mass Mm and the calculated instantaneous mass Mc. can be done during a flight of the aircraft arithmetically or statistically. Advantageously, this comparison makes it possible to have a continuity of this information of an instantaneous mass estimated M.
    [0016] Thus, when the measured instantaneous mass Mm and the instantaneous mass calculated Mc- are available simultaneously, they can be compared and obtain the estimated instantaneous mass M. On the other hand, when the instantaneous mass measured Mm is not available, the aircraft in a non-flight phase covered by usable performance curves, no comparison is possible. However, the availability of instantaneous mass calculated Mc. Nevertheless, it is possible to provide an estimated instantaneous mass M equal to this calculated instantaneous mass M. Since the unavailability of the measured instantaneous mass Mn is only transient, an estimated instantaneous mass M can quickly be provided again. An aircraft comprises at least one tank, and generally several tanks, in which the fuel is stored. This instantaneous mass calculated Mc of the aircraft can be determined from the amount of fuel remaining in the aircraft or the amount of fuel consumed from an initial time to such as the takeoff of the aircraft. For example, an instantaneous volumetric flow rate of 15 Dv for consumption of a fuel supplying the power plant is measured and the instantaneous temperature T of this fuel is measured. An instantaneous mass flow rate Dm of the fuel is then calculated. The mass Mcc of the fuel consumed is then determined from an initial moment to by integration of the instantaneous mass flow rate Dm, of the fuel from the initial moment to. We also know the total initial mass Mo of the aircraft at the initial moment to. Finally, the calculated instantaneous mass Mc is calculated which is equal to the total initial mass Mo at which the mass Mcc of the consumed fuel is subtracted.
    [0017] It is also possible to measure a volume VcR of the fuel remaining in all the tanks and the instantaneous temperature T of this fuel. The MCR mass of this fuel remaining in all the tanks is then calculated. The McRto mass is also known of the fuel remaining in the tank set and the total initial mass Mo of the aircraft at the initial moment. The mass Mcc of the fuel consumed is then determined from an initial moment to subtracting the mass Mau. fuel of the mass McRt of the fuel remaining at the moment t. The calculated instantaneous mass Mc is finally calculated. which is equal to the total initial mass Mo of the aircraft at the initial moment to which the mass Mcc of the consumed fuel is subtracted. Similarly, it is possible to measure a volume VcR of the fuel remaining in all the tanks and the instantaneous temperature Ti of this fuel. The mass Mo is then calculated fuel remaining in all the tanks. Also known is the initial mass Mi off fuel of the aircraft at the initial time to. The calculated instantaneous mass Mc is then calculated. which is equal to the initial mass Mi excluding fuel of the aircraft to which the McR mass of the remaining fuel is added.
    [0018] The measured instantaneous mass Mm and the calculated instantaneous mass Mc can then be compared in the course of the flight of the aircraft arithmetically and thus determine a first difference between them. Thus, if this first deviation is less than or equal to a predetermined error threshold, the measured instantaneous mass Mm is considered as an estimated instantaneous mass M which is reliable and usable, whereas if this first deviation is greater than this error threshold, the measured instantaneous mass Mm is considered as an instantaneous mass estimated M not usable and an alert is issued to inform the pilot of the aircraft.
    [0019] Furthermore, if this first deviation is large and positive, it can also make it possible to identify either an unloading of a payload or the release of a load under sling. On the other hand, if this first deviation is negative and significant in absolute value, it can be deduced that a loading of a payload or the hooking of a load under sling has been carried out.
    [0020] Indeed, the measured instantaneous mass Mm determined from the aircraft performance curves takes into account any change in the total mass of the aircraft, including the loading or unloading of a payload. In contrast, the instantaneous mass calculated Mc. takes into account only the variation of the fuel quantity of the aircraft, regardless of the total weight of the aircraft. In the case of a first large deviation in absolute value, the pilot can confirm the loading or unloading of a payload, thereby confirming the reliability of the measured instantaneous mass Mm. It is also possible to compare a first difference between the mass instantaneous measured Mm and an initial mass as well as a second difference between the instantaneous mass calculated Mc. 15 and this initial mass. The instantaneous mass measured Mm and the instantaneous mass calculated Mc. are for example determined at a time t and the initial mass is determined at the initial time to. This initial time to corresponds to a particular moment of the flight, such as the loading or unloading of a payload for example, and this initial mass is a fixed and constant value. This initial moment to is preferably the moment of takeoff of the aircraft and the initial mass is the mass of the aircraft at takeoff. We can then determine a second difference between the first difference and the second difference. If this second difference is less than or equal to the error threshold, the measured instantaneous mass Mm is considered as an instantaneous mass estimated M reliable and usable whereas if this second difference is greater than this error threshold, the instantaneous mass measured Mm is considered an instantaneous mass estimated M not usable and an alert is issued to inform the pilot of the aircraft.
    [0021] In addition, this second difference, while important in absolute value, may also make it possible to identify as previously mentioned the loading or unloading of a payload of the aircraft.
    [0022] The values of the measured instantaneous mass Mn and the calculated instantaneous mass Mc can also be compared statistically. during a flight of the aircraft to determine a value of the estimated instantaneous mass M consolidated, reliable and accurate.
    [0023] It is thus possible to compare the evolution of the instantaneous mass of the aircraft by virtue of the different values of the measured instantaneous mass Mn and the calculated instantaneous mass Mc. of the aircraft during a flight by statistical analysis. One can also compare by statistical analysis the evolution of the first and second 15 differences during a flight. For example, it is possible to calculate a first average or a second average, respectively, of the first deviations or second deviations determined between the instant t and the initial instant to during the flight of the aircraft. Then, this first average or this second average can be compared with the error threshold in order to determine whether the measured instantaneous mass Mm can be considered as an instantaneous mass estimated M reliable and consolidated and, therefore, usable. In addition to the first average or the second average, a first or second standard deviation associated with the first and second deviations can be calculated respectively, and then the first or second standard deviation can be analyzed to determine whether the measured instantaneous mass Mn can be considered as an instantaneous mass estimated M reliable and consolidated.
    [0024] It is also possible, from the values of the measured instantaneous mass Mni and the calculated instantaneous mass Mc. determined between a time t and the initial time to, recalculate the initial mass of the aircraft at the initial time to and thus verify that this initial mass is constant between the instant t and the initial moment te. Such a statistical analysis can be performed from the values of the instantaneous mass measured Mm and the fuel consumption of the aircraft during a flight. Moreover, such a statistical analysis can be performed by means of at least one Kalman filter and thus make it possible to consolidate the estimated instantaneous mass M of the aircraft by comparing the values of the instantaneous mass measured Mm and the instantaneous mass calculated Mc. for example by the mass of fuel consumed Mcc by the aircraft. A measurement vector Zni is firstly determined such that (Mm), Min being the instantaneous measured mass of the licc aircraft and Mcc being the mass of fuel consumed by the aircraft, a vector of state X to be determined such that X (1140), M being an instantaneous estimated mass of the aircraft and Mo 20 being the initial total mass of the aircraft, -on defines an equation of state i = AX + B.Dmi, with A = 111) (0) and B = [1j, Dm, being the instantaneous mass flow rate of the fuel which is equal to the derivative of the consumed fuel mass Mcc of the aircraft, being the derivative of the state vector X, - we define a measurement equation Zm = CX, with C = [_ 11 ° - we apply the equation of state and the measurement equation to a Kalman filter in order to determine the state vector X, and consequently, the estimated instantaneous mass M and the total initial mass Mo of the aircraft In a similar way, using at least one Kalman filter, it is possible to It is also possible to consolidate the estimated instantaneous mass M of the aircraft through the remaining McR fuel mass in the tanks of the aircraft. The present invention also relates to a device for estimating the instantaneous mass of a rotary wing aircraft. A rotary wing aircraft generally comprises a power plant having at least one engine and a main power gearbox, the main power gearbox rotating at least one main rotor and one anti-torque rotor. The aircraft also comprises a plurality of sensors providing measurements on the environment of the aircraft as well as on the operation of the aircraft and such equipment, such as its powerplant and the main rotor in particular. This device for estimating the instantaneous mass comprises at least one calculation means and at least one memory, a memory storing aircraft performance curves and calculation instructions. The calculating means receives the measurements of the sensors and can apply the calculation instructions in order to implement the method for estimating the instantaneous mass of a rotary wing aircraft previously described.
    [0025] This device for estimating the instantaneous mass may comprise at least one Kalman filter in order to statistically compare the evolution of the measured instantaneous mass Mm and the evolution of the calculated instantaneous mass Mc. of the aircraft during a flight and thereby to determine a consolidated, accurate and reliable value of the estimated instantaneous mass M of the aircraft.
    [0026] The invention and its advantages will appear in more detail in the following description with examples given by way of illustration with reference to the appended figures which represent: FIG. 1, a rotary wing aircraft provided with a device for estimating the instantaneous mass of the aircraft, FIG. 2, a block diagram of a method for estimating the instantaneous mass of the aircraft; FIGS. 3 to 7, performance curves of the aircraft; FIG. 8, an architecture using a Kalman filter for determining the estimated instantaneous mass M, and FIG. 9, a curve representing the evolution of the initial mass Mo estimated by this architecture. 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.
    [0027] The aircraft 10 also comprises a plurality of sensors 4-9, a dashboard 15, a powerplant 20 provided with two turbine engines 21,22 and a main power transmission gearbox 23 rotating the main rotor 11 The sensors 4-9 make it possible to measure different information relating to the environment of the aircraft 10, the state and the operation of the aircraft 10 and the main rotor 11. The aircraft 10 finally comprises a device 1 for estimating the instantaneous mass of the aircraft 10 provided with a calculation means 2 and a memory 3.
    [0028] Atmospheric sensors 4,5 make it possible to measure atmospheric characteristics relating to the environment of the aircraft 10 and are, for example, a means for measuring the atmospheric pressure Po and a means for measuring the outside temperature Tc around the aircraft 10. The sensors 6.7 make it possible to measure flight characteristics relating to the speed and the position of the aircraft 10. The speed sensor 6 measures, for example, the speed relative to the air of the aircraft. aircraft 10 in three preferred directions of the aircraft 10 10 such as longitudinal, transverse and elevation directions. This speed relative to the air of the aircraft 10 can be decomposed into a horizontal velocity Vh and a vertical velocity Vz, the horizontal and vertical directions being defined in a terrestrial reference by means of attitude determination means and of heading 16 as an AHRS device for the English expression "Attitude and Heading Reference System" that the aircraft 10 comprises. Moreover, the vertical speed Vz is generally determined by using a calculation of the derivative of the variation of the altitude of the aircraft measured by measurements of the static pressure. The sensor 7 is for example a radioaltimeter determining the height Hz of the aircraft 10 relative to the ground. Finally, the power sensors 8, 9 and 8 ', 9' make it possible to measure power characteristics relating respectively to the operations of the main rotor 11 and the rear rotor 13 of the aircraft 10. The sensors 8, 9 are, for example means for respectively measuring the torque CR and the actual rotational speed NR of this main rotor 11. Similarly, the sensors 8 ', 9' are measuring means respectively of the CRAC torque and the actual speed of rotation NRAc of this tail rotor 13. Additional sensors not shown in the figures can also make it possible to measure power characteristics relating to the operation of the power plant. These additional sensors are, for example, means for measuring the torque Cim and the NRim rotation speed of a main output shaft of the power plant 20.
    [0029] The device 1 can implement a method for estimating the instantaneous mass of the aircraft 10 whose block diagram is shown in FIG. 2. The memory 3 stores aircraft performance curves and flight instructions. calculation. The calculation means 2 applies these 15 calculation instructions in order to implement this method. The process according to Figure 2 comprises six steps. During a first step 51, flight characteristics of the aircraft 10 are measured, such as the horizontal speed Vh, the vertical speed Vz of the aircraft 10 via the speed sensor 6 and the height 10 of the aircraft 10 relative to the ground via the radioaltimeter 7. During a second step 52, the power characteristics of the aircraft 10 are measured, such as the torque CR and the actual speed of rotation. NR of the main rotor 11, via the power sensors 8.9 and the CRAC torque and the actual rotation speed NRAc of said anti-torque rotor 13, via the power sensors 8 ', 9'. During a third step 53, atmospheric characteristics relating to the environment of the aircraft are measured, such as the atmospheric pressure Po and the temperature To of the air around the aircraft 10 via atmospheric sensors 4.5. During a fourth step 54, the flight power W of the aircraft 10 is determined. This flight power W, ensuring the lift and the displacements of the aircraft 10, is distributed between the main rotor 11 and the Rear rotor 13. This flight power W is therefore equal to the sum of the powers consumed by the main rotor 11 and by the rear rotor 13. In addition, this flight power W is equal to the total power supplied by the installation 20 to which is subtracted an accessory power W acc needed to supply the different equipment of the aircraft 10. During a fifth step 55, an operating point 15 of the aircraft 10 is determined on at least one set of aircraft performance curves 10 as a function of the flight characteristics of the aircraft 10, the atmospheric characteristics and the power of flight W. These series of performance curves consist of several courts bes, as shown in Figures 4 to 7, depending on atmospheric characteristics or characteristics of the aircraft 10 such as its mass for example. These series of performance curves are specific to the aircraft 10 and make it possible to characterize the operation of the aircraft 10 and in particular its flight power W according to the different flight phases. These series of performance curves cover the main flight phases of the aircraft 10 which are the hover for the first curves shown in FIG. 4, the ascending flight for the second curves represented in FIG. 5 or the flight in FIG. bearing for the 3036789 28 third and fourth curves shown respectively in Figures 6 and 7. In order to define the flight phase of the aircraft 10 and, consequently, the series or series of performance curves to be taken into account, 5 can use a selection algorithm using the values of the horizontal speed Vh and the vertical speed Vz of the aircraft 10. In fact, a hover corresponds to a horizontal speed Vh and a vertical speed Vz of the aircraft 10 substantially zero. An ascending flight, which concerns both a rising and descending flight, corresponds to a vertical speed Vz and a horizontal velocity Vh that are not zero. A level flight corresponds to a vertical speed Vz substantially zero and a horizontal velocity Vh nonzero. During a sixth step 56, the measured instantaneous mass Mm of the aircraft 10 is deduced from the performance curves represented in FIGS. 4 to 7, the flight characteristics of the aircraft 10, the atmospheric characteristics and the flight power W. In the case where the aircraft 10 is hovering, the series of first performance curves shown in FIG. 4 according to a first formula (NR0) 3 = k.erm (NRol. n- "NR Each first curve of this series corresponds to a pair of atmospheric pressure and air temperature around the aircraft 10.
    [0030] The reduction coefficient Cf such that = (Po -) is first calculated as a function of the atmospheric pressure Po expressed in millibar (mb) and the temperature To of the air around the aircraft expressed in kelvin (K) measured via atmospheric sensors 4,5. Then, a first value A / is calculated such that A_ = if (NR0) 3 NR0 is a rotational target speed of the main rotor 11. Then, it is determined, thanks to the first curve of the performance of the aircraft 10 corresponding to the atmospheric pressure PO and the temperature Te of the air around the aircraft 10, and as a function of the first value A /, a second value Az 2 Mm (NR (A 2 10 such that A2 = f ri 2 (1 '; Rc) 1 = a, as indicated in FIG. 4. The influence coefficient k is then determined according to the height Hz of the aircraft 10 with respect to the ground measured for example by The coefficient of influence k is specific to each aircraft 10 and defined by a ground influence curve 15 shown in FIG. 3 as a function of the height Hz of the aircraft 10 with respect to the ground. expressed as a function of the number of diameters of the main rotor 11. On the curve shown in FIG. The influence k is between 0.9 and 1.1. The coefficient of influence k is generally equal to unity, which corresponds to a hovering flight of the aircraft 10 in an HES zone, that is to say a zone excluding ground effect. Finally, the measured instantaneous mass X of the aircraft 10, such as NR Mm = k, is calculated from the second value A2. A2. In the case where the aircraft 10 is in ascending flight, a series of second performance curves shown in FIG. 5 is used according to a second formula 30 = f2 (Vz). also a series of third performance curves shown in Figure 6 according to a third formula (- / = f3 (1771) This third formula Vy corresponds in fact to a level flight of the aircraft 10 with a horizontal speed Vh'equal at the optimum climb speed V. y of the aircraft 10, the second ratio C1 being obtained for such e Vy speed Vh 'of the aircraft 10. Each second and third curve corresponds to a pair of atmospheric pressure and temperature air around the aircraft 10.
    [0031] Firstly, the second performance curve corresponding to the atmospheric pressure Po and the temperature Tg of the air around the aircraft 10 is determined first and, depending on the vertical speed Vz of the aircraft 10, a third value A3 such that A3 = f2 (11Z) = a p47, as shown in FIG. 5. Then, the reduction coefficient u is calculated such that ci = To Then, the third value A3 is calculated, of the flight power W and the reduction coefficient u a fourth (17 "w, value A4 such that A4 = e A = [(-) 3 Vy is then determined by virtue of the third performance curve corresponding to the pressure atmospheric Po and the temperature To of the air around the aircraft 10, and as a function of the fourth value A4 which is such that A4 = [r) = f3 (e1 ÷ -1), a e Vy fifth value As such that A5 = (11rna), as indicated on the 3036789 31 figure 6. The fourth and fifth values concerned are respectively In FIG. 6, we finally write down A4.2 and A5.2. Finally, from the fifth value As and the reduction coefficient, the measured instantaneous mass Mm, such that Mm = A5.0, is calculated. In the case where the aircraft 10 is in level flight, two series of third and fourth performance curves represented in FIGS. 6 and 7 are used respectively according to the third formula (-1v2 -) = f3 H and a fourth formula e Vy 10 (nw, = [(--3 Each fourth curve corresponds to a (- Vh) Mml e) Vy Vy o- f4 third ratio C: L7.no The measured instantaneous mass Mm is then determined which makes it possible to solve simultaneously the third and fourth formulas, according to the values of the flight power W and the horizontal speed Vh of the aircraft 10. In order to solve these two formulas simultaneously, for example, the sixth value A6, such as Vh3 A6 = (-Vy) as a function of the horizontal speed Vh and the optimum climb speed Vy of the aircraft 10.
    [0032] Then, using the series of fourth performance curves and based on this sixth value A6, the first pairs Bi formed of a third value A3 and a fifth value (W, value A5 such that A3 = pif) are determined. LT) 1) ify and A5 = & n. Thus, as shown in FIG. 6, five first pairs 115.i respectively formed of a third value A3.1-A3.5 and a fifth value As./-A5.5 are obtained. Each fifth value A5.1 corresponds to a fourth performance curve and each third value A3.1 forms with the sixth value A6 a point of this fourth curve. Then, we calculate the reduction coefficient o- where To = Then calculates from each third value A3.1-143.5, the flight power W and the reduction coefficient o- (If) / a fourth value A4.1-A4.5 such that A4 = A = K-11 1. On 3 ° - V y A4i then forms second couples Ci [A5. formed respectively of a fourth value A4.1-A4.5 and a fifth value As./-A5.5. Of these second couples Ci, only one defines, as indicated in FIG. 6, an operating point situated on the third performance curve corresponding to the atmospheric pressure Po and to the temperature To or located close to this third one. performance curve. The second pair C2 formed of a fourth value A4.2 and a fifth value A5.2 solving the third formula is thus determined as a function of the atmospheric pressure Po and the temperature Ta. The other second pairs Ci, C3-05 do not indeed make it possible to solve the third formula taking into account the atmospheric pressure Po and the temperature To. For example, it can be seen that the operating point corresponding to the second pair C3 is well located on a fourth curve of 30 performances, but which does not correspond to the atmospheric pressure Po and the temperature To. Finally, from the fifth value A5.2 of this second pair C2 the measured instantaneous mass Mm of the aircraft 10 5 such that Mm = This estimation method thus makes it possible to determine in flight the instantaneous mass measured Mm accurately. Advantageously, this instantaneous mass measured Mm can be used to optimize the use of the aircraft 10 by determining for example with precision the power of flight necessary for the realization of a particular maneuver of the aircraft 10 or the payload transportable by the aircraft 10. This measured instantaneous mass Mm can then be displayed on an instrument or a screen present on the dashboard 15 of the aircraft 10 in order to inform the pilot. In addition, in order to obtain an estimated instantaneous mass M of the reliable and consolidated aircraft, the measured instantaneous mass Mm can be compared with a calculated instantaneous mass Mc. of the aircraft 10 obtained differently, for example from the variation of the amount of fuel present in the aircraft 10. This comparison of the measured instantaneous mass Mm and the instantaneous mass calculated Mc can be made during a flight of the aircraft 10 arithmetically or statistically. The aircraft 10 shown in FIG. 1 comprises two tanks 25, 26 in which the fuel of the aircraft 10 is stored. The calculated instantaneous mass Mc. of the aircraft 10 can be determined from the amount of fuel remaining in the aircraft 10 or the amount of fuel consumed from an initial time to such as the takeoff of the aircraft 10.
    [0033] 3036789 34 It is then possible to compare the measured instantaneous mass Mm and the calculated instantaneous mass Mc by arithmetic during the flight of the aircraft 10. by determining a first difference between this measured instantaneous mass Mm and this calculated instantaneous mass M. The measured instantaneous mass Mm is then considered as an instantaneous mass estimated M that is reliable and usable when this first difference is less than or equal to an error threshold predetermined. In addition, this first deviation can make it possible to identify a change in the payload of the aircraft 10 when its absolute value is large. This change may correspond to unloading or loading at least a portion of this payload. The values of the measured instantaneous mass Mm and the calculated instantaneous mass Mc can also be compared statistically. during a flight of the aircraft 10 to determine an estimated instant mass M consolidated, reliable and accurate. For example, it is possible to calculate a first average of these first deviations determined during the flight of the aircraft since the initial instant te. Then, we can compare this first average with the error threshold as previously mentioned or analyze the first standard deviation of this first average. It is thus possible to compare the evolution of the estimated instantaneous mass M of the aircraft 10 by virtue of the different values of the measured instantaneous mass Mm and of the instantaneous mass calculated Mc during a flight by statistical analysis. Such a statistical analysis can be performed by means of a Kalman filter 31. A diagram of such an architecture is shown in FIG. 8. It is thus possible to consolidate the estimated instantaneous mass M by comparing the values of the mass. instantaneous measured Mm and instantaneous mass calculated Mc. for example by the mass of fuel Mcc consumed by the aircraft 10. M 5 First determines a measurement vector Zm (mmcc), the measured instantaneous mass Mm and the mass of fuel consumed Mcc being determined by a computer 32. As previously described, the measured instantaneous mass Mm is determined from the performance curves 10 of the aircraft 10 and the mass of fuel consumed Mcc is determined from the fuel consumption, more precisely from the instantaneous mass flow Dm; from this fuel for an initial moment to. The calculator 32 and the Kalman filter 31 are preferably integrated in the calculation means 2.
    [0034] A state vector X (M) to be determined is then defined, Mo M being the estimated instantaneous mass of the aircraft 10 and Mo being its initial total mass at the initial moment t0. The total initial mass Mo is therefore a constant and its derivative is zero. Furthermore, the estimated instantaneous mass M is equal to the measured instantaneous mass Mm and the mass of fuel consumed Mcc is the difference between the initial total mass Mo and the estimated instantaneous mass M while taking into account errors or inaccuracies in measurements. , such as M = Mm Arne, and Mcc = Mo - M Ames- It can also be written that the estimated instantaneous mass M is equal to the difference between the total initial mass Mo and the mass of fuel consumed Mcc to the errors or inaccuracies of Close measurements, such as M = M0 -Mcc + Arne ,. In fact, the derivative 3036789 36 of this estimated instantaneous mass M is equal to the derivative of the consumed fuel mass Mcc which is the instantaneous mass flow rate Dm of the fuel, the initial total mass Mo being a constant. So we have M = -Dm.
    [0035] We then define a state equation = AX + B.Dmi and a measurement equation Zm = CX with the following matrices A. [00 00], - 1 1 01 B = [and C = 1, being the derivative of state vector 0. Lastly, the state equation and the measurement equation 10 are applied to the Kalman filter 31 in order to determine the state vector X. In this way, values of the estimated instantaneous mass M are obtained. of the total initial mass Mo at each instant t during the flight of the aircraft 10. Moreover, since the matrix A is a null matrix, the realization of the Kalman filter 31 can be simplified with respect to the generic representation of the FIG. 8 deleting the link containing this matrix A. In addition, a state noise Brétat and a noise related to the Brmes measurements are to be taken into account in these equations which are written as follows: = A. X + B. Dm + Brétat and Zm = C. X + Brmes. The value of this state noise Brétat is for example defined according to the confidence and / or the representativity of this state equation while the value of the noise related to the Brmes measurements is defined according to the margin of precision related to the measurement performance of the measurements. 25 sensors themselves. Consequently, if the difference between the measured instantaneous mass Mn 'and the estimated instantaneous mass M is less than or equal to an error margin corresponding to this noise related to the Brmes measurements and to this state noise Brétat, this mass 3036789 37 Estimated instant M is considered reliable and usable. On the other hand, if the difference between the measured instantaneous mass Min and the estimated instantaneous mass M is greater than this margin of error, this estimated instantaneous mass M is considered as unreliable and not usable. Moreover, the initial total mass Mo is normally a constant which corresponds to the total mass of the aircraft 10 at the initial moment to. Advantageously, the method according to the invention then makes it possible to identify variations of this total initial mass Mo corresponding in particular to a loading or to an unloading of a payload of the aircraft. The graph of FIG. 9 represents the variation of X (values of the total initial mass Mo defined by the state vector 111) as a function of time t. There is a discrete decrease Mo 15 of this total initial mass Mo at time tA which corresponds in fact to a measurement or calculation error, such that the use of a series of performance curves does not correspond to in the case of theft of the aircraft 10. There is also a sudden and lasting decrease in the total initial mass Mo from 20 instant tB. This decrease in the total initial mass Mo being durable and substantially constant after the instant tB in fact makes it possible to identify an unloading of a payload of the aircraft 10. 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. 30
    权利要求:
    Claims (20)
    [0001]
    REVENDICATIONS1. Method for estimating the instantaneous mass of a rotary wing aircraft (10), said aircraft (10) comprising a powerplant (20) provided with at least one engine (21,22) and a gearbox main power unit (23), said main power transmission gearbox (23) rotating at least one main rotor (11) and one anti-torque rotor (13), characterized in that, during a flight, -on measuring the flight characteristics of said aircraft (10), such as the horizontal speed Vh, the vertical speed Vz of said aircraft (10) and the height Hz of said aircraft (10) relative to the ground, -on measurement of the power characteristics of said aircraft (10), such as the torque CR and the actual rotational speed NR of said main rotor (11) and the torque CRAC and the actual rotational speed NRAc of said anti-torque rotor (13), measure atmospheric characteristics relative to the environment of said aircraft (10), such as the pressure at mospheric Po and the temperature To of the air around said aircraft (10), -on determines a flight power W of said aircraft (10), -on determines a point of operation of said aircraft (10) on at least one series of curves the performance of said aircraft (10) according to said flight characteristics of said aircraft (10), said atmospheric characteristics and said flight power Wdudit aircraft (10), and - we deduce a measured instantaneous mass Mm said aircraft (10).
    [0002]
    2. Method according to claim 1, characterized in that a selection algorithm is used to determine each series of performance curves of said aircraft (10) to be used according to the values of said horizontal speed Vh and said vertical speed. Vz of said aircraft (10). 5
    [0003]
    3. Method according to any one of claims 1 to 2, characterized in that, when said aircraft (10) is hovering, said horizontal speed Vh and said vertical speed Vz being substantially zero, functional characteristics of said aircraft (10). ) being defined in particular by a series of first 10 performance curves according to a first formula. (NR0) 3 = ice1 rm W being said flight power of said cr NR J 1 aircraft (10), where o is a reduction coefficient, k being a coefficient of influence of the ground on the behavior of said aircraft (10) as a function of said height Hz of said aircraft (10) with respect to said ground, Mn being said instantaneous measured mass of said aircraft (10), NR0 being a speed rotation setpoint of said main rotor (11), NR being said actual rotation speed of said main rotor (11) and fi being a first function represented by a series of first performance curves of said main rotor (11) According to said aircraft (10), said reduction coefficient is calculated such that u = (e), a first value A / is calculated such that A1 o- "U112) = IN (N120) 3 - is determined, by virtue of a first performance curve of said aircraft (10) according to said first function fi corresponding to the flight conditions of said aircraft (10) and as a function of said first value A1, a second value A2 2 M771 (NR " 2). Mm cr NR such that A2 = f1H 0- NR k - said coefficient of influence k is determined according to said height Hz of said aircraft (10) with respect to said ground, and 3036789 40 - from said second value Az said mass is calculated instantaneous measured Mm of said aircraft (10), such that NR) 2 Min = k.A2.a (.-. NRo
    [0004]
    4. Method according to any one of claims 1 to 2, characterized in that, when said aircraft (10) is in ascending flight, said horizontal speed Vh and said vertical speed Vz being non-zero, functional characteristics of said aircraft ( 10) being defined in particular by two series of seconds and third performance curves respectively according to a second formula (Ij I w, = f2 (Vz) and a I (- O-) Vy third formula (141 = f3Cel), W being said flight power 0- Vy of said aircraft (10), u being a reduction coefficient, Mm being said instantaneous measured mass of said aircraft (10), Vz being said vertical speed of said aircraft (10), Vy being the optimum speed of said aircraft (10), f2 and f3 being respectively a second and a third function represented by two series of performance curves of said aircraft (10), a first ratio (11a) being obtained for a vertical speed Vz that lwith said aircraft (10), said flight power W corresponding to said vertical speed Vz, a second ratio () being obtained cf Vy for a speed Vh 'of said aircraft (10) only horizontal and equal to said optimum climb speed Vy, an optimum power W 'corresponding to said horizontal speed Vh', - is determined by a second performance curve 25 of said aircraft (10) according to said second function f2 corresponding to the flight conditions of said aircraft (10) and 3036789 41 according to said vertical speed Vz of said aircraft (10) a (W 'third value A3 such that A3 = f2 (17Z) = 1: ÷ 7) / (w, a) 1Ty - said reduction coefficient o- is calculated a = T is calculated from said third value A3, said flight power W and said reduction coefficient o- a (If) / fourth value A4 such that A4 = e A = r) 3 Cr Vy-on determines through a third performance curve of said aircraft (10) according to the adite third function f3 and corresponding to the flight conditions of said aircraft (10) and in function of said fourth value A4 which is such that 3 (1 A4 = [r) 1 = 1. r) 'e Vy a fifth value AS such that A5 = (-fn), and - from said fifth value As is calculated said measured instantaneous mass Mm such that Mm = As.a . 15
    [0005]
    5. Method according to any one of claims 1 to 2, characterized in that, when said aircraft (10) is in level flight, said vertical speed Vz is substantially zero, the functional characteristics of said aircraft (10) being defined in particular by two series of third and fourth performance curves according to a third formula (-) = f3 (-) and a fourth Vy formula = [(Vh) 3 Min], W being said formula J411 ify), flight power said aircraft (10) being a reduction coefficient, Mm being said instantaneous measured mass of said aircraft (10), Vh being said horizontal speed of said aircraft (10), Vy being the optimum climb speed of said aircraft (10), f3 and fi being respectively third and fourth functions represented by two series of performance curves of said aircraft (10), a first ratio (f) being obtained for any horizontal speed Vh of said aircraft, said flight power W corresponding to said horizontal speed Vh, a second ratio (-1f) being obtained for a speed Vh 'of said aircraft (10) Vy only horizontal and equal to said optimum climb speed Vy, an optimum power W' corresponding to said horizontal velocity Vh ', said measured instantaneous mass Mm is determined which makes it possible to simultaneously solve the third and fourth formulas, according to said flight power W and said horizontal speed Vh. 15
    [0006]
    6. Method according to claim 5, characterized in that a sixth value A6 is calculated such that A6 = (Vh) 3 as a function of said horizontal speed Vh and of said optimum climb speed Vy of said aircraft (10). a series of fourth performance curves of said aircraft (10) is determined according to said fourth function f 1 and as a function of said sixth value A 6 of the first pairs formed of a third value A3 and a fifth (17) / mm A5 value such that A3 = cr (cy)) and A5 = -, vy 25-, said reduction coefficient cr is calculated such that Ci = (- po), To be calculated from each third value A3, of said flight power W and said reduction coefficient u a (17) / fourth value A4 such that A4 = a A = [(117-) Vy1, in order to form second pairs formed of a fourth value A4 5 and of a fifth value A5, the second pair formed of a fourth value A4 and a fifth I value As solving said third formula through a third performance curve of said aircraft (10) according to said third function f3 and corresponding to the flight conditions of said aircraft (10), and - one calculates from said fifth value As said second couple solving said third formula said measured instantaneous mass Mn, such that Mm = A5.cr.
    [0007]
    7. Method according to any one of claims 1 to 6, characterized in that - a calculated instantaneous mass / 1/1 is calculated, which is determined from the variation of the quantity of fuel present in said aircraft (10 ), said measured instantaneous mass Mn is compared with said calculated instantaneous mass Mc. to obtain an estimated instantaneous mass M of said reliable and consolidated aircraft (10).
    [0008]
    8. Method according to claim 7, characterized in that, in order to compare the values of said measured instantaneous mass Xi and said calculated instantaneous mass Mc during a flight of said aircraft (10) arithmetically, 3036789 44 - a first difference between said measured instantaneous mass Mm and said calculated instantaneous mass Mc is determined; if said first difference is less than or equal to an error threshold, said measured instantaneous mass Mm is considered as an instantaneous mass estimated M reliable and usable and, -if said first deviation is greater than said error threshold, said measured instantaneous mass Mm is considered as an instantaneous mass estimated M not usable and an alert is issued. 10
    [0009]
    9. Method according to claim 7, characterized in that the values of said measured instantaneous mass Mm and of said calculated instantaneous mass Mc during a flight of said aircraft (10) are compared statistically.
    [0010]
    10. Method according to claim 9, characterized in that a first difference between said measured instantaneous mass Mm and said calculated instantaneous mass Mc is determined, a first average of said first determined discrepancies between a time t and an initial instant being calculated. if said first average is less than or equal to an error threshold, said measured instantaneous mass Mm is considered to be an estimated instantaneous mass M that is reliable and usable and, if said first average is greater than said error threshold, Said measured instantaneous mass Mm is considered as an instantaneous mass estimated M that can not be used and an alarm is emitted.
    [0011]
    11. The method as claimed in claim 9, wherein a first difference between said measured instantaneous mass Mm and said calculated instantaneous mass Mc is determined, a first average of said first determined discrepancies between a time t and an initial time to, and a first associated standard deviation, -on analyzes said first average and said first associated standard deviation to determine if said measured instantaneous mass Mm is considered as an instantaneous mass estimated M reliable and usable.
    [0012]
    12. The method of claim 9, characterized in that the values of said measured instantaneous mass Mm and said calculated instant mass Mc are compared. during a flight of said aircraft (10) via at least one Kalman filter (31).
    [0013]
    13. A method according to any one of claims 7 to 12, characterized in that one measures an instantaneous volumetric flow rate Dvi of 20 consumption of a fuel supplying said power plant (20), -on measures the instantaneous temperature T, said fuel, -on calculates an instantaneous mass flow Dm, said fuel, -on determines a Mcc mass of said fuel consumed from an initial moment to by integration of said instantaneous mass flow Dm, said fuel from said initial moment to, and -on calculates said fuel instantaneous mass calculated Mc. which is equal to the total initial mass Mo of said aircraft (10) at said initial time to which said mass Mcc of said consumed fuel is subtracted.
    [0014]
    14. A method according to any one of claims 7 to 12, characterized in that, said aircraft (10) comprising at least one reservoir (25,26) in which said fuel is stored, -one measure Vo volume said fuel remaining in all of said tanks (25,26), the instantaneous temperature Ti of said fuel is measured, a mass McR of said fuel remaining in said set of said tanks (25,26) is calculated; a mass Mcc of said fuel consumed from an initial time to subtracting said McRto mass from said fuel remaining in said set of said tanks (25,26) to said initial moment of said mass McRt of said fuel remaining at time t, and calculating said calculated instantaneous mass Mc which is equal to the total initial mass Mo of said aircraft (10) at said initial moment ta to which said mass Mcc of said consumed fuel is subtracted.
    [0015]
    15. Method according to any one of claims 7 to 12, characterized in that, said aircraft (10) comprising at least one reservoir (25,26) in which said fuel is stored, -on measures a VcR volume of said fuel remaining in all of said tanks (25,26), - the instantaneous temperature 7- of said fuel is measured, a mass Mo is calculated said fuel remaining in said set of said tanks (25,26), calculates said calculated instant mass Mc. which is equal to the initial mass M / off fuel of said aircraft (10) to said initial instant 5 to which is added said mass McR of said remaining fuel.
    [0016]
    16. The method of claim 13, characterized in that for comparing the values of said measured instantaneous mass Min and said calculated instant mass Mc 10. during a flight of said aircraft (10) statistically via a Kalman filter (31), M is determined a measurement vector Zn, such that Zm (mmcc), Mn, being said instantaneous mass measured and Mcc being said consumed fuel mass of said aircraft (10), -on defines a state vector X to be determined such that X (mo), M being an estimated instantaneous mass and Mo being the initial total mass of said aircraft ( 10), one defines a state equation i = A.X + B.Dmi, with A = FO OF LO 0.1 and B = Dm i being said instantaneous mass flow rate of said fuel equal to the derivative of said fuel mass consumed Mcc of said aircraft (10), being the derivative of said state vector X, 1 OF -on defines an equation of measurements Zm = CX, with C = r - applying said state equation and said measurement equation 25 to a Kalman filter (31) for determining said state vector X, and hence said estimated instant mass M and said mass in itiale total Mo said aircraft (10). 3036789 48
    [0017]
    17. Method according to any one of claims 1 to 16, characterized in that said flight power W is equal to the sum of the power consumed by said main rotor (11) and 5 of the power consumed by said anti-torque rotor ( 13).
    [0018]
    18. A method according to any one of claims 1 to 16, characterized in that said flight power W is equal to a motive power supplied by said power plant (20) to which is subtracted an accessory power Wu. supplying equipment of said aircraft (10).
    [0019]
    19. Device (1) for estimating the instantaneous mass of a rotary wing aircraft (10), said aircraft (10) comprising a powerplant (20) provided with at least one engine (21,22) and a main power transmission gearbox (23), as well as a plurality of sensors (13-19) providing measurements on the operation and environment of said aircraft (10), said main power transmission gearbox (23); ) rotating at least one main rotor (11) and one anti-torque rotor (13), characterized in that said instantaneous mass estimation device (1) comprises at least one calculating means (2) and at least one a memory (3), a memory (3) storing series of performance curves of said aircraft (10) and calculation instructions, said calculating means (2) receiving said measurements of said sensors (13-19) and applying said calculation instructions in order to implement the method of estimating the mass ins of a rotary wing aircraft (10) according to any one of claims 1 to 18.
    [0020]
    20. Device (1) according to claim 19, characterized in that said device for estimating the instantaneous mass (1) comprises at least one Kalman filter (31) in order to compare the evolution of said instantaneous measured mass. Mm and said instant mass calculated Mc. said aircraft (10) during a flight and to determine an instantaneous estimated mass M.
    类似技术:
    公开号 | 公开日 | 专利标题
    EP3098579B1|2019-07-03|A method of estimating the instantaneous mass of a rotary wing aircraft
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    同族专利:
    公开号 | 公开日
    FR3036789B1|2017-05-26|
    US20170010148A1|2017-01-12|
    EP3098579B1|2019-07-03|
    EP3098579A1|2016-11-30|
    US10545047B2|2020-01-28|
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    法律状态:
    2016-05-20| PLFP| Fee payment|Year of fee payment: 2 |
    2016-12-02| PLSC| Publication of the preliminary search report|Effective date: 20161202 |
    2017-05-23| PLFP| Fee payment|Year of fee payment: 3 |
    2018-05-22| PLFP| Fee payment|Year of fee payment: 4 |
    2019-05-22| PLFP| Fee payment|Year of fee payment: 5 |
    2020-05-22| PLFP| Fee payment|Year of fee payment: 6 |
    2022-02-11| ST| Notification of lapse|Effective date: 20220105 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1501117A|FR3036789B1|2015-05-29|2015-05-29|METHOD OF ESTIMATING THE INSTANTANEOUS MASS OF A ROTATING WING AIRCRAFT|FR1501117A| FR3036789B1|2015-05-29|2015-05-29|METHOD OF ESTIMATING THE INSTANTANEOUS MASS OF A ROTATING WING AIRCRAFT|
    EP16168032.7A| EP3098579B1|2015-05-29|2016-05-03|A method of estimating the instantaneous mass of a rotary wing aircraft|
    US15/161,620| US10545047B2|2015-05-29|2016-05-23|Method of estimating the instantaneous mass of a rotary wing aircraft|
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