Method and assistance system for detecting a degradation of flight performance

Abstract

The invention relates to a method and to a device for detecting a degradation of flight performance of an aircraft that is in flight, wherein current flight status data of the aircraft that is in flight are first determined. A flight performance index is then calculated on the basis thereof. Furthermore, on the basis thereof, a nominal flight performance reference index is determined by means of a flight performance model, wherein a degradation of flight performance can be inferred by comparing the two indices.

Claims

1. A method for detecting a degradation of flight performance of an aircraft in flight, comprising steps of: a) determining a current flying state of the aircraft in flight, from a plurality of values comprising flying state parameters that influence or characterize flight performance of the aircraft, which are at least partially recorded by sensors provided on the aircraft, calculating a current flight performance indicator from one or more flying state parameters of the currently determined flying state by an electronic evaluation unit, b) calculating, by an electronic evaluation unit, a current flight performance indicator the calculating being configured to calculate the current flight performance indicator as an overall change in energy over a predetermined period of time, based at least in part on one or more flying state parameters of the determined current flying, c) determining, by the electronic evaluation unit, a nominal flying performance reference indicator from a flight performance model of the aircraft in flight and one or more flying state parameters of the currently determined flying state, wherein the flight performance model of the aircraft in flight replicates a nominal flight performance of the aircraft in a non-degraded flying state, and wherein the flight performance model is provided to the electronic evaluation unit, and d) comparing, by the electronic evaluation unit, the current flight performance indicator and the nominal flight performance reference indicator and, if the comparing indicates a deviation between the current flight performance indicator and the nominal flight performance reference indicator that is greater than a predetermined limit value, detecting a degradation of flight performance of the aircraft in flight.

2. The method as claimed in claim 1, wherein the step of detecting the degradation of flight performance is configured to detect that the deviation establishes aircraft icing.

3. The method as claimed in claim 1 wherein the current flying state parameters are selected from the group consisting of a current flying speed, a change over time of the current flying speed, a current altitude, a change over time of the current flying altitude, a current overall aircraft mass, a change over time of the overall aircraft mass, an engine characteristic for determining engine performance or engine thrust, a load factor, a current lift characteristic of the aircraft, a current dynamic pressure, and a current aircraft configuration.

4. The method as claimed in claim 1 wherein the current lift A is calculated as the current lift characteristic of the aircraft as a flying state parameter according to the formula
A≡(n.sub.z).sup.a.Math.g.Math.m where A is the lift, (n.sub.z).sup.a is a load factor in the lift axis of the aircraft, g is the acceleration due to gravity and m is the overall aircraft mass.

5. The method as claimed in claim 4, wherein a load factor (n.sub.z).sup.a in a lift axis of the aircraft is calculated according to the formula
(n.sub.z).sup.a≡−(n.sub.x).sup.f.Math.sin(α)+(n.sub.z).sup.f.Math.cos(α) where (n.sub.x).sup.f is the load factor in the longitudinal axis of the aircraft, (n.sub.z).sup.f is the load factor in the vertical axis of the aircraft and a is the angle of attack.

6. The method as claimed in claim 1 further comprising: determining an engine characteristic, wherein the engine characteristic comprises a low-pressure shaft speed of at least one engine of the aircraft, wherein the engine characteristic is among the flying state parameters.

7. The method as claimed in claim 1, wherein calculating the current flight performance indicator includes calculating the overall change in energy according to the formula
Ė.sub.ov=m.Math.V.sub.TAS.Math.{dot over (V)}.sub.TAS+½.Math.{dot over (m)}.Math.V.sub.TAS.sup.2+m.Math.g.Math.{dot over (H)}+{dot over (m)}.Math.g.Math.H where (Ė.sub.ov) is the overall change in energy over the predetermined time period, H is the current flying altitude, ({dot over (H)}) is the change over time of the flying altitudes over the predetermined time period, V.sub.TAS is the current flying speed with respect to the air flowing around, {dot over (V)}.sub.TAS is the change over time of the flying speed with respect to the air flowing around over the predetermined time period, m is the current overall aircraft mass, ({dot over (m)}) is the change over time of the overall aircraft mass over the predetermined time period and g is the acceleration due to gravity.

8. The method as claimed in claim 1 wherein the nominal flight performance reference indicator is a nominal overall change in energy over a predetermined time period, a flight performance model being provided, from which the nominal overall change in energy for the current flying state is derivable on the basis of the current flying state.

9. The method as claimed in claim 8, further comprising calculating a differential resistance coefficient ΔC.sub.W based on a current overall change in energy as a current flight performance indicator and a nominal overall change in energy as a nominal flight performance reference indicator according to the formula $\Delta C_W=\frac{{\stackrel{.}{E}}_{\mathrm{ref}}-{\stackrel{.}{E}}_{\mathrm{ov}}}{q.Math.S.Math.V_{\mathrm{TAS}}}$ where (Ė.sub.ov) is detected as the current overall change in energy, (Ė.sub.ref) as the nominal overall change in energy, V.sub.TAS as the speed of the aircraft with respect to the air, q as the dynamic pressure and S as the reference surface area of the aircraft, and detecting a degradation of the flight performance if the differential resistance coefficient ΔC.sub.W is greater than the limit value.

10. The method as claimed in claim 1 wherein the current flight performance indicator is determined while also taking into account a variation in wind experienced by the aircraft.

11. The method as claimed in claim 10, wherein a wind component of a change over time of the speed of the aircraft with respect to the air while taking into account the variation in wind experienced according to the formula ${\stackrel{.}{V}}_{\mathrm{TAS},{\stackrel{.}{\stackrel{.fwdarw.}{V}}}_k}=\frac{{\stackrel{.}{u}}_{k,g}{u}_{a,g}+{\stackrel{.}{v}}_{k,g}{v}_{a,g}+{\stackrel{.}{w}}_{k,g}{w}_{a,g}}{V_{\mathrm{TAS}}}$ where {dot over (u)}.sub.k,g, {dot over (v)}.sub.k,g, {dot over (w)}.sub.k,g are the three components of the path acceleration vector (time derivative of the path speeds) in the geodetic system of coordinates, u.sub.a,g, v.sub.a,g, w.sub.a,g are the three components of the vector of the current flying speed with respect to the air, V.sub.TAS is the speed of the aircraft with respect to the air and V.sub.TAS,{right arrow over ({dot over (V)})}.sub.K is the change in the current flying speed because of a change in the path speed, and the current flight performance indicator is given as an overall change in energy over a predetermined time period and is calculated according to the formula
Ė.sub.ov=m.Math.V.sub.TAS.Math.{dot over (V)}.sub.TAS+½.Math.{dot over (m)}.Math.V.sub.TAS.sup.2+m.Math.g.Math.{dot over (H)}+{dot over (m)}.Math.g.Math.H where {dot over (V)}.sub.TAS={dot over (V)}.sub.TAS,{right arrow over ({dot over (V)})}.sub.K.

12. The method as claimed in claim 1 wherein values of flying state parameters for determining a sideslip state are determined, a compensation value of the current flight performance indicator is calculated in dependence on these determined values of flying state parameters for determining the sideslip state and the comparison is carried out in dependence on the compensation value.

13. The method as claimed in claim 12, wherein the compensation value is a resistance compensation value, which is calculated according to the formula $\Delta C_{W\beta ,\mathrm{Comp}}\equiv \frac{n_y.Math.m.Math.g.Math.\sin \hat{\beta }}{q.Math.S_F}$ where ΔC.sub.W,comp is the resistance compensation value, n.sub.y is a lateral load factor, m is the overall aircraft mass, g is the acceleration due to gravity, β is the sideslip angle, q is the dynamic pressure and SF is the reference surface area of the aircraft.

14. The method as claimed in claim 1 wherein a reduced flight performance model is provided to the electronic evaluation unit in the form of a multidimensional table, each flying state parameter that is relevant to the flight performance model being replicated by a dimension of the table, each dimension of the table having a plurality of interpolation points, which are the predetermined values of the respective flying state parameters, and at least one nominal flight performance reference indicator being stored for each pair of interpolation values comprising values of the various flying state parameters.

15. The method as claimed in claim 14, wherein the flying state parameters forming the dimensions of the table include any among and/or any combination or sub-combination of flying speed with respect to the air flowing around, a lift characteristic of lift of the aircraft, an engine characteristic of engine performance, flying altitude, and an aircraft configuration.

16. An assistance system for detecting a degradation of flight performance of an aircraft in flight, the assistance system being configured for performing the method for detecting the degradation in flight performance as claimed in claim 1.

17. The assistance system as claimed in claim 16, wherein the assistance system is configured for performing a detecting of aircraft icing based on a recognized degradation of flight performance.

18. An aircraft comprising an assistance system as claimed in claim 16.

Description

(1) The invention is explained by way of example on the basis of the accompanying figures, in which:

(2) FIG. 1 shows a schematic representation of an assistance system;

(3) FIG. 2 shows a simplified representation of a tabular flight performance model.

(4) FIG. 1 schematically shows the assistance system 10, which may for example be an electronic data processing system within the avionic systems of an aircraft. However, it is also conceivable that the assistance system 10 is provided outside the aircraft, a communicative connection between the assistance system on the one hand and the aircraft on the other hand then having to exist in order to be able to transmit the data of the flying state that are necessary for the calculation and the detection to the assistance system 10 and to transmit a possible detection of icing back again into the aircraft. Use as “post-flight analysis” is also conceivable.

(5) In the further explanations, it is however assumed that the assistance system 10 is a component part of an aircraft.

(6) The assistance system 10 is connected via an interface 11 to the data bus 12 of the avionic system of the aircraft, in order to be able to record the flying state parameters necessary for detection. By way of this data bus 12, the assistance system 10 is indirectly in connection with the sensors fitted as standard in the aircraft and can thus record the flying state parameters necessary for detection that are measured with the aid of the sensors during the flight and access them from the data bus 12 via the interface 11.

(7) The assistance system 10 has furthermore a digital data memory 13, in which the flight performance model 14 is stored in the form of a multidimensional table. The multidimensional table has the advantage, however, that the determination of the flight performance reference indicator is possible without any special computing effort, since it is obtained directly from the table in dependence on the specific values of the flying state parameters of the flying state. In addition, there is the possibility of interpolating between flight performance reference indicators if the values of the flying state parameters of the flying state do not directly replicate the corresponding interpolation points.

(8) As shown in FIG. 2, such a tabular representation of the digital flight performance model may for example consist of five dimensions, a dimension being respectively provided for the flying speed, a lift characteristic, an engine characteristic, the flying altitude, and also an overarching aircraft configuration.

(9) Returning to FIG. 1, the assistance system 10 has furthermore an electronic evaluation unit 15, which is designed for detecting an icing state. For this, the evaluation unit 15 has a reference module 16, which is designed for calculating or determining a nominal flight performance reference indicator. In the exemplary embodiment of FIG. 1, the nominal flight performance reference indicator is the overall change in energy Ė.sub.ref, which is also stored in the table in the flight performance model according to FIG. 2.

(10) The reference module 16 is in this case connected to the digital data memory 13, in which the flight performance model 14 is stored, in order to be able to access the table 14 stored there. Furthermore, the reference module 16 is connected in terms of signaling to the data bus 12 via the interface 11, in order to be able to determine the flying state parameters necessary for the calculation and determination of the overall change in energy Ė.sub.ref and their current values.

(11) In the exemplary embodiment of FIG. 1, the reference module 16 requires at least the current flying speed V.sub.TAS with respect to the air flowing around, the current barometric height H, the current overall mass m, values with respect to the load factors (at least in the lift axis, advantageously also in the longitudinal and transverse axes), an engine characteristic and also items of information with respect to the aircraft configuration. The engine characteristic may be for example the low-pressure shaft speed of the engines of the aircraft (for example as an averaged characteristic over all the engines). If the aircraft has two or more engines, it is conceivable that the low-pressure shaft speed is averaged over the engines.

(12) With respect to the aircraft configuration, items of information that reflect the current aircraft configuration are transmitted. These are items of information with respect to the landing gear (retracted, extended) and also items of information with respect to lift-changing measures, such as for example high-lift systems, slats, landing flaps. These items of information are therefore advantageous because, by changing the aircraft configuration in such a way, the aerodynamics of the aircraft are influenced, and consequently the overall resistance of the aircraft is changed. In order to prevent that, when there is a change of the aircraft configuration, and consequently an accompanying change of the overall resistance, a variation of the flight performance caused by icing is not inferred, the individual possible aircraft configurations are also taken into account in the flight performance model, so that a correct nominal flight performance reference indicator can also be determined for each aircraft configuration.

(13) For the determination of the nominal overall change in energy Ė.sub.ref, first a lift characteristic A is required, obtained according to the formula
A≡(n.sub.z).sup.a.Math.g.Math.m

(14) The load factor (n.sub.z).sup.a can in this case be calculated according to the formula)
(n.sub.z).sup.a≡−(n.sub.x).sup.f.Math.sin(α)+(n.sub.z).sup.f.Math.cos(α)
where the angle α is the angle of attack.

(15) On the basis of the current aircraft configuration, the multidimensional table of the flight performance model 14 that matches the current aircraft configuration is then determined from the digital data memory. Then, on the basis of the values of the current flying altitude, the lift characteristic, the flying speed and also the engine characteristic from the table, the nominal overall change in energy Ė.sub.ref over time is determined and temporarily stored in the reference module 16.

(16) The nominal overall change in energy Ė.sub.ref over time is in this case the characteristic of the flight performance of the aircraft in the un-iced state and can to this extent be understood as an idealized value. This reference indicator of the flight performance may in this case either be provided generally for the aircraft type of the aircraft or be adapted specifically to the aircraft, for example if the aircraft is already somewhat older, resulting in a changed flight performance. This allows the overall system to be much more accurate.

(17) Furthermore, the evaluation unit 15 has a flying state module 17, which is likewise connected by the interface 11 to the data bus 12 and can then calculate on the basis of corresponding flying state parameters a current flight performance indicator in the form of an overall change in energy Ė.sub.ov. For this, the flying state module 17 receives as flying state parameters at least the current flying speed V.sub.TAS, the current altitude H, the overall aircraft mass m and also a change over time of the overall aircraft mass {dot over (m)}.

(18) On the basis of these values, it is then possible with the aid of the formula
{dot over (E)}.sub.ov≈(g.Math.{dot over (H)}.Math.m)+(g.Math.H.Math.{dot over (m)})+(V.sub.TAS.Math.{dot over (V)}.sub.TAS.Math.m)+(½.Math.V.sub.TAS.sup.2.Math.{dot over (m)})
to determine the current overall change in energy over time as a flight performance characteristic.

(19) In order to avoid atmospherically induced instances of false detection, a quasi steady state of the atmosphere is advantageously assumed. The change over time of the norm of the incoming-flow velocity vector V.sub.TAS may comprise not only a component from the change in the path speed (aircraft movement in the quasi steady-state, homogeneous wind field) but in addition a component from the experienced change over time of the wind (flying through a steady-state and/or inhomogeneous wind field), with
{dot over (V)}.sub.TAS={dot over (V)}.sub.TAS,{right arrow over ({dot over (V)})}.sub.K+{dot over (V)}.sub.TAS,{right arrow over ({dot over (V)})}.sub.w

(20) The component {dot over (V)}.sub.TAS,{right arrow over ({dot over (V)})}.sub.K is dominated here decisively by the characteristics of the aircraft, and consequently correspondingly comprises the flight performance. By contrast, the change in the incoming flow due to a change of the wind {dot over (V)}.sub.TAS,{right arrow over ({dot over (V)})}.sub.w results from the non-steady atmosphere through which the aircraft is moving. With the estimated component of the wind, in for example a geodetic system of coordinates (as advantageously explained below), these two components can be analytically separated. The following applies:

(21) ${\left({\stackrel{.fwdarw.}{V}}_{\mathrm{TAS}}\right)}_g={\left({\stackrel{.fwdarw.}{V}}_k\right)}_g-{\left({\stackrel{.fwdarw.}{V}}_w\right)}_g.Math.{\left({\stackrel{.}{\stackrel{.fwdarw.}{V}}}_{\mathrm{TAS}}\right)}_g=\frac{d}{\mathrm{dt}}\left({\left({\stackrel{.fwdarw.}{V}}_k\right)}_g-{\left({\stackrel{.fwdarw.}{V}}_w\right)}_g\right)$