Note: Descriptions are shown in the official language in which they were submitted.
CA 02510115 2005-06-16
Device and method for damning at least one of a rigid body mode and elastic
mode of an aircraft
Field of the invention
The present invention relates to the modification of a rigid body mode or an
elastic
mode of an element. In particular, the present invention relates to a damping
device
for damping at least one of at least one rigid body mode and at least one
elastic mode
of an aircraft and to a method of damping at least one of at least one rigid
body mode
and at least one elastic mode of an aircraft.
Technological background
A device having a sensor, a regulation or control device connected to the
sensor and
an actuator for driving controlling, guiding and/or regulating surfaces of the
aircraft is
known.
Summary of the invention
The invention further concerns a method for damping at least one rigid body
mode
and/or at least one elastic mode of an airplane, especially for blast load
reduction and
increase of stability and comfort in an airplane having at least one sensor
means as
well as having a regulation unit connected with the one or the several sensor
means
and at least one actuator acting upon controlling, guiding and/or regulating
surfaces of
the aircraft.
There may be a need for a lowering of arbitrary local and global structural
loads in
aircrafts, which are induced e.g. by blast induced and/or manoeuvering induced
excitation of rigid body modes and/or elastic modes.
The used short term "mode" concerns a "rigid body mode" and/or an elastic mode
of
the aircraft.
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2
Moreover, the device according to an exemplary embodiment of the invention or
the
method according to an exemplary embodiment of the invention, respectively,
may
serve for increasing comfort as well as increase stability of the aircraft by
means of a
damping of the respective "rigid body modes" and/or elastic eigenmovements.
From the state of the art, a plurality of regulation systems for decreasing
rigid body
modes and/or elastic modes, generally also called modes, in aircrafts are
known.
Theoretically optimal turns out to be, for example, the design of a regulation
system
having a leading back of state vector which admittedly until now failed due to
the
necessity of an "exact observer" who implies a very high computational effort.
Alternatively, the eigenmovements or modes ascribed to the airplane, according
to a
back transform from the modal space into the physical space, as far as
measurement
technology is concerned, may be determined by means of a linear combination of
different acceleration values. In order to come along with a reduced number of
sensors in this case though, a filtering of the signal, mostly by means of
filters of low
order, is necessary for band limiting.
As a further possibility for determining the signal to be lead back from out
of only one
measurement value, for example a laborious filtering is applied in the state
of the art.
In this context, a subtractive method is concerned, wherein from the
measurement
signal, by connecting low pass filters and notch filters in series, all signal
portions
besides the wished mode are filtered out.
Further, it is o$en necessary to adapt the regulation system to the
eigenfrequencies of
the airplane which change, for example, as a consequence of the mass change
due to
the continuous fuel consumption.
With respect to the selection and driving of the controlling, guiding and
regulating
surfaces, up to now in known regulation systems only integral indicators are
considered, as for example the lowering of the effective value of a load or an
acceleration at a certain position.
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Moreover, previous regulation systems mostly concern special solutions for
very
special vibration problems, which are usually problems of comfort. Thereby,
the
applicability of such regulating systems is restricted to the body structure
of an
airplane, and may not offhand be transferred to other components. Further,
until now,
due to the high computational effort, it is only possible, with a justifiable
effort, to
influence a strongly limited number of modes. In this context, for the most
part a rigid
body mode or the first body or wing bending mode is concerned. In order to
create the
necessary leading back signal, there have to be applied several acceleration
sensors,
which not only increases the system's technical effort, but also the
probability of
malfunction.
Although regulating systems which detect a certain mode out of a measurement
signal
by means of simple filtering of a certain mode may be technically realised,
due to the
still too broad-band leading back, in spite of the effected band limiting,
they still
excite other, unwanted eigenmovements of the airplane and thereby influence
the
flight behaviour in a negative way. In turn, therefrom again may result
disadvantages
in the region of the flight behaviour as well as an increase of loads and
accelerations)
at other components of the airplane.
In order to face the shown problems, the modal design of regulation is applied
in
known regulation systems. The modal leading back requires a very great system
complexity though, since in principle for each eigenmovement to be influenced,
an
individual leading back signal has to be created in a computationally involved
way.
This means that from a measurement signal, all modes besides the desired one
have to
be filtered out in a subtractive method by means of notch filters. After a
subsequent
phase correction, the signal may then be switched onto the controlling,
guiding and
regulating surfaces. Now, since each mode requires a different phase
correction, there
has to be built up one such branch for each respective mode accordingly. It is
thereby
not possible, to keep more than one mode within the measurement signal, since
the
phase can only be adapted for one mode.
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It is common to all known regulation systems, that their design is connected
to a high
effort of optimising. The target functions in such optimisations, are in most
cases
effective values of the load or the acceleration at a certain body position of
the
airplane, as a consequence of which the possible excitation of other modes is
not
detected.
In contrary to this, the device according to an exemplary embodiment of the
invention
or the method according to an exemplary embodiment of the invention,
respectively,
are believed to allow the purposeful lowering of individual load peaks in the
load
spectrum or acceleration spectrum of the body structure, not only the lowering
of an
integral value of several modes.
Further, in known regulation systems, the controlling, guiding and regulating
surfaces
are only determined with respect to their global effectiveness in case of
certain load
problems, characterised by effective values. But if the controlling, guiding
and
regulating surfaces are not adapted to the modes in an optimal manner, further
modes
will be influenced even in case of an optimal leading back signal. Up to now,
usually
only one controlling, guiding and regulating surface is selected. This leads
to a very
low efficiency of the known regulation systems, so that greater control
amplitudes and
thereby longer actuator ways become necessary. To the one hand, from this
follows
an increase of the air resistance, and on the other hand an increased actuator
wear.
An exemplary embodiment of the present invention is based on the aim to create
a
device as well as a method for damping rigid body modes and/or elastic modes
of
aircrafts, which avoids the indicated disadvantages of the before mentioned
regulation
systems.
According to an exemplary embodiment, by means of the controller or a
regulator, at
least one respective rigid body mode and/or at least one respective elastic
mode of the
aircraft may be modified, allowing for a reduction of the rigid body modes
and/or
eigenmovements and/or elastic modes of the aircraft, and thereby a substantial
reduction of structural loads.
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According to an exemplary embodiment, the rigid body modes and/or
eigenmovements and/or elastic modes of the aircraft, which may lead to load
problems, comfort problems or stability problems, may be reduced by means of
controlling, guiding and regulating surfaces disposed at the aircraft and
controlled by
means of the controller. The regulation is thereby effected depending on the
measurement values detected by a sensor and processed by means of filter
elements.
The regulation is effective in selected small frequency ranges ofthe rigid
body
eigenmovement and/or the elastic mode of the aircraft. At first, the approach
according to this exemplary embodiment is general and may be applied to
arbitary
local and/or global load problems, stability problems and comfort problems or
a
combination of these problems in aircrafts of any kind.
It is believed that it is advantageous that a device according to an exemplary
embodiment of the present invention or a method according to an exemplary
embodiment of the present invention, respectively, acts frequency selective
and
thereby, as far as the processing of the leading back signal is concerned,
when
modifying a certain mode, all others usually remain un-effected. Thereby
neither the
rigid body eigenmovements to be obtained, nor the elastic eigenmovements to be
obtained are negatively influenced.
The output side of a device according to an exemplary embodiment of the
present
invention or a method according to an exemplary embodiment, respectively, is
based
on a modal criterion which designs the phase position and the amplitude
distribution
of the controlling, guiding and regulating surfaces according to the nature of
the
eigenvector of the relevant mode. Thereby is secured, that even when an
optimal
leading back signal is given, besides the ones) to be influenced, no further
modes are
modified. When the deflections of the controlling, guiding and regulating
surfaces are
optimally adapted with respect to amplitude and phase, a maximum influencing
of the
modes to be modified in the sense of the design is reached with minimal
control
effort. In comparison to known regulation systems or methods, respectively,
the
control effort may be reduced to 10-15%. Thereby, the lifetime of the
actuators is
increased, while at the same time the air resistance is reduced and a thereby
a
reduction of fizeI consumption is caused.
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Moreover, a device according to an exemplary embodiment of the present
invention,
or a method according to an exemplary embodiment of the present invention,
respectively, may only require one (acceleration) sensor, whereupon a high
availability and fail-safety is obtained. At the same time, the application of
an internal
sensor increases the availability, since this cannot be influenced by outer
influences
like for example bird strike etc. Further, it may be possible, that a device
according to
an exemplary embodiment of the present invention participates to use an
already
existing sensor, so that no further additional equipment is necessary.
By means of an exemplary embodiment of the invention, the necessary system
complexity is considerably lowered, because to the one hand, only the
application of
one sensor may be necessary, and to the other hand an optimised, additive
filter
structure is applied.
This reduction of the system complexity in relation to known regulation
systems is
possible by use of resonance filters which at a low order of the system
already provide
the possibility to lead back small frequency intervals from the measurement
signal to
the actuators, and thereby to the controlling, guiding and/or regulating
surfaces.
Thereby, concerningthe signals side, all modes which are not to be modified,
remain
unchanged due to their exclusion from the leading back signal. The reduction
of the
necessary degree of order of the filters is about between 60 and 70 %, as
compared to
known subtractive methods.
Further, there results a great time advantage in the design phase of the
device
according to an exemplary embodiment of the invention, since all parameters
may be
determined analytically, so that extensive optimisations may be omitted.
Thereby, a
saving of time of 80 to 90 % in the design phase results, as compared to other
methods.
According to another exemplary embodiment of the invention, a frequency signal
of a
frequency measurement arrangement may be provided to the controller
representing
the frequency of at least one rigid body mode and/or at least one elastic
mode.
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Thereby, a what is believed to be optimal adaptation of the device according
to the
exemplary embodiment if the invention is given, when physical frame conditions
change, as for example during a loss of mass of the aircra8 as a consequence
of
continuous fuel consumption. Further, the security with respect to modelling
errors
which stem from the design phase is increased. The detection and analysis of
the rigid
body modes as well as the elastic eigenmovements of the aircraft for
adaptation of the
device according to an exemplary embodiment of the invention thereby do not
have to
occur in realtime. Since the mass depending frequencies only change slowly, a
regularly, in medial time intervals occurring verification and, as the case
may be,
anew calibration of the device according to an exemplary embodiment of the
invention is sufficient.
The device according to an exemplary embodiment of the invention or the method
according to an exemplary embodiment of the invention, respectively, may be
distinguished in that a provision for the eigenvector of the to be modified
rigid body
modes and/or elastic modes is effected, when selecting the control surfaces
for
maximising the efficiency, and in that the technical realisation of the device
is
effected by resonance filters for securing a small system complexity.
Short description of the drawings
Further exemplary embodiments of the present invention are depicted in the
following
drawings and will be described in the following.
Figure 1: shows a schematic description of two bending modes of an aircraft.
Figure 2: shows a progression of the lateral force QY normalised to the
maximum
lateral force Qymax, plotted versus the normalised body length of an
aircraft in case of an excitation with lateral blasts.
Figure 3: shows a spectrum of the lateral force QY normalised to the maximum
lateral force QYm$x, plotted versus the frequency f normalised to the
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g
maximum frequency fmaX in a front body region of the aircraft, when in a
dominant mode.
Figure 4: shows an acceleration spectrum in the region of the position of a
sensor
means, when in a dominant mode.
Figure 5: shows a simplified schematic structure of a device according an
exemplary embodiment of the invention.
Figure 6: shows detailed structure of a regulation device for modifying a
mode.
Figure 7: shows progression of the lateral force QY normalised to the maximum
lateral force Qym~, plotted versus the normalised body length of the
airplane without and with application of the device according an
exemplary embodiment of the invention.
Figure 8: shows a spectrum of the lateral force QY normalised to the maximum
lateral force Qyn,aX, plotted versus the frequency f normalised to a
maximum frequency f",aX in the front body region of the airplane without
and with application of the device according to an exemplary
embodiment.
Figure 9: shows an acceleration spectrum nY/ nym$x, plotted versus the
frequency f
normalised to the maximum frequency f",aX in the region of the sensor
means in the front body region of the airplane with and without
application of the device.
Figure 10: shows a structure of a device for simultaneous modifying several
modes.
Figure 11: shows a progression of the lateral force Qy normalised to the
maximum
lateral force Qym~, plotted versus the normalised body length of the
airplane, when damping several modes without and with application of a
device according to an exemplary embodiment of the present invention.
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Detailed description of exemplary embodiments:
Figure 1 shows an airplane 1, the body of which vibrates in a first and a
second
bending mode 2,3. In the illustration of figure 1, the airplane 1 moves in the
direction
of flight 4. Throughout the following, the term "mode" is used as abbreviation
of the
terms "rigid body mode", "rigid body movement mode" and/or "elastic
eigenmovement or mode", respectively. The airplane 1 has a body 5, a nose 6
and a
back 7. In the region of the nose 6, a sensor means 8 is disposed. In the
region of the
back 7, there are disposed controlling, guiding and regulating surfaces 9, for
example
in form of ailerons, trim surfaces, spoilers and rudders.
By means of a fixed law of regulation, it is not possible, to effectively damp
two
adjacent eigenmovements of the airplane 1 having closely neighboring
frequencies, as
well as different phase positions (knot positions), because the phase position
can only
be adapted to one eigenmovement at a time. If, for example, the nose 6 of the
aircraft
1 moves in the direction of the arrow 10 with a certain velocity v, according
to a law
of regulation assumed here, for example a force F had to act at the back 7 in
the
direction of the arrow 11. If the regulation means now were to act in the same
way for
all modes, according to this law of regulation, the second bending mode 3
would
unintentionally be excited simultaneously. Thereby, the nose 6 is moved in the
direction of an arrow 12 with a velocity v, and a force F in the direction of
the arrow
13 would in the range of the back 7 act upon this, regulated by the regulation
means.
As a result, the application of the same law of regulation to the first and
the second
bending mode 2,3 would lead to a destabilising of the 2. bending mode.
Figure 2 shows the progression of the lateral force Qy normalised to the
maximum
lateral force Qymax, plotted versus the normalised body length of an airplane
in case of
an excitation with continuous blasts.
At the vertical axis of the diagram, there is plotted the progression of the
lateral force
QY, respectively referring to a maximum lateral force with respect to a
certain
position x/x",aX at the body 5 of the airplane 1. A curve progression 14
represents the
CA 02510115 2005-06-16
maximally permittable lateral force Qy in the respective position x/Xmax in
the body of
the airplane 1. A curve progression 15 reflects the actual progression of the
lateral
force QY, which results without the application of the inventive device. It
may be
taken from the diagram of figure 2 that in a region of the accentuation 16, a
overloading of the body structure or the body 5 of the airplane 1 may occur.
This
region is between the nose 6 and the region of the body 5 in which usually the
airfoils
are disposed. In this region, the allowable maximum value of the lateral force
Qy is
exceeded to a remarkable extent, as is shown by the curve progression 1 S,
which
extends above the curve progression 14.
»gure 3 shows a spectrum of the lateral force Qy, normalised to the maximum
lateral
force QY max for a position at the body's front part, plotted versus the
frequency f
normalised to a maximum frequency fm~, of the airplane 1, when in a dominant
mode.
At the vertical axis of the diagram, the progression of the lateral force QY,
each
related to a maximum lateral force QYmaX at a certain normalised frequency
f/fmax is
plotted. In the region of an accentuation 17, a dominant elastic mode 18 is
clearly to
be seen, which significantly contributes to increased load values within the
body
structure or the body 5, respectively, of the airplane 1.
The diagram shown in figure 4 shows an acceleration spectrum in the region of
the
position of a sensor means, when in a dominant mode
To this end, the sensor means 8 is preferably situated in the front body
region of the
airplane 1. At the vertical axis of the diagram, the progression of the
acceleration
nY/nym,~ in case of certain normalised frequencies f/fmaX is plotted. In the
region of an
accentuation 19, a dominant eigenmovement 20 shows-up again. The maximum of
this dominant eigenmovement 20 is in the range of about 33% of the maximum
frequency fm~.
In figure 5, the schematic structure of the device 21 according to an
exemplary
embodiment of the present invention is shown.
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The illustration of figure 5 substantially concerns the measurement of the
lateral
acceleration ny in a front body region of an airplane 22. The airplane 22, for
example
a big transport airplane or passenger airplane in a long version, has a body
23, a nose
24, a back 25, as well as airfoils 26. In the region of the nose 24, a sensor
means 27 is
disposed. A front body region 28 of the airplane 22 is laterally loaded by
(lateral)
blasts 29. The sensor means 27 serves to measure the lateral acceleration
values nY
generated by the blasts 29 impacting upon the body 23. As a consequence of the
blasts
29, rigid body modes and/or elastic aircraft modes are excited in the body 23,
as is
symbolised by the arrows 30,41. A measurement signal 31 outputted by sensor
means
27 after a respective measurement technological processing, for example in an
amplifier not shown here and/or in an anti-abasing filter is led to a
regulation
arrangement or controller 33 and a frequency measurement arrangement 34 in
parallel. From the frequency measurement arrangement 34, a frequency signal 35
is
led to the regulation arrangement 33. The frequency signal 35 represents the
frequency of at least one rigid body mode and/or at least one elastic
eigenmovement
of the airplane 22. The frequency measurement arrangement 34 also extracts the
frequency signal 35 from the measurement signal 31 of the sensor means 27, for
example by means of a digital analysis within a digital calculation unit. For
the
operation of the device 21 according to an exemplary embodiment of the present
invention, a sensor means 27, for example in form of an acceleration sensor
for
detecting a lateral acceleration nY, is usually sufficient. Accordingly,
accelerations in
other spatial directions can be detected, to which end the modified sensor 27
may be
provided.
Within the regulation arrangement 33, a mixed leading back signal or feedback
signal
36 (comprising several lives) is generated, which, by means of actuators, not
shown in
detail in figure 5, is led back again to controlling, guiding and regulating
surfaces 40
of the airplane 22 like, for example, aileron 39, rudder 38, spoiler 37, as
well as other
guide and trim surfaces as, for example, elevator, "Mini-Teds", "Tabs",
"Goumey-
Flaps", "Canards"and so on. The effect of the leading back signal 36, lead
back from
the regulation arrangement 33 via the actuators to the airplane 22 again is
symbolised
by the arrow 41.
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Due to the rigid body modes and/or elastic eigenmovements of the airplane 22,
detected by the sensor means 27 which within the regulation arrangement 33 are
subject to a suitable modification for damping the unwanted modes and as a
leading
back signal 36 by means of the actuators are subsequently led back to the
airplane 22
via the controlling, guiding and regulating surfaces 40 there results an
effective
damping of at least one rigid body mode and/or at least one elastic
eigenmovement of
the whole airplane 22.
Figure 6 shows the detailed structure of the regulation arrangement for
modifying a
mode.
The regulation arrangement or controller 42 substantially comprises a
regulation line
43. The regulation line 43 is constituted of a resonance filter 44, the phase
correction
units 45, 46, the amplifiers 47, 48 and an inverter 49.
To the resonance filter 44 is supplied a measurement signal 50, for example,
in form
of lateral acceleration values nY measured by a sensor means in a front body
region of
the airplane. A signal 51 of the resonance filter 44 is then led to the phase
correcting
units 45, 46 in parallel. The phase correcting units 45,46 are each
respectively
followed by one of the amplification factor units 47,48. The amplification
factor unit
48 is followed by an inverter 49. The amplification factor units 47, 48, as
well as the
inverter 49 form several leading back signals or feedback signals 52, which
are led
back via actuators which are not shown in detail in figure 6 to controlling,
guiding and
regulating surfaces of the airplane. Due to the connecting of the leading back
signals
52 to respective controlling, guiding and regulating surfaces of the airplane,
especially
to aileron, rudder, spoiler, guiding surfaces, trim surfaces and the like,
there results a
damping of at least one rigid body mode and/or at least one elastic
eigenmovement of
the airplane as a whole. In the specific embodiment of figure 6, for example,
the
leading body signal 53 is led back to at least one aileron of the airplane, a
leading
back signal 54 (or feedback signal) is lead back to at least one aileron of
the airplane,
and a leading back signal 55 (or feedback signal) is led back to at least one
spoiler,
one trim surface or the like of the airplane.
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The inverter 49 allows for a simple structure of the regulating unit 42,
because the
leading back signal 54 for actuating the actuators for the aileron, and the
leading back
signal 55 for actuating the actuators for operating the spoilers have a phase
shift of
180° with respect to each other, so that by means of the inverter 49
from the leading
back signal 54 for the aileron, the leading back signal 55 for actuating the
spoiler may
be generated by means of a simple inversion of sign in the inverter 49.
Thereby, a
phase correction unit may be economised. According to the number of modes or
elastic modes of the airplane to be influenced, respectively, several
regulated lines or
regulated branches, respectively, have to be set in parallel. In a preferred
manner,
within the framework of the signal processing for each rigid body mode and/or
elastic
eigenmovement of the airplane to be modified, the regulation arrangement 42
has a
separate regulated line, the structure of which corresponds to the regulated
line 43,
which has been explained further above within the framework of the explanation
of
figure 6. For signal processing within the regulation means 42, a filter of
third order
may be sufficient. Alternatively, a filter of higher or lower order may be
provided.
In the embodiment shown in figure 6 the resonance filter 44 is applied as
filter of
second order. The measurement signal 50 is at first supplied to the reasonance
filter
44. The resonance filter 44 thereby follows the transfer functionTF:
2* D* H62
TF -
s2+ 2* D* V16* s+ W(72
The phase correcting units 45, 46 of first order obey the transfer function
-c* s+1
TF-
c* s + 1
Further there are provided two amplification factor units 47, 48 which follow
the
phase correcting units 45, 46. In the above transfer functions, D refers to
the damping
and wo corresponds to the undamped eigenfrequency of the resonance filter 44.
By
means of the constant c, the phase shift in the phase correcting units 45,46
may be
adjusted. The function of the resonance filter 44 is a discrete realtime
fourier analysis.
Thereby, there exists an easily realisable possibility, as far as system
technology is
CA 02510115 2005-06-16
14
involved, to obtain information concerning the searched vibration mode from
the
physical sum signal in form of the measurement signal 50 which exists in form
of a
discrete acceleration signal from the sensor. In an exemplary embodiment of
the
invention, the regulation means 42 is realised by means of a digital
calculator unit.
The adjustment of the parameters contained in the transfer functions may
require an
analysis of the rigid body modes and the elastic eigenmovement of the
airplane, in
which a device according to an exemplary embodiment of the present invention
shall
be implemented.
To the one hand, the resonance filter 44 is described by its resonance
frequency which
corresponds to the resonance frequency of the airplane eigenmovement and/or
elastic
mode to be modified, and on the other hand by means of its damping, which
shall turn
out preferably low, in order to reach the required small band characteristic
of the
resonance filter 44.
The simple form of the resonance filter 44 stresses the advantage of the
selective
approach chosen here, as compared to previously known regulation systems. By
means of only one filter of second order with only two parameters per mode
which
may easily be determined, the intended signal information, the required signal
information concerning the mode to be damped, that is the rigid body mode
and/or the
elastic mode to be damped may be selected directly by means of the device or
the
method in accordance with an exemplary embodiment of the present invention. In
contrast, in case of the subtractive method according to the state of the art,
in case of
an airplane having n modes, up to n-1 unwanted modes have to be filtered out.
This
means, there arises the (n-1)-fold effort in the design and in the realisation
of such
regulation systems.
The subsequent modification of the phase in the phase correcting units 45, 46
serves
to compensate the sensor's dead time, the signal processing time in the
regulating
arrangement 42, the phase delay of the actuators, the phase delay of the in
stationary
aerodynamics of the airplane, the provision of the mixed leading back signals
(or
CA 02510115 2005-06-16
1S
feedback signals) 52 (synthesis of velocity and acceleration) as well the
phase
position of the complex eigenvector of the airplane's relevant mode to be
damped
The mixed leading back signals or feedback signals 52 serve to obtain the
resonance
frequency of the mode to be modified in the closed regulation-loop. This is
necessary
to secure the operability of the resonance filter 44. By the term "mixed
leading back"
in this context is meant the combined leading back of acceleration signals and
velocity signals. By means of the filtering, the velocity signal is obtained
from the
acceleration signal, so that no any additional separate sensor is required to
this end.
Taking into consideration the phase position and the local amplitude values of
the
complex eigenvector of the movement allows for a what is believed to be
optimal
adaptation of the regulation arrangement 42 to the modes of the airplane.
A device according to an exemplary embodiment of the present invention or a
method
according to an exemplary embodiment of the present invention is based on a
modal
view and, derived therefrom, on a criterion for phase adaptation and amplitude
adaptation of the guiding, controlling and regulating surface deflections by
means of
the amplification factor units 47,48 to the eigenvector of the to be modified
elastic
eigenmovement and/or the rigid body mode.
This may allow for a maximum influenceability of the movement when a control
effort and regulation effort is minimal, and on the other hand the appearance
of linear
distortions (excitation of other modes with the frequency of the target mode)
is
prevented. The interrelationship between the mode (eigenvector of the mode) to
be
affected, which means to be damped, and the input signal or kind of input
circuitry
required for influencing only the one mode which is lead back, no other modes,
serves
as a basis.
In the specific case, this means that the damping of a body mode requires the
presence
of ailerons, rudders, spoilers or the like, operated by actuators. If, for
adapting the
regulating arrangement 42, a state vector feedback is assumed (the relevant
part of the
state vector is created by filtering the original vector), it may be shown in
modal form,
that, for exclusive modification of a discrete eigenmovement, the number of
necessary
CA 02510115 2005-06-16
16
control inputs has to be equal to the number of physical degrees of freedom
From this
it may be followed, that theoretically, the application of compensation
forces, created
from leading back or feedbac acceleration (or position, respectively) and
velocity to
all mass-points of the aircraft, would be necessary for complete damping of
all
relevant modes. In practice, this is not may not be realized in an ideal
manner, due to
the high number of degrees of freedom Taking into account these theoretical
boundary conditions, a what is believed to be an optimal approach may be
obtained
by means of the available guiding, controlling and regulating surfaces.
From a load analysis according to the diagram of figure 3 at a selective
position in the
front body region of the airplane, the dominance of the elastic mode 18 can be
seen,
which significantly attributes to the overall mechanical stress of the
airplane. With
respect to this one dominant elastic mode 18, in the following shall at first
be
explained, how this may be damped effectively by means of a device according
to an
exemplary embodiment of the present invention. In this respect, the regulation
arrangement 42 comprises only one regulation line 43, since at first only the
single
dominant mode 18 shall be damped.
The diagram of figure 7 at first shows the progression of the lateral force QY
normalised to the maximum lateral force Qymax, plotted versus the normalised
body
length of the airplane without and with applying a device according to an
exemplary
embodiment of the present invention, as compared to a progression of the
maximally
acceptable load.
At first all three curve progressions shown have in common that there is a
discontinuity-position. in the region of a body position x/x"~X of about 40
percent,
This discontinuity is approximately in a region, in which the airfoils are
disposed at
the airplane.
At first, a curve progression 56 illustrates the maximum acceptable lateral
force
QY~QYmax wherein the lateral force Qy is respectively related to a maximal
lateral
force Qyn,aX, which must not be exceeded in any case at a certain body
position x/xmaX
of the airplane. A curve progression 57 symbolises the progression of the
lateral force
CA 02510115 2005-06-16
17
Qy/ Qym~ which results without the application of a device according to an
exemplary embodiment of the present invention. In this respect, it shall be
noted that
there results an overloading in the range of an accentuation 58, which means
in the
front body region of the airplane. A curve progression 59 eventually reflects
the
progression of the lateral forces QY/ QYmaX resulting from the application of
a device
according to an exemplary embodiment of the present invention. Thereby, the
curve
progression 59 is always considerably below the maximally acceptable lateral
forces
QY/QYmax according to the curve progression 56.
Subsequent to respective design and dimensioning of the regulation arrangement
according to the criteria stressed further above, within the frame of the
description of
figure 6, for the airplane having a device according to an exemplary
embodiment of
the present invention, there are obtained substantially reduced stresses in
the front
body region. Especially the dominant mode creating this stress (see figure 3)
was
effectively damped by means of a device according to an exemplary embodiment
of
the present invention.
The diagram of figure 8 shows the spectrum of the lateral force Qy, normalised
to the
maximum lateral force Qym~, plotted versus the frequency f normalised to a
maximum frequency fmax in the front body region of the airplane without and
with
application of the device.
The curve-progression 61 represents the spectrum of the lateral force Qy/QYmaX
without application of a device according to an exemplary embodiment of the
present
invention or a method according to an exemplary embodiment of the present
invention, respectively. The curve progression 62 reflects the spectrum of the
lateral
force Qy/QymaX though, which results from applying a device according to an
exemplary embodiment of the present invention or a method according to an
exemplary embodiment of the present invention, respectively. It can be seen
that in
the region of about 25 percent of the maximum frequency of the spectrum fmaX,
a
significant lowering of the amplitude of the dominant mode of the lateral
forces
QY/QYmax of the body of the airplane results from the application of the
device.
CA 02510115 2005-06-16
1g
Thereby, the mechanical load of the body or the body structure of the
airplane,
respectively, is substantially reduced.
The diagram of figure 9 shows the acceleration spectrum ny/nymaX plotted
versus the
frequency f/ f~X normalised to a maximal frequency in the range of the sensor
means
in the front body region of the airplane without andwith application of a
device
according to an exemplary embodiment of the present invention.
At the vertical axis of the diagram, the lateral acceleration values nY, each
related to a
maximum acceleration value nY",~ in the region of the position of the sensor
means
are plotted versus the frequency ratio f/f~X
A curve progression 63 represents the acceleration values nY/nYt"aX without
the
application of the device, whereas a curve progression 64 reflects the
acceleration
values ny/ny",aX that result from the application of the device. From the
representation
of the diagram of figure 9, it can be seen, that the curve progression 64
which results
from the application of a device according to an exemplary embodiment of the
present
invention is substantially below the curve progression 63 in the region of a
dominant
mode 65, so that reduced acceleration values nY/nYm~ result. The acceleration
values
ny/nymax at certain positions of the body of the airplane are relevant for
stability
problems and comfort problems.
As a consequence of the application of a device according to an exemplary
embodiment of the present invention or a method according to an exemplary
embodiment of the present invention, respectively, there may result a
substantial
improvement of the stability and the comfort of the airplane.
Figure 10 illustrates a detailed inner structure of an embodiment of the
regulation
arrangement or controller 66 within a device according to an exemplary
embodiment
of the present invention for modifying two modes at the same time.
The internal structure of the regulation arrangement 66 in principle
corresponds to the
structure of the regulation arrangements already described, which are directed
to
CA 02510115 2005-06-16
19
damping one mode. In contrary to the already described regulation
arrangements, the
regulation arrangement 66 has two regulation lines 67,68. The regulation line
67
substantially comprises a resonance filter 69 as well as two parallel
following phase
correction units 70,71, wherein following to each of it, an amplification unit
72 and
73 is disposed. Accordingly, the regulation line 68 comprises a resonance
filter 74, to
which the phase correcting units 75,76 are following in parallel. The phase
correcting
units 75, 76 themselves each are followed bythe amplification units 77 and 78.
To the
resonance filters 69,74 is supplied a measurement signal 79, for example, in
the form
of a lateral acceleration signal nYdetected by the sensor means. Further, the
regulation
arrangement 66 has the adding units 80,81. The adding units 80,81 are only
necessary,
if two or more modes, i.e. rigid body mode and/or elastic modes, are to be
damped. In
this case, the output signals of the respective resonance filters 69, 74 (one
per mode,
respectively) are adjusted individually with respect to phase and amplitude by
means
of the phase correcting units 70,71,75,76 and the amplifying units
72,73,77,78, and
are subsequently combined into one respective leading back signal or feedback
signal
84,87,89 per controlling, guiding and/or regulating surface.
A leading back signal or feedback signal 82 of the amplification factor unit
72 is led
to the adding unit 80. In order to create a leading back signal or feedback
signal 84, a
leading back signal or feedback signal 83 of the amplification unit 77 is led
to the
adding unit 80 as well. The leading back signal 84 serves to actuate
respective
controlling, guiding and/or regulating surfaces of the airplane, for example
in form of
the rudder, by means of actuators, which are not shown in further detail. A
leading
back signal 85 of the amplification factor unit 73 is led to the adding unit
81. Further,
a leading back signal 86 of the amplification factor unit 78 is also led to
the adding
unit 81 for creating a leading back signal 87. The leading back signal 87 in
tum serves
to actuate respective controlling, guiding and regulating surfaces of the
airplane, like,
for example, the ailerons, by means of actuators not shown in further detail.
Eventually, the leading back signal 87 behind the adding unit 81 is led in
parallel via
an inverter 88 for creating a leading back signal 89. The leading back signal
89 serves
for actuating respective controlling, guiding and regulating surfaces of the
airplane,
especially for actuating the spoilers by means of actuators not shown in
further detail.
In using the inverter 88, an allowable phase correcting unit is economised,
since the
CA 02510115 2005-06-16
necessary leading back signals 87,89 between the ailerons and the spoilers of
an
airplane in this case show a phase shi$ of 180°.
In the diagram of figure 10, it is clearly to be seen that for damping each
rigid body
mode and/or elastic eigenmovement of the airplane there is necessary an
individual
respective regulation line 67,68. Thereby, all components of the regulation
arrangement 66, with respect to the modes to be damped of the airplane have to
be
configured individually. In contrary to the known regulation arrangements
there are
not filtered out all unwanted modes besides the one to be damped in a
subtractive and
therefore involved procedure, but by means of the resonance filters 69,74 only
the
relevant modes are passed for further processing in the regulation arrangement
66.
By means of a device according to an exemplary embodiment of the present
invention, not all disturbing modes are damped equally. It is possible that a
strongly
disturbing mode is damped considerably strong while a weaker mode- a mode
which
possibly is not disturbing at all- is even amplified a bit. If the integral
over all modes
in the frequency range from 0 to 10 Hz is considered, in any case, in an
integral point
of view, there results a substantial damping, as compared to a system not
damped by
means of a device according to an exemplary embodiment of the present
invention.
Figure 11 shows the progression of the lateral force Qy normalised to the
maximum
lateral force Qymax, plotted versus the normalised body length x/xmax of the
airplane,
when damping two modes, each with and without application of the device, as
well as
compared with a further curve.
A curve progression 90 reflects the maximally acceptable loading values. A
curve
progression 91 symbolises the progression of the lateral forces without the
application
of a device according to an exemplary embodiment of the present invention.
From the
comparison of the curve progressions 90,91, it results that the allowed limit
values of
the normalised lateral force Qy/QYmaX is transgressed at least in the range of
an
accentuation 92. By means of the application of a device according to an
exemplary
embodiment of the present invention, there results a falling below the
maximally
CA 02510115 2005-06-16
21
allowed lateral forces over the whole body length of the airplane, as is
illustrated by a
curve progression 93.
According to a method according to an exemplary embodiment of the present
invention (see figure 5), the measurement signal 31 is supplied to the
regulation unit
33, for example in form of the lateral acceleration signal ny of the
regulation unit 33
within a device according to an exemplary embodiment of the present invention.
The
measurement signal 33 is detected by means of a sensor means 27 in the front
body
region of the airplane 22, and in this context represents at least one actual
elastic
eigenmovement and/or at least one elastic eigenmovement of the airplane 22.
Within
the regulation arrangement 33, the processing of the measurement signal 31 to
a
leading back signal 36 is effected by respective filter means. The internal
structure of
the regulation arrangement 33 thereby follows the structure explained within
the
frame work of the description of figure 6 and figure 11. According to a
further
embodiment of the regulation arrangement 33 for carrying out aspects according
to
the invention (which is not shown in the figures in further detail), it is
possible to
provide an arbitrary number of parallel filter lines.
The leading back signal 36 during the carrying out of the procedure according
to the
invention by means of actuators not shown in further detail is lead back to
respective
controlling, guiding or regulating surfaces of the airplane 22 especially
aileron,
rudder, spoiler and the like, for damping at least one critical rigid body
mode and/or at
least one elastic eigenmovement. Thereby an effective damping of the critical
rigid
body modes and/or elastic modes of the airplane results.
Prior to an effective application of a device according to an exemplary
embodiment of
the present invention or carrying out a method according to an exemplary
embodiment of the present invention, the regulation arrangement at first has
to be
accordingly configured during the development phase of the airplane.
This is effected by a complete analysis of the transfer functions and modes of
the
integral flight-dynamics model of the airplane. The results of the complete
analysis
are then used for selecting suitable sensor means, suitable positions for
disposing the
CA 02510115 2005-06-16
22
sensor means as well as suitable controlling, guiding and regulating surfaces
for
optimal damping critical rigid body modes and/or elastic modes. Preferably, as
sensor
means are used acceleration sensors which are positioned in the front body
region of
the airplane.
According to a variant of a method according to an exemplary embodiment of the
present invention, at the same time, the frequencies of the rigid body modes
to be
modified and/or the elastic eigenmovements of the airplane are detected by
means of
the frequency measurement arrangement. To this end, the frequency measurement
arrangement outputs a frequency measurement signal, which is supplied to the
regulation arrangement. Thereby, an adaptation to different mass relations of
the
airplane during the course of the flight, for example caused by the continuous
consumption of fuel, is taken into account.
By means of a phase and amplitude consideration of the relevant modes, the
resonance filters, the phase correcting units as well as the amplification
factor units
are configured in such a way that when leading back the leading back signals
outputted by these units to the controlling, guiding and regulating surfaces
of the
airplane, an optimal damping of the selected rigid body modes and/or the
elastic
modes results.
By means of a method according to an exemplary embodiment of the present
invention, it is possible, as a result, to effectively damp the amplitude of
at least one
rigid body mode and/or at least one elastic eigenmovement of the airplane,
wherein
the regulation technological effort is as small as possible. In comparison to
the known
methods, which are based on a subtractive removal of unwanted modes besides
the
mode to be modified, there results a considerably reduced system complexity.
A device according to an exemplary embodiment of the present invention, as
well as a
method according to an exemplary embodiment of the present invention may be
applied to arbitrary local and global load problems, stability problems and
comfort
problems, especially in big airplanes. Thereby for the first time results the
possibility,
in using only one measurement signal of a discrete sensor means, for example
in form
CA 02510115 2005-06-16
23
of a lateral acceleration sensor nY, to specifically modify an arbitrary
number and an
arbitrary combination of rigid body modes and/or elastic eigenmovements of the
airplane, without influencing modes to be preserved in a negative way. This is
made
possible by the structure of a modal regulation without the necessity of an
observer,
so that, as a result, a considerably reduced system complexity results.
According to an
exemplary embodiment of the invention, the adaptation of the deflection to the
selected controlling, guiding and regulating surfaces to the local amplitudes
of the
modes for allowing a maximal damping may be possible, when, at the same time,
a
control effort is minimal. The required deflections of the selected
controlling, guiding
and regulating surfaces are only 10 to 15 percent, as compared to the previous
methods or devices, respectively. Thereby, by means of a device according to
an
exemplary embodiment of the present invention or a method according to an
exemplary embodiment of the present invention the loading of the actuators may
be
reduced. In turn, this may allow for a a reduced air resistance of the
airplane which
may allow to reduce the consumption of fuel at the same time.
At the same time, a device according to an exemplary embodiment of the present
invention or a method according to an exemplary embodiment of the present
invention, respectively, may prevent the generation of unwanted linear
distortions in
the regulation loop and accounts for no further unwanted modes being excited.
Finally, it is noted that a device according to an exemplary embodiment of the
present
invention or a method according to an exemplary embodiment of the present
invention is not to be seen as limited to applications for reduction of
structural load or
damping of vibration, respectively, in big airplanes. Application
possibilities for
reduction of structural load or damping of vibration, respectively, are given
in
connection to all bigger spatial structures, like, for example, also at ships,
bridges and
buildings.
CA 02510115 2005-06-16
24
List of reference characters
1 airplane
2 first bending mode
3 second bending mode
4 direction of flight
body
6 nose
7 back
8 sensor means
9 controlling, guiding and regulating surfaces
11 arrow
12 arrow
13 arrow
14 curve progression
curve progression
16 accentuation
17 accentuation
18 dominant mode
19 accentuation
dominant eigenmovement
21 device
22 airplane
23 body
24 nose
back
26 airfoils
27 sensor means
28 front body region
29 blasts
arrow
31 measurement signal
33 regulation arrangement
34 frequency measurement arrangement
CA 02510115 2005-06-16
35 frequency signal
36 leading back signal
37 spoiler
38 rudder
39 aileron
40 controlling, guiding and regulating surfaces
41 arrow
42 regulation arrangement
43 regulation line
44 resonance filter
45 phase correcting unit
46 phase correcting unit
47 amplification factor unit
48 amplification factor unit
49 inverter
50 measurement signal
51 signal
52 leading back signals
53 leading back signal
54 leading back signal
55 leading back signal
56 curve progression
57 curve progression
58 accentuation
59 curve progression
60 curve progression
61 curve progression
62 curve progression
63 curve progression
64 curve progression
65 dominant mode
66 regulation arrangement
67 regulation line
CA 02510115 2005-06-16
26
68 regulation line
69 resonance filter
70 phase correcting
unit
71 phase correcting
unit
72 amplification
factor unit
73 amplification
factor unit
74 resonance filter
75 phase correcting
unit
76 phase correcting
unit
77 amplification
factor unit
78 amplification
factor unit
79 measurement signal
80 adding unit
81 adding unit
82 leading back signal
83 leading back signal
84 leading back signal
85 leading back signal
86 leading back signal
87 leading back signal
88 inverter
89 leading back signal
90 curve progression
91 curve progression
92 accentuation
93 curve progression
