Title of the invention
A method for real time structural integrity monitoring & control of an
aircraft.
2. Field of Invention
The present invention relates to a method comprising of a novel real time
structural integrity monitor & controller, by real time monitoring of strain
and vibration values at critical location on aircraft. In case of reaching near
to maximum permissible limit of strain or vibration values, an
instantaneous automatic control of aircraft speed carried out‘ Additionally,
two indications to Pilot is made available, one to perform best possible
operation, as the requirement may be, like level the aircraft to reduce
strain or vibration using real time processed values and sécond displays
as history for awareness, so that Pilot shall not repeat the same profile
again where strain or vibration exceeds the level or may decide to land
safely. The computation is based on real time strain sensor data, vibration
sensor data, Global positioning system (GPS) data and aircraft speed
data.
3. Background of invention
The background of invention addressed two important aspect ,first is the
growing costs of flight testing leads to an optimization of number of test
flight required for flight evaluation of newly integrated & modified structures
exposed externally mainly on wings &fuselage. Secondly during routine
flight, monitoring of structural integrity of the structure of an aircraft which
might have been weakened due to corrosion, cracked surface or loose
structural interconnect. The corroded cracked or loose structure may not
be noticed generally prior to sortie, it is noticed generally during aircraft
overhauling. The changes in aircraft external body, loose structural
interconnect, corroded or cracked may simultaneously lead to unintended & unacceptable phenomenon like compression or elongation at
structurally fixed attachment or generation of the criticaI vibration values.
The dynamic load may be increased during flight and further needs to
avoid any additional increase in dynamic load otherwise required for
strengthening of the structure. Until now it has been practised of gradually
extend and evaluate aerodynamic load on aircraft joint structure for
preventing "over compression or over elongation" at the attachment
beyond technically specified limit by incremental offline analysis. Also
needs gradual validation of vibration level at each critical point on aircraft
during design phase. The problem in hand involves firstly ascertain
charaéteristic envelope of respective elasticity of aircraft frame or
component which becomes increasingly less damped at increasing
dynamic or impact loads and secondly unacceptable elongation and
compression leading to structural failure. Normally, such phenomenon like
vibration leads to failure of the respective structure The mathematically
computed vibration damping system used in high performance aircraft
needs to increase the speed causing vibration to a prescribed maximum
level. To validate this mathematical model in optimal manner, this novel
integrity control system is designed. This novel system which comprises
strain gauges, accelerometers, controllers and adjusting elements,
controls the characteristic behaviour of the structure. In other words, such
system achieves a real time control based on computed aerodynamic
damping and strain values. The novel system contains two databases; first
the vibration database containing maximum permissible values allowed
as adjusting elements, which are periodically validated at the characteristic
frequency of the attachment or wing, in order to achieve maximum safety
on the aerodynamic damping effect, second the strain database containing
maximum permissible values allowed as maximum compression, or
elongation permissible .Pilot has real time control in flight based on
database. Also ,there is requirement of ,In order to avoid high vibration
with respect to increase with speed, additional safety mechanism and measurement systems (including sensors) must be monitored and
analysed such as using installed devices on an aircraft, in response to
signals generated by sensor means providing a vibration representing
value. Such sensors are integrated with control system which provide a
control signal fo'r the control of the damping means in response to a speed
and attitude which exceeds a certain selectable threshold speed value.
Thé sensors are strain or vibration sensors which provide the respective
control signal for the damping means in response to an strain and
vibration level which exceeds in its amplitude or frequency a
selectable threshold amplitude or frequency value. These threshold
values signal or anticipate a contour of vibration or strain. The damping
means are then maintained stationary in the damping position until the
undesired vibration causing instability has been damped sufficiently or
until the aircraft speed is reduced to a level below the threshold or critical
level. According to the invention the damping means are not moved
repeatedly for the same vibration situation by displaying list of aircraft
speed and attitude values causing instability..ln typical prior art, in case of
integration of new structural component on aircraft ,it is required that
structural integrity of new component to be tested with aircraft envelbpe viz
pitch ,roll and speed level , manually operated aircraft controls were
selected by Pilot to call off the flying decision and envelope is gradually
expanded based on offline analysis of previous flight data. When airbraft
is moved to close limit and it is in various phase, i.e. when the lever is set
to the rapid roll position and any aircraft vibration is detected, under such
circumstances, the effort of reducing the aircraft speed operation can be
extremely reduced by the usage of automatic integrity controller . Pilot can
select auto mode using a toggle on-off switch mounted on cockpit to select
whether‘automatic or manual control is required in case of test sorties
.Along with this switch. novel safety controller using various inputs like
throttle input, aircraft current speed input, forward acceleration input, strain
and vibration in real time , are processed, making the task of aircraft speed control autométic in place of any selection by the Pilot. A novel safety
control system for controlling a speed control proVided to an aircraft
includes cut off of throttle position and operating a control surface when
aircraft is reaching safe profile. The novel controller controls the aircraft,
when a predetermined limit condition is satisfied. A throttle position is cut
‘
off, which helps to reduce aircraft speed and indicate Pilot to cérw out
certain manual operation as per requirement. In summary, prior art had
following drawbacks:
‘
a) For Pilot, there is no auto control system to operate when
structural integrity limit is reached find is detected and
automatically operate other system for speed reduction.
b) No poséibility to know the safe strain, vibration limit to decide
speed reduction in real time.
c) No possibility to know the structure weakening due to corrosion
level or structurally loose interconnect ,unsafe during flight
These problems were substantially overcome by the system which is
the subject of this invention.
4. Summary of the Invention
The invention relates to a method for the structure integrity monitoring &
controlling of a high speed aircraft, for expanding the aircraft envelope in
much optimized manner, monitor structure weakening due to corrosion,
crack or structurally loose interconnect which are unsafe during flight. For
this, the strain gauges and vibration sensor are placed into a suitable
position in accordance to mathematically model indicating probable
location of high stress and vibration levels. The levels may cause a
catastrophic structural failure in the local area of an aircraft during flight
.As the local stress point reaches to maximum pgssible limit or closer to
the threshold point during flight to thereby controller provides a suitable
reduction in strain and vibration values. The control system activates
control which includes reduction in speed and levelling the aircraft when needed and stay in an activated effective position for a length of time
sufficient for decreasing the vibration and or strain values using
speed .The control activation takes place rapidly in about 20 to 40
milliseconds from the time the sensors pick up or provide a signal for such
control activation. During trial phase ,it is essential to monitor the safety of
aircraft ,specially vibration and strain condition of aircraft .Also corrosion
or crack or loose interconnect Im>ay deteriorate performance of structure
component . In case of any exigency condition on aircraft ,certain set of
manual operation also has to be performed ,which is more manual
reaction time, to alter the movement of aircraft Especially for envelope
related decision, in case of change of speed and manoeuvring
requirement, it has to be determined very precisely based on current
aircraft speed and acceleration & sensor data of aircraft. This decision
depends upon Pilot’s skill‘ There was no automatic method to compute in
real time .the safe levels using set of auto sensing and auto operation
.There is high chance nonjudgmental error which may be catastrophic .Also,
in case based on error in manual judgement Pilot may opt for ejection in
case of catastrophic aircraft structure failure .Hence as an additional safety
measure, Pilot is provided with additional feature of auto integrity control
using novel auto control system. This auto control system does real time
computation of safe profile based on various aircraft inputs and does auto
operation. Auto operation helps the aircraft to reduce speed using throttle
helps to stop acceleration .Also pilot has feature of real time display of
control system outputs automatically. Also Pilot have indication for
additional manual operation to control aircraft further .This lead Ito
maximum possible control using throttle operation ,Levelling the aircraft
This ultimately helps Pilot to bring aircraft lat safe speed and eventually
damp the vibration on the aircraft or strain as the case may be. The
interface between a Pilot and the control Surfaces of an aircraft, especially
in case of nearing vibration during manoeuvring, needed much attention.
The Pilot control interfaces are simple but needed a great deal of pilot skill to control aircraft in case of Vibration. The novel controller helps, Aircraft,
with uses of microcontroller based electronics and interfaces to assist the
pilot for semi—automated operation and make envelope condition of flight
safer. In no "automatic” interface, the pilot interacts with an manual user
interface‘ using a dedicated selection indicators in the cockpit .This
controller also enables Pilot to select auto operation mode during aircraft
speed exigency in case of vibration. This feature is helpful for trainee Pilot
with less flying experience. Algorithm in novel controller is constantly
working to make the and control the interface safer, easier to understand
and automatically operate, and more effective, efficient and reliable way.
In terms of up gradation on old aircraft , limited digital control laws are
used to implement control laws that use a reference command based on
vibration sensor at aircraft area, GPS data, aircraft speed ,throttle position
,acceleration and combination of thereof. Aircraft speed in conjunction with
a attitude profile and sensor data is considered as a reference command.
In No “automatic “aircraft, the Pilot manually moves (deflects) a throttle
stick and operate to provide a control command during speed flight
phases. In this case, a manual operation provides reduction in aircraft
speed with stability‘ This type of control provides average handling
qualities while in normal integrity condition but not necessarily vibration
condition. More specifically, any sudden aircraft vibration condition’ often
do not provide suitable action time, and ih particular to reduce speed
during speed. Generally speaking, speed is initiated by increasing the
aircraft‘s throttle position as high as possible to increase the acceleration
to a desired amount. The thrust ideally is smoothly decreased from
maximum to idle as the vibration may progress. However in case of aircraft
vibration in short span of time, based on current aircraft speed,
acceleration has to be stopped .For this, the information of sensors is
used to change the control law to control aircraft speed stability near to the
safety. Particularly, a control law in case of vibration with positive speed
stability and throttle control operation that provides a control law with positive speed stability. Change in the speed can only be accomplished
while parameter like speed and acceleration is within control limit. Under
Such circumstances, the effort of control aircraft based on human
judgement needs little time .In automated system the aircraft speed can be
‘
extremely reduced by the usage of a microcontroller processing which in
turn momentary change the throttle position, making easier the task of
aircraft speed reduction without the pilot intervention. This controller
provides good handling qualities during rapid acceleration and rapid roll ,
with the benefit of not needing or reading manually sensor information in
safety-critical aircraft vibration condition .In an exemplary illustrative non-
Iimiting implementation, a controller based on sensor information, the
control is presented. For example, the controller computes the possibility
of control operation based on a set of aircraft parameters and without of
the pilot interruption e‘g‘ throttles, control surface, or any adequate
operation. The aircraft parameters include, but are not limited to, in this
example, acceleration, and speed. The technology herein aims to propose
an aircraft control system and
'a
method of adding.negative speed with
Stability characteristics ,where the aircraft is set to the a particular
configuration , i.e. when the throttle lever is set to the increasing position,
with vibration detected in aircraft, with requiring use of sensors ,GPS
information. The effort of reducing the aircraft speed during maneuvering
can be extremely reduced by the usage of a output of controller . Since the
illustrative reconfigured control system for speed, in addition to manual air
control capability, an auto control process is performed similarly to a
conventional aircraft: the pilot not will be required to keep the control of
aircraft control surface position in order to reduce the aircraft speed which
reduces significantly the pilot workload. Once the target speed is reached,
and vibration indication will grow, the pilot can decide based on this
current speed reference value by selecting appropriate selection. As long
as the Switch is selected, the reference control is continuously
synchronised to the current airspeed. When the switch is selected, the current airspeed is latched as a new control. The disengagement of the
speed acceleration is indicated as a flag also in the primary display. A non-
Limiting advantage of the illustrative solution is a control law that provides
Suitable handling qualities during aircraft vibration condition. This
eliminates the failure case of using erroneous information during training
phase of pilot and allows the control of the aircraft without reduction of
safety margins. In one example non-limiting implementation, additional
hardware or physical parts are needed to implement the proposed solution
when compared to the aircraft in the basic configuration. An example non-
Limiting illustrative system provides a flight control safety mode and
method that provides aircraft speed control through the usage of a switch
and controller. Configured for manoeuvring, the engagement of the
proposed mode adds positive reduction in speed. A way to the flight
control system detects that the aircraft is configured for safe speed.
However, any other sensor used in aeronautical industry could be used to
detect the flight phase, for instance, but not limited to, sensors. A way to
the pilot to change the aircraft speed when positive speed Stability is
engaged. The pilot interceptor may be any of a plurality of devices used in
aeronautics industry to serve aircraft interface with a human pilot, e.g.
throttle, or relevant operation. Till the max allowed speed is reached a on off
switch with controller operates to select the control operation. A mean
of processing data and computing outputs, based on a determined logic,
and commanding the control Surfaces and control operation. A mean of
commanding the control surface according to the command given by this
mean of processing data and computing outputs. A set of sensors which
senses the flying parameters of the aircraft and the position of aircraft .
to be used in a logic module that decides if the auto control mode is to be
engaged. Once control engaged, a set of operations which controls the
control of the flight vehicle is used in a logic module.

CLAIMS
We Claim
1. A method & apparatus for real time structural integrity monitoring &
control of an aircraft which may deteriorate due to corrosion ,crack, loose
interconnect and during development phase; wherein said method uses a
10 novel electrical interconnect (Ref Fig1) called as 3 aircraft structural
integrity control [200] wherein, the novel integrity controller[200] contains,
input section named as multi input synchronised data acquisition section
containing dedicated interface as GPS input interface [103] for providing
real time GPS data ,aircraft speed input [104] for providing real time
15 aircraft speed data, vibration sensors input[107] for providing real time
aircraft acceleration data, throttle position[105] for current throttle position
,strain gauges [106] for providing real time strain data ,output section
namely real time strain boundary indication[110] for indicating higher
boundary of strain values, real time vibration boundary indication[109] for
20 showing appropriate vibration values and output section as a control
signal [108] to control aircraft speed in case auto mode ,LCD pane|[112]
for display of aircraft speed & attitude parameter with sensor values where
it has reached to boundary condition to avoid repetition in same sortie.
25 2. The method and apparatus as said in claim 1, wherein Novel
integrity controller [200] receives 28VDC from 28VDC source [100] from
existing 28VDC source of an aircraft and power supply to the controller
[200] is by using a dedicated switch named as auto/man selection switch indication of auto man selection is presented by auto man selection
indication [102] to confirm that auto mod is selected during speed control.
3. The method and apparatus as said in claim 1, wherein Novel
integrity controller [200] ,Micro controller [111] has nine logical blocks(ref
fig 2),first logical block is vibration database database[701] which is a
memory unit which stores vibration, speed , location profile data indicative
of characteristics of vibration at various locations, Second logical block is
strain database [702] which is memory unit which stores strain and ,spiced
and location profile data indicative of characteristics of the strain at various
locations, third logical block is algorithm 2 block[703],fourth logical block is
data processing block [704], fifth logical block is indication
module[705],sixth logical block is algorithm 1 block[706],seventh logical
block is holding circuit[707],eighth logical block is input/output module[708]
,ninth logical block is GPS and speed data handling module[709].
4. The method and apparatus as said in claim 1, wherein, the novel
controller[200] has a data processing block[704] ,Input/output module[708]
which has multiple input line and output line holding circuit[707] which
gives control signal[108],the indication module[705] further includes an
indication for real time exigency of vibration and strain values, the logical
block named real time data processing block[704] is dedicated for
'
processing GPS data [103], aircraft speed input[104], data processing
block [704] performs combines aircraft speed and attitude information,
This all inputs~are handled by input/output module[708], inputs are stored
in store speed [301], store current GPS data [302], GPS values stored in
GPS data [302], output of data processing block [704] to fed to algorithm 2
block [703] and algorithm 1block [706].
5.
'
The method and apparatus as said in claim 1, wherein the logical
block stain database [702] has profiling data as shown in fig 9, a set of
various points on aircraft mathematically maximum permissible strain
values, speed and attitude ,each sensor on various location on aircraft is
mapped with array of speed and attitude, This set of possible data base
not varying with time ,resulting data items after processing and appropriate
unit conversion are sent to algorithm 1 block[706], performs data
processing, algorithm for algorithm 1 block[706] thus process data for the
plurality of inputs, the Algorithm 1 block[706] generates signal for control
10 signal[108] processed by a relay R1 which ensure that appropriate level is
maintained for varying time duration ,the output of real time data
processing block[704] generates various signal as an indication from pilot
to do monitor visually and will indicate Pilot to bring attitude to neutral
position to avoid further strain.
15
6. The method and apparatus as said in claim 1, wherein the logical
block vibration database [701] is having profiling data as shown in fig 10,
has a set of various points on aircraft, mathematically maximum
permissible vibration values, speed and attitude ,Each point of sensor
20 location is mapped with array of speed and attitude ,This set of possible
data base not varying with time, data items after processing sent to
algorithm 2 block[703], Algorithm for safe vibration computation[703] thus
process data for the plurality of inputs, the Algorithm 2 block[703]
generate signal for control signal[108] ,handled by relay R1 which ensure
25 that appropriate level is maintained for varying time duration ,The output of
real time data processing block [704] generate various signal as an
indication from pilot to do monitor visually and will indicate Pilot to bring
attitude to neutral position to avoid further vibration ,microcontroller
[111]|ogica| block GPS and speed data hollding circuit [709] is also act as input to perform computation of current position of aircraft after
interpolating real time GPS data [103], In the preferred embodiment,
based on resulting data items control signal[108] is generated ,This control
signal is conditioned at holding circuit [707] and made available to further
system via input/output module [708], The sensors data received via
input/output [708] block, The control signal for control signal holding circuit
[707] is further used by relay circuitry R1,The relay R1, switching circuit
works on availability of pulse generated from OR gate D1, which is in turn
based on control signal [108]. couples in parallel with all possible
momentary signal for holding signal from vibration sensor [107] , a
processing of the signals in auto controller [200], Controller [200]
computes potential control level on the basis of the sensed : integrity
control throggh automatic manipulation of a throttle lever irrespective of
any position selected by pilot ,Here throttle position is passed through
relay R1 and once algorithm in microcontroller detects that auto mode is
selected and possible to apply control; throttle position passed through
relay is cut off automatically, the throttle to momentary OFF condition
switch in OFF condition, based on algorithm 1 block [706] gives signal(ref
fig 2) to indication module [709] and indication of strain boundary [110]to
generate assist indication which needs Pilot intervention which in turn
connected with a plurality.
7. The method and apparatus as said in claim 1 (Ref Fig.7), wherein
The input & output module[708] acquires Real time raw vibration values
from sensor 1[501] provides vibration from sensor, likewise it handles
vibration up to Nth Sensor ,The block algorithm 2 block[ 703] carry out fast
Fourier transform using module FFT of Raw vibration values for FFT block
1[502] for first sensor likewise in plurality for Nth sensor ,The data is fed to
block compute gZ/Hz values for compute 1[503] and likewise for Nth sensor ,The output is fed to comparator 1[504] which compares from
vibration database [701] and further gives output to control signal holding
circuit for 1[504] likewise for Nth signal, The signal is be high in variable
length till vibration levels reaches to safe values The output is passed to_
block indication module [705] which has component indication [109] for
indication to Pilot and generation of control signal [108] for onward
circuitry, same is also given to LCD panel [112] for permanent indication to
Pilot for unsafe values for no repetition of same profile .
8. The method and apparatus as said in claim 1, wherein (Fig.8)
shows a processing of the signals in auto controller [200], the input &
output module[708] acquires Real time raw strain values from sensor
15
20
25
1[601] provides strain from sensor, likewise it handlés strain up to Nth
Sensor ,The block algorithm 1 block[706] carry out engineering conversion
using module engineering conversion 1[602] for first sensor likewise in
plurality for Nth sensor, output is fed to comparator 1[604] which compares
from vibration database [701] and further gives output to control signal
holding circuit for 1[504] likewise for Nth signal, The signal is be high in
variable length till strain levels reaches to safe values The output is
passed to block indication module [705] which has component indication
[109] for indication , same is also given to LCD panel [112] for permanent
indication to Pilot for unsafe values for no repetition of same profile .