Note: Descriptions are shown in the official language in which they were submitted.
6 3 8
WO94/15~2 PCT/GB94/00016
AERODYNAMIC PRESSURE SENSOR SYSTEMS
The present invention relates to aerodynamic pressure
sensor systems for sensing aerodynamic pressures applied
to a component movable relative to a surrounding gaseous
medium and is particularly although not exclusively
concerned with sensor systems for use on control surface
components of an aircraft such as wing and fin structures
- where the aerodynamic pressures thereon vary over the
surface in response to predetermined variations in one or
more aerodynamic states of the component in relation to
the medium.
One such aerodynamic state is the airspeed of the
aircraft which is required for all-aircraft. The current
practice is to use pitot heads for this purpose.
Other aerodynamic states are angle of incidence, ~, and
its time rate of change, ~, which are also required for
more sophisticated aircraft, and which are obtained by
means of separate systems eg, vanes.
Pitot heads and vanes both cause drag, are easily
damaged, require duplication for redundancy and are
relatively difficult to maintain and integrate with on-
board computers for data management.
.
A pitot head is a hollow tube that projects forward into
the incident airflow and measures the total pressure of
the airflow. Because it protudes from the surface it
causes additional drag and is liable to damage,
particularly on the ground. It is relatively cumbersome
to integrate with an overall airdata system, and it
reguires rigorous maintenance and a separate installation
for redundancy.
W094/15~2 PCTIG~94/00016
;'t)'`'~ "~
A separate sensing system consisting of a vane mounted on ~
a protruding rod measures the aircraft angle of :
incidence, ~, and its time rate of change, a. It is
however also prone to accidental damage and causes ~
additional drag. ::
The use of Fly-by-Wire (FBW) technology in most all new
aircraft designs requires information on ~ and ~, as well ~
as other information, such as sideslip ang~e. Also in :
increasing demand is information on the aircraft
attitude. Furthermore, there is at present no way to ~:
determine in-flight wing deformation due to aeroelastic
effects and there is no on-board information on aircraft
trimmability.
Furthermore, at Mach numbers above 4 (hypersonic flight)
the pitot heads and vane systems become impracticable to
use.
For incomprPssible flow, which can be assumed for low
sub-sonic flight, airspeed measurement makes use of the
classical Bernouli's equation which relates total
pressure, P~, and local static pressure, Ps~ to the free
stream velocity V~ as follows:
PT P5 ~ P~
where p~ is the free stream air density, or
PT ~ PS = ~ POVE
where VE is the equiYalent airspeed and p is the air
density at sea level
For compressible flow, which n~eds to be assumed for high
sub-sonic and supersonic flight, the pressure-velocity
WO94/15~2 2 131 6 3 ~ PCTIGB94/00016
relationship in Euler form is:
~ Ve2 = a2~ [( ) ]
where a~ is the free stream speed of sound, and
~ is the adiabatic index of air
Thus, the equivalent airspeed VE is derived as a function
of the total pressure PT and the static pressure Ps~ For
current systems, PT is measured by a pitot head on the
aircraft and Ps by a static source on the aircraft. PT
and Ps may be in error due to local variations in flow
around the aircraft. This error, the so called Position
Error, is derived by flight test calibration.
Other ~orrections need to be made which include
instrument and lag errors, and a compressibility
correction is required, in order to obtain the corrected
equivalent airspeed. Currently, such corrections are
applied manually by the pilot.
It is an object of the present invention to provide an
improved aerodynamic pressure sensor system for an
aircraft for determining an aerodynamic state of the
aircraft.
Acc~rding to a first aspect of the present invention
there is provided a component movable relative to a
surrounding gaseous medium to produce over a
predetermined face region thereof aerodynamic pressures
thereon which vary in a predetermined manner in response
to predeterminéd variations in an aerodyna~ic state or
WO94/15~2 PCT/GB94/00016
:~,i3~63~ ~'
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states of the component in relation to the medium,
wherein the component includes an aerodynamic pressure
sensor system comprising a sensor array of pressure
sensitive elements which occupy predetermined locations '~,'
in the face region and each of which generates an output
signal representative of the aerodynamic pressure at the
location and signal generating means responsive to the ''
output signals to generate by reference to the
aerodynamic pressures which the output signals represent ~-
a condition signal representing variations in the
aerodynamic state or one or more of the aerodynamic
states of the component in relation to the medium. ;
In an embodiment of the invention hereinafter to be
described, the component has a leading edge profile
formed by a frontal profile surface which extends
outwardly and rearwardly from a predetermined reference
plane and the predetermined face region occupied by the
sensor array lies in the frontal profile surface.,
Preferably, the sensor array so extends over the frontal
prsfile surface as to provide pressure sensitive elements
on the frontal pr~file surface on each side of the
reference plane.
It is a characteristic of control surface components,
such as wing and fin structures of an aircraft, that
stagnation pressures are developed on the components and
that the positions of these pressures are subject to
variation over the surface of the component in response
to variations in one or more-of the aerodynamic states of
the aircraft or component in relation to the medium.
Tn ~iew of the above characteristic, in an embodiment of
the invention hereinafter to be described, the component
has a frontal profile surface which is such as to create
~ . .
W094/15~2 PCT/GB94/00016
21'.3163~
on the surface an aerodynamic stagnation pressure, the
position of which is.subject to variation over the
surface in response to variations in the aerodynamic
state or states of the component in relation to the
medium and the sensor array is arranged so to extend over
the frontal profile surface that the pressure sensitive
elements are responsive to variations in pressures
arising from variations in the position of the stagnation
pressure.
In an embodiment of the invention hereinafter to be
described the pressure sensitive elements of the array
are so located as to form a column of pressure sensitive
elements in the face region. Preferably, the pressure
sensitive elements of the array are such as to form one
or more further columns of pressure sensitive elements in
the face region.
In an embodiment of the invention hereinafter to be
described each element of the first column of elements
forms with each corresponding element of the other column
or each of the other co~umns of elements a row of
elements extending over the face region in a dire~tion
transverse ~o the columns of elements. Preferably, the
sensor array comprises pressure sensitive elements so
arranged as to form a multiplicity of columns of elements
and a multiplicity of rows of elements. The columns may
conveniently be juxtaposed in the array and the elements
juxtaposed in the or each column~
In an embodiment of the invention hereinafter to be
described the sensor array is so positioned over the
frontal profile surface that the pressure sensitive
elements of the or each column of elements extend over
the frontal profile surface on each side of the reference
WO94/15~2 pcTlGs94loool6
3 ~1 3 `~
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plane.
In an embodiment of the in~ention in its simplest form
the signal generating means generates in response to the ~-
output signals a condition signal representing the speed
of the component relative to the medium by reference to
the aerodynamic pressures over the frontal profile
surface represented by the output signals.
In an embodiment of the invention hereinafter to be
described the component has a component profile which
includes the leading edge profile and a trailing edge
profile, the predetermined reference plane passes through ~-
the centres of curvature of the leading edge profile and
the trailing edge profile and the component profiIe is
such as to generate lift for predetermined angles of
incidence of the reference plane to the direction of
advance movement of the component with respect to the
medium. -
In an embodiment of the invention hereinafter to be
described the component is a main supporting surface
component such as a wing structure and the signal
generating means generates in response to the output
signals from the pressure sensitive elements an incidence
signal representing the angle of incidence of the
component as measured between the reference plane and the
direction of advance movement of the component relative
to the medium.
In addition, the signal generating means may be arranged
to generate in response to the output si~nals an angle of ;
incidence rate signal representing a time rate of change
of the angle of incidence of the component with respect
to thë medium.
WO94/15~2 2 ~ 3 ~ S 3 ~ PCT/GB94/00016
The component may alternatively be a vertically arranged
aircraft control surface component providing directional
stability of the aircraft, such as a vertical tail fin,
the component having a component profile which includes
the leading edge profile and a trailing edge profile,
with the predetermined reference plane passing through
the centres of curvature of the leading edge profile and
the trailing edge profile and the component profile being
such as to generate stabilising side thrust for
predetermined angles of incidence of the reference plane
to the direction of advance movement of the component
with respect to the medium.
The signal generating means may then be arranged to
generate in response to the output signals an angle of
sideslip signal representing the angle of sideslip of the
component as measured between the reference plane and the
direction of advance movement of the component relative
to the medium.
In the embodiment of the invention hereinafter to be
described the component is an elongate component wh-ich
extends in a direction transverse to the direction of
advance movement of the component relative to the medium
and which is so shaped as to produce under predetermined
load conditions and during relative advance movement with
respect to th~ medium angles of incidence which vary
along the transverse~direction. The sensor array may
then comprise a first sensor sub-array which extends over
the frontal profile surface of the component at a .
predetermined first location thereof and a second sensor
sub-array which extends over the frontal profile surface
at a location spaced in the transverse direction from the
~irst predetermined location and the signal generating
means ma~ then generate in response to the output signals
WO 94115832 ~ 3 1 53 8 ~CT/GB94/00016 .
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from the pressure sensitive elements of the two sub-
arrays incidence signals representing the a~gle ~f
incidence of the component at each of the locations,
whereby an angle of twist signal can be generated to
represent the angle of twist of the component between the
two locations.
Where the component is an aircraft main supporting
surface or control surface component cantilevered from -~
the aircraft body, the first sensor sub-array may be
located in a root region of the component and the second
sub-array may be located in a tip reg~on of the ~-
component.
According to a second aspect of the present invention
there is provided an aircraft including one or more
components according to the first aspect of the
invention.
In an embodiment of the invention according to its second
aspect and as hereinafter to be described two of the --
components according to the first aspect of the invention
are port and starboard wing structures and another of the
components according to the first aspect of the invention
is a vertically arranged tail fin structure.
The pressure sensitive elements may conveniently take the
form of magnetic tapes for accurately sensing aerodynamic
pressures or pressure sensitive piezo elecric cells
arranged to form the sensor array.
In the embodiments of the invention hereinafter to be
described the pressure sensitive elements are flush
mounted in the component and as a consequence eliminate
the drag which would arise from the ~se of pitot-head and
W094t1s~2 2131 6 3 ~ PCTIGB94/00016
~-vanes.
Embodiments of the invention will now be described by way ~-
of example with reference to the accompanying drawings in
which:-
Fig l is a schematic perspective view of an aircraft with
control surface wing and fin structures embodying
aerodynamic pressure sensor arrays according to the
invention
Fig 2 is a profile section of the port wing structure of
the aircraft shown in Fig l, taken on the line II-II in
Fig 1 and showing one of the sensor arrays.
Fig 3 is a scrap perspective view drawn to an enlarged
scale of the port wing struc~ure of the aircraft shown in
Fig l, illustrating the sensor array shown in Fig 2
Fig 4 is a schematic graphical representation
illustrating ~he change of local pressure coefficient
over the leading edge profilP of the port wing structure
of the aircraft illustrated in Fig l as measured by the
array of pressure sensitive elements provided in the :
leading edge profile,
Fig 5 is a profile section of the vertical tail fin
structure of the aircraft shown in Fig l, taken on the
line V-V in Fig l
Fig 6 is a scrap perspectiYe view drawn to an enlarged
scale of the vertical tail ~in structure of the aircraft
shown in Fig l, illustrating one of the sensor arrays
embodied in the structure
WO94/15~2 PCT/GB94100016
,~ L~
'.
-- 10
Fig 7 is a flowchart of an airdata handling system for
operating on outputs from the pressure sensitive elements
of sensor arrays provided on the aircraft illustrated in
Fig l to produce output displays or output signals
representing principal aerodynamic states of the aircraft
or of its control surface components.
~eferring first to Fig l, an aircraft ll includes a -
fuselage body 12, port and starboard main supporting
surface wing structures 13 and 14 with engines 15 and 16
and a tail unit 17 including a vertically arranged
control surface fin structure 18 and port and starboard -
control surface elevator structures l9 and 20.
The port wing structure 13 is shown in section in Fig 2.
It has an aerofoil profile 21 and includes a leading edge ;
profilé 22 and a trailing edge profile 23. A chord line
24 is shown in chain dot line. It is as conventionally
defined the straight line through the centres of
curvature of the leading and trailing edge profiles 23
and is the reference line from which angles of incidence
~ are measured.
The leading edge profile 22 of the wing structure 13
includes a sensor array 25 which as best seen in Fig 3
comprises pressure sensitive elements 26 arranged in a
multiplicity of columns 27 and rows 28. The columns 27
of elements 26 extend as shown over the leading edge
profile 22 on each-side of the chord line 24, with the
columns 27 extending further over the leading edge
profile on the underside of the wing structure than on
the upper side of the structure. The rows 28 of elements
26 extend spanwise along the leading edge profile 22 as
shown.
.
. . ~
WO 9411~2 2131 6 3 ~ PCT/GB94100016
The pressure sensitive elements 26 of the array 25 are,
as shown, flush mounted in the leading edge profile 22 of
the wing structure 13 and comprise magnetic tape sensors
which can acurately produce in association with an output
circui~ signals representative of the local aerodynamic
pressures applied to the element. The elements 26 may,
if desired, be protected by retractable shielding to
guard against external impacts on the ground and heating
arrangements may also be provided to protect the elements
from the problems of icing. :
.
The pressure sensitive eIements 26 of the sensor array 25
occupy juxtaposed locations on the leading edge profile
22 in both the chordwise and spanwise directions of the
profile. As a result, they become subject during flight
of the aircraft ll to local aerodynamic pressures which
have magnitudes which are dependent on their location in
the region covered by the array. Furthermore, the local
aerodynamic pressures vary in dependence upon the angle
of incidence of the wing structure during flight.
Such variations in aerQdynamic pressure over the region
covered by the sensor array 25 is graphically represented
in Fig 4 in which the local pressure coefficient at each
pressure sensitive element location 26' is plotted for
successive columns 27 of the elements 26. As will be
seen, the maximum local pressure coefficient occurs at a
predetermined stagnation point for each column 27 of the
array 25 at a position on the leading edge profile
between two of the pressure sensitive elements and that
the pressure coefficient falls off on each side of the
stagnation point progressively over the upper and lower
surfaces of the leading edge profile.
In addition to the variation in the local pressure
WO94/1~2 PCTIGB94100016
~3~S3 - 12 -
coefficients over the leading edge profile covered by the
sensor array 25, the local pressure coefficients also
vary over the leading edge profile as the angle of
incidence of the wing structure change during flight,
insofar as the stagnation pressure changes its location
and provides a maxiumum aerodynamic pressure at another
or other locations of the elements 26 of the array 25.
Thus, the pressure sensitive elements 26 can be arranged
to generate output signals representative of the local -
aerodynamic pressures at the locations which they occupy
on the leading edge profile 22 to provide pressure :
distribution information which includes the stagnation
pressure at the stagnation point and which can be
utilised to generate as hereinafter to be described :,
outputs representative of the indicated or equivalent
airspeed, the angle of incidence of the wing structure
and if desired the time rate of change of the incidence
angle.
Static pressures are sensed at conventional static ports
by sensors producing electrical output signals.
The port wing structure 13 includes a further sensor
array 29 in the tip section of the structure 13, which is
composed of pressure sensitive elements in columns and
rows in the same manner as the sensor array 25, with the
columns and rows extending chordwise and spanwise over
the leading edge profile 22 in the tip region in the same
manner and by the same amount as the columns and rows 27
and 28 of the array 25, the pressure sensitive elements
providing output si~nals representat~ve of the
aerodyn~mic pressures at the locations of the elements in
the same manner as the elements 26 of the array 25.
WO94/15~2 21 31 6 3 ~ PCT1GB94100016
- 13 - ;
The output signals from the pressure sensitive elements
26 of the two arrays 25 and 29 can be used simply to
generate an output representing the indicated airspeed of
the aircraft for example by averaging the outputs from
the two arrays. Additionally, the output signals can be
used to produce an output representing the angle of .
incidence of the wing structure 13 also by averaging the ~:
output signals from the two arrays. More importantly,
however, the output signals can with advantage be used to
generate an output representing the angle of twist of the :~
wing structure 13 as measured by the difference between
the angle of incidence at the array 25 and the angle of
incidence at the array 29.
Th~ starboard wing structure 14 also includes two further
spaced sensor arrays one of which (not shown) is arranged
in the root section of the wing structure 14 at a
position corresponding to that of the array 25 on the
wing structure 13 and an array 30 located at the tip
section of the wing structure 14 and corresponding to the
array 29 on the tip section of the structure 13.
The sensor arrays provided on the wing structures 13 and
14 may simply provide for accurate measurement of
indicated airspeed and where desired the incidence angle
and, if also desired, the time rate of change of the
incidence angle and the angles of twist of the wing
structures.
It is however considered to be of advantage to provide :
also an output representative of the angle of sideslip of
the aircraft relative to its direction of movement
through the air and for this purpose further sensor
arrays are provided on the vertical tail fin structure 18
of the aircraft, as now to be described.
WO94/l5~2 PCTIGB94/00016
3~63~ ~
- 14 -
As will be seen from Fig 1, a sensor array 31 is embodied .
in the fin structure 18 in the root region thereof and a ~-
further sensor array 32 is embodied in the structure 18
at the tip section of the structure. -::
: ,:
The fin structure 18 is shown in section in Fig 5. It
has a low drag profile 33 and includes a leading edge
profile 34 and a trailing edge profile 35. A chord line
36 is shown in chain-dot line. It is, as conventionally
defined, the straight line through the centres of
curvature of the leading and trailing edge profiles 34
and 35 and is the reference line from which angles of
sideslip are measured. The low drag profile 33 differs
from the aerofoil profile 21 shown in Fig 2 insofar as it
is symmetrical with respect to the chord line 36.
The sensor array 31 including in the leading edge profile
34 of the fin structure 18 is best seen in Fig 6. It
comprises pressure sensitive elements 26 arranged in a
multiplicity of columns 27 a~d rows 28. The columns 27
extend as shown over the leading edge profile 34 on each
side of the chord line_36 and by equal amounts on each
side of the fin structure 18. The rows 28 extend as
shown spanwisz along the leading edge profile.
The pressure sensitive elements 26 of the array 31 are,
as shown, flush mounted in the leading edge profile 34
and conveniently comprise magnetic tape sensors which can
accurately produce in association with an output circuit
signals representative of the local aerodynamic pressures
applied to the element. As proposed for the sensor
arrays provided on the wing structures 13 and 14, the
elements 26 of the array 31 may also if desired be
protected by retractable shielding to guard against
external impacts on the ground and heating arrangements
WO 94/15832 PCTIGB94tO0016
2131~38
may also be provided to protect the elements from the ,
problems of icing.
The pressure sensitive elements 26 of the sensor array 31
occupy juxtaposed locations on the leading edge profile
34 of the fin structure 18 both in the chordwise and
vertical direction of the profile. As a result, they :
become subject during flight of the aircraft to local
aerodynamic pressures which have magnitudes which are
dependent upon the location of the element in the region
covered by the array. Furthermore, these local -
aerodynamic pressures vary in dependence upon the angle
of sideslip of the fin structure 18 during flight.
Such variations in local aerodynamic pressure over the
region covered by the sensor array 31 corresponds closely
to that for the array 25 as graphically represented in
Fig 4. Again, the maximum local pressure coefficient
occurs at a predetermined stagnation point for each
column 27 of the array 31, to each side of which the
aerodynamic pressure falls off progressively over port
and starboard surfaces_of the leading edge profile 34.
In addition to the variation in local aerodynamic
pressure over the leading edge profile covered by the
sensor array 31, the local aerodynamic pressures also
vary over the leading edge profile as the angle of -
sideslip of the fin structure 18 change5 during flight,
insofar that the stagnation pressure changes its location
and provides a maximum local aerodynamic pressure at
ano~her or other locations of the elements 26 of the
array 31 in dependence upon the angle of sideslip of the
aircraft.
The pressure sensiti~e elements 26 of the array 31 are
WO94/15~2 pcTlGs94loool6
?.~3~3~
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arranged to generate output signals representative of the
local aerodynamic pressures at the locations which they
occupy on the leading edge profile 34 of the fin
structure 18 to provide pressure distribution information
which identifies the stagnation pressure at the
stagnation point on the leading edge profile and which
can then be utilised to generate an output representative
of the angle of sideslip of the f in structure 18 with
respect to the direction of movement of the aircraft
through the air. The information can of course also if
desired by used alternatively or additionally to generate
outputs representative of the indicated or equivalent
airspeed of the aircraft.
The fin structure 18 includes a further sensor array 32
in the tip section of the structure, which is composed of ~;-
pressure sensitive elements in columns and rows in the
same manner as the sensor array 31, with the columns and
rows extendin~ chordwise and vertically over the leading
edge profile 34 in the tip section in the same manner and
as the columns 27 and rows 28 of the array 31, the
pressure sensitive elements providing output signals
representative of the aerodynamic pressures at the
locations of these elements in the same manner as the
elements 26 of the array 31.
The output signals from the pressure sensiti~e elements
of the two arrays 31 and 32 can be used simply to
generate an output representing the indicated or
equivalent airspeed of the aircraft. In particular,
however, the output siqnals are used to produce an output
representing the angle of sideslip of the fin structure
18 by averaging the output signals from the two arrays.
In addition, the output signals can be used in special
circumstances which demand it to generate an output
W094115~2 PCT/GB94/00016
~ l 31~ ~8
representing the angle of twist as measured by the
difference between the angle of sideslip as measured at
the array 31 and the angle of sideslip as measured at the
array 32.
The positions of the sensor arrays provided on the wing
structures 13 and 14 and the fin structure 18 are
carefully selected so that the elements 26 are responsive
to local variations of pressure near to and including the
stagnation point. The output signals generatçd by the
pressure sensitive elements then provide aerodynamic
pressure distribution information which can be processed
by an air-data system to produce outputs representative
of one or more selected aerodynamic states of the
aircraft, such as airspeed, incidence, sideslip angle and
wing twist and along with data from inertial-navigation
and global position systems, attitude angles and ground
speed. -
A typical generalised air-data handling system is .
schematically illustrated in Fig 7 in flowchart form.
Inputs transmitted to ~nd outputs produced by the system
are represented by abbreviations which are conventionally
accepted but which are for convenience set out in the
following table of symbols:- -
Svmbols
IAS Indicated Airspeed
CAS Calibrated Airspeed
EAS Equivalent Airspeed
TAS True Airspeed :~
a Angle of Incidence
a Time Rate of change of angle of incidence
B Sideslip angle
INS Inertial Navigational System :
WO94/15~2 PCT/GB94/00016
~,~3~G~ :
- 18 -
GPS Global Positioning System
EFCS Electronic Flight Control 5ystem
In the flowchart illustrated in Fig 7, a central
processor 37 is provided with inputs 38 to 4l and
produces outputs 42 to 44. The inputs 38 include the
output signals from the arrays provided on the wing
structures 13 and 14 and the fin structure 18, which are
computed in the processor 37 to generate total
aerodynamic pressures PT. Input signals from one or more
static ports which are provided by sensors producing ~;
electrical signal outputs represent the static pressure '
Ps~ The inputs 39 comprise data stored in memory and
representing position error, instrument errors,
compressibility and air density changes. The inputs 40
comprise stagnation pressure position values for given
angles of inclination a and sideslip angles ~. Input 4l
provides information as to the fuel used.
As to the outputs, output 42 includes displays of
(PT-PS)' CAS, IAS, EAS, and TAS. The outputs 43, also in
display form, comprise ~, ~ and ~ computed by the
processor 37 from the inputs 38 and stored data 40. The
outputs 44 include wing deformation presented as an angle
of twist and is computed from the inputs to the
processor. The outputs 44 further include ground speed,
wind speed and aircraft attitude also computed by the
processor 37 from the inputs to the processor as well as
data provided by the Inertial Navigation and Global
Position systems 45, which are also supplied with output
from the processor 37 for application to the electronic
flight control system 46.
The sensor arrays are deployed on the aircraft to measure
aerodynamic surface pressure distribution over small
WO94/1~2 PCTIGB94/00016
213163~
- 19 -
regions in several carefully selected aircraft locations,
and to feed their output signals to the processor 37
already stored with the necessary correction factors and
geometric details required to compute accurate speeds
(IAS/CAS/EAS/TAS), a, a, sideslip angle, and aircraft
attitude. The final choice of parameters presented
depends on the degree of sophistication desired in the
system. The invention furthermore eliminates surface
protrusion and combines speed and incidence measuring
systems into one integrated new system.
By comparing with stored data on stagnation position
obtained by flight test calibration the local overall
aircraft incidence can be determined. Also by comparing
relative values of incidence at the wing-tip and wing-
root arrays with corresponding calibrated data for the
unloaded wing, the degree of structural wing twist can be
determined.
The sideslip angle of the a~rcraft can be derived in a
similar manner from the position of the stagnation point
on the fin structure 1~ derived from the appropriate
sensor arrays. The use of a number of columns of -
elements for the sensor array in each region, each giving -~
the stagnation pressure and position, provides a more
accurate average value of the quantities, and also
ensures that failure of one or more columns will not
affect the system adversely. -~
It is also to be understood that the sensor arrays can be -
installed at any place on the surface where local flow
field information is required.
It is estimated that the total weight of the sensor -~
gystem (sensor array and electric cable) is less than
WO94115~2 PCT/GB94/00016
?.~3'S~'~
- 20 -
that of a conventional pitot-head static-porl system for
the reasons (i) that the relatively heavy pitot-head tube
is replaced by light pressure sensor arrays, and (ii)
that the conventional pressure tubes from the pitot head
and static port systems to the cockpit instruments are
replaced ~y electrical cables, the weight of which may be
decreased further by using multiplex data transmission.
Choices for pressure sensitive elements are the following
given in decreasing order of sensitivity:
(i) Magnetic tape - a relatively new method in
accurately sensing pressure;
(ii) Piezo-electric cells
Other possible choices are strain gauges and vacuum
tubes, but these are unlikely to surpass the capabilities
offered by the elements referred to above.
Stagnation pressures encountered at high subsonic speeds
typi~ally vary from 3.L lb/in2 at sea leve1 (360 knots
airspeed) to 0.4 lb/in2 at 50000ft altitude ~around 0.4
Mach). Away from stagnation, as velocity increases,
there could be about 50% reduction in the level of
pressure head readings. This is well within the range of
capability offered by the system hereinbefore described.
For high performance military aircraft the low end of the
range is of the order of 0.2 lb/in2.
The invention provides a system for obtaining data on an
aircraft in flight for the determination of aerodynamic
states such as speed, incidence, attitude and the like.
Compared with the conventional pitot-tube/~-vane system
it has the following advantages:
WO94/15~2 pcTlGs94loool6
(i) Reduction in drag.
(ii) Less prone to accidental damage
(iii) Performs the functions of pitot-tube and the ~-
vane systems with a single type of system.
(iv) Built-in redundancy provides greater accuracy,
reliability and safety.
(v) Pro~ides aircraft attitude information.
(vi) Provides information on in-flight twist of wing
and fin structures.
(vii) Easier system inte~ration - facilities real- -
time computation of required data.
(viii) Compatible with EFCS and can be integrated with
an inertial navigational system and/or Global
Position System.
(ix~ Could be used to supply data to an in-~light
trimming system to improve cruise efficiency.
:
The flush-mounted sensor arrays are installed at several
chosen locations on an a-rcraft to measure the local
variation of pressure near to and including the
stagnation point. The electrical outputs from the arrays
are processed by the air-data system to give airspeed,
incidence, sideslip angle and wing twist, and~ along with
data from an inertial navigation and Global Position
System to give attitude angles and ground speed.
The use of surface pressure sensor arrays flush-installed
at the leading edges of the lifting surfaces has the
advantage of drag reduction, easier airdata integration,
more detailed cockpit display, improved
accuracy/safety/reliability/redundancy/maintainability,
and possible weight savings.
A great ad~antage of the proposed system is that all
sensor array output signals are electronic, a~d storage
WO94/15~' PCTIGB94/00016
~,~3~G3 ~ ~
- 22 -
of the corrections in on-board computers enables
corrected airspeeds and other aerodynamic states of the
aircraft to be calculated automatically and displayed on
cockpit instruments.
Calibration of the sensor arrays, in order to obtain the
corrections necessary to derive accurate aircraft state
data will be required. Since all calculations are
performed in real-time on on-board computers oth~er
details eg wind speed and ground speed can also be
obtained directly as cockpit displays, given inputs from
an INS or a GPS.
In current practice, while real time computation of
airspeed is possible with on-board computers, the
invention offers a simpler method of integration to the
system by the very nature of having electrical signals at
the source.
