EP0629166A1 - Aerodynamic pressure sensor systems - Google Patents

Aerodynamic pressure sensor systems

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Publication number
EP0629166A1
EP0629166A1 EP94904225A EP94904225A EP0629166A1 EP 0629166 A1 EP0629166 A1 EP 0629166A1 EP 94904225 A EP94904225 A EP 94904225A EP 94904225 A EP94904225 A EP 94904225A EP 0629166 A1 EP0629166 A1 EP 0629166A1
Authority
EP
European Patent Office
Prior art keywords
component
aircraft
aerodynamic
predetermined
elements
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP94904225A
Other languages
German (de)
French (fr)
Inventor
Ajoy Kumar Kundu
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Short Brothers PLC
Original Assignee
Short Brothers PLC
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Short Brothers PLC filed Critical Short Brothers PLC
Publication of EP0629166A1 publication Critical patent/EP0629166A1/en
Withdrawn legal-status Critical Current

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Classifications

    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
    • G01P5/00—Measuring speed of fluids, e.g. of air stream; Measuring speed of bodies relative to fluids, e.g. of ship, of aircraft
    • G01P5/14—Measuring speed of fluids, e.g. of air stream; Measuring speed of bodies relative to fluids, e.g. of ship, of aircraft by measuring differences of pressure in the fluid
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01P—MEASURING LINEAR OR ANGULAR SPEED, ACCELERATION, DECELERATION, OR SHOCK; INDICATING PRESENCE, ABSENCE, OR DIRECTION, OF MOVEMENT
    • G01P13/00—Indicating or recording presence, absence, or direction, of movement
    • G01P13/02—Indicating direction only, e.g. by weather vane
    • G01P13/025—Indicating direction only, e.g. by weather vane indicating air data, i.e. flight variables of an aircraft, e.g. angle of attack, side slip, shear, yaw

Definitions

  • 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.
  • 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 requires rigorous maintenance and a separate installation for redundancy.
  • 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, ⁇ . It is however also prone to accidental damage and causes additional drag.
  • V E is the equivalent airspeed and p. is the air density at sea level
  • a. is the free stream speed of sound
  • y is the adiabatic index of air
  • the equivalent airspeed V E is derived as a function of the total pressure P ⁇ and the static pressure P s .
  • P ⁇ is measured by a pitot head on the aircraft and P s by a static source on the aircraft.
  • P ⁇ and P s 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.
  • 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 predetermined variations in an aerodynamic state or states of the component in relation to the medium
  • 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.
  • 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.
  • the sensor array so extends over the frontal profile surface as to provide pressure sensitive elements on the frontal profile surface on each side of the reference plane.
  • control surface components such as wing and fin structures of an aircraft
  • 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.
  • the component has a frontal profile surface which is such as to create 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.
  • the pressure sensitive elements of the array are so located as to form a column of pressure sensitive elements in the face region.
  • 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.
  • each element of the first column of elements forms with each corresponding element of the other column or each of the other columns of elements a row of elements extending over the face region in a direction transverse to the columns of elements.
  • 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.
  • 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 plane .
  • 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.
  • 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 profile 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.
  • 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.
  • the signal generating means may be arranged to generate in response to the output signals an angle of incidence rate signal representing a time rate of change of the angle of incidence of the component with respect to the medium.
  • 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.
  • the component is an elongate component which 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 the 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 first predetermined location and the signal generating means may then generate in response to the output signals from the pressure sensitive elements of the two sub- arrays incidence signals representing the angle of 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.
  • 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 region of the component.
  • an aircraft including one or more components according to the first aspect of the invention.
  • 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.
  • the pressure sensitive elements are flush mounted in the component and as a consequence eliminate the drag which would arise from the use of pitot-head and ⁇ -vanes .
  • Fig 1 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 1, 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 structure of the aircraft shown in Fig 1, illustrating the sensor array shown in Fig 2
  • Fig 4 is a schematic graphical representation illustrating the change of local pressure coefficient over the leading edge profile 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 1, taken on the line V-V in Fig 1
  • Fig 6 is a scrap perspective' view drawn to an enlarged scale of the vertical tail fin structure of the aircraft shown in Fig 1, illustrating one of the sensor arrays embodied in the structure
  • 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 1 to produce output displays or output signals representing principal aerodynamic states of the aircraft or of its control surface components.
  • an aircraft 11 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 19 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 profile 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.
  • 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 circuit 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 elements 26 of the sensor array 25 occupy juxtaposed locations on the leading edge profile 22 in both the chordwi ⁇ e and spanwise directions of the profile. As a result, they become subject during flight of the aircraft 11 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.
  • Fig 4 Such variations in aerodynamic 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.
  • 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.
  • the local pressure coefficients In addition to the variation in the local pressure 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.
  • 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 signals representative of the aerodynamic pressures at the locations of the elements in the same manner as the elements 26 of the array 25.
  • 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.
  • 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.
  • the 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.
  • 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 and 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 spanwise 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.
  • 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 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.
  • the local aerodynamic pressures also vary over the leading edge profile as the angle of sideslip of the fin structure 18 changes during flight, insofar that the stagnation pressure changes its location and provides a maximum local aerodynamic pressure at anorher or other locations of the elements 26 of the array 31 in dependence upon the angle of sideslip of the aircraft.
  • the pressure sensitive elements 26 of the array 31 are 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 fin 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 extending 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 sensitive 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.
  • the output signals 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.
  • the output signals can be used in special circumstances which demand it to generate an output 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 generated 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.
  • FIG. 7 A typical generalised air-data handling system is schematically illustrated in Fig 7 in flowchart form. Inputs transmitted to and 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:-
  • a central processor 37 is provided with inputs 38 to 41 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 P ⁇ .
  • Input signals from one or more static ports which are provided by sensors producing electrical signal outputs represent the static pressure P s .
  • 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 41 provides information as to the fuel used.
  • output 42 includes displays of (P ⁇ -P s ) , CAS, IAS, EAS, and TAS.
  • the outputs 43 also in display form, comprise a , a 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 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 , 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.
  • 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 aircraft can be derived in a similar manner from the position of the stagnation point on the fin structure 18 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.
  • the sensor arrays can be installed at any place on the surface where local flow field information is required.
  • the total weight of the sensor system is less than that of a conventional pitot-head static-port 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 by electrical cables, the weight of which may be decreased further by using multiplex data transmission.
  • strain gauges and vacuum tubes are unlikely to surpass the capabilities offered by the elements referred to above.
  • Stagnation pressures encountered at high subsonic speeds typically vary from 3.1 lb/in 2 at sea level (360 knots airspeed) to 0.4 lb/in 2 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/in 2 .
  • 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: (i) Reduction in drag.
  • Position System. (ix) could be used to supply data to an in-flight trimming system to improve cruise efficiency.
  • the flush-mounted sensor arrays are installed at several chosen locations on an aircraft 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.
  • a great advantage of the proposed system is that all sensor array output signals are electronic, and storage 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.
  • the invention offers a simpler method of integration to the system by the very nature of having electrical signals at the source.

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Indicating Or Recording The Presence, Absence, Or Direction Of Movement (AREA)
  • Measuring Fluid Pressure (AREA)

Abstract

Elément (13) structurel d'avion qui, en vol, produit sur une zone (22) frontale prédéterminée de l'avion des pressions aérodynamiques qui varient de manière prédéterminée en réponse à des variations prédéterminées d'un ou de plusieurs états aérodynamiques de l'avion, comprenant un système de capteur de pression aérodynamique formé d'un réseau (25) de capteurs constitué d'éléments (26) sensibles à la pression qui occupent des emplacements prédéterminés dans la zone frontale et qui génèrent chacun un signal de sortie représentant la pression aérodynamique au niveau de leur emplacement respectif; et d'un système générant des signaux qui réagit aux signaux de sortie afin de générer, en se fondant sur les pressions aérodynamiques que les signaux de sortie représentent, un signal ou plusieurs signaux de condition représentant les variations de l'état aérodynamique ou d'au moins un des états aérodynamiques de l'avion. Les systèmes de capteurs sont incorporés dans les structures d'ailes et de plan fixe de queue.Structural element (13) of an aircraft which, in flight, produces on a predetermined frontal zone (22) of the aircraft aerodynamic pressures which vary in a predetermined manner in response to predetermined variations of one or more aerodynamic states of the aircraft, comprising an air pressure sensor system formed of an array (25) of sensors consisting of pressure sensitive elements (26) which occupy predetermined locations in the frontal area and which each generate an output signal representing the aerodynamic pressure at their respective location; and a signal generating system responsive to the output signals to generate, based on the aerodynamic pressures that the output signals represent, one or more condition signals representing changes in aerodynamic state or at least one of the aerodynamic states of the aircraft. The sensor systems are incorporated into the wing and tail stabilizer structures.

Description

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, α, and its time rate of change, α, 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 requires rigorous maintenance and a separate installation for redundancy. A separate sensing system consisting of a vane mounted on a protruding rod measures the aircraft angle of incidence, α, and its time rate of change, α. It is however also prone to accidental damage and causes additional drag.
The use of Fly-by-Wire (FB ) technology in most all new aircraft designs requires information on and , as well as other information, such as sideslip angle. 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 incompressible flow, which can be assumed for low sub-sonic flight, airspeed measurement makes use of the classical Bernouli's equation which relates total pressure, Pτ, and local static pressure, Ps, to the free stream velocity Vβ as follows:
p - p = in v2 where pβ is the free stream air density, or
Pτ - Ps = ^PoVE 2,
where VE is the equivalent airspeed and p. is the air density at sea level
For compressible flow, which needs to be assumed for high sub-sonic and supersonic flight, the pressure-velocity relationship in Euler form is :
where a., is the free stream speed of sound, and y is the adiabatic index of air
Thus, the equivalent airspeed VE is derived as a function of the total pressure Pτ and the static pressure Ps. For current systems, Pτ is measured by a pitot head on the aircraft and Ps by a static source on the aircraft. Pτ 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 corrections 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.
According 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 predetermined variations in an aerodynamic state or 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 profile surface as to provide pressure sensitive elements on the frontal profile 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.
In view 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 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 columns of elements a row of elements extending over the face region in a direction transverse to 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 plane .
In an embodiment of the invention 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 profile 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 signals an angle of incidence rate signal representing a time rate of change of the angle of incidence of the component with respect to the medium. 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 which 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 the 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 first predetermined location and the signal generating means may then generate in response to the output signals from the pressure sensitive elements of the two sub- arrays incidence signals representing the angle of 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 region 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 use of pitot-head and α-vanes .
Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which:-
Fig 1 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 1, 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 structure of the aircraft shown in Fig 1, illustrating the sensor array shown in Fig 2
Fig 4 is a schematic graphical representation illustrating the change of local pressure coefficient over the leading edge profile 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 1, taken on the line V-V in Fig 1
Fig 6 is a scrap perspective' view drawn to an enlarged scale of the vertical tail fin structure of the aircraft shown in Fig 1, illustrating one of the sensor arrays embodied in the structure 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 1 to produce output displays or output signals representing principal aerodynamic states of the aircraft or of its control surface components.
Referring first to Fig 1, an aircraft 11 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 19 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 profile 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 α 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. 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 circuit 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 elements 26 of the sensor array 25 occupy juxtaposed locations on the leading edge profile 22 in both the chordwiεe and spanwise directions of the profile. As a result, they become subject during flight of the aircraft 11 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 aerodynamic 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 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 signals representative of the aerodynamic pressures at the locations of the elements in the same manner as the elements 26 of the array 25. 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.
The 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. 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 and 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 spanwise 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 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 changes during flight, insofar that the stagnation pressure changes its location and provides a maximum local aerodynamic pressure at anorher or other locations of the elements 26 of the array 31 in dependence upon the angle of sideslip of the aircraft.
The pressure sensitive elements 26 of the array 31 are 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 fin 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 extending 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 sensitive 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 signals 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 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 generated 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 and 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:-
Symbols
IAS Indicated Airspeed
CAS Calibrated Airspeed
EAS Equivalent Airspeed
TAS True Airspeed a Angle of Incidence Time Rate of change of angle of incidence β Sideslip angle
INS Inertial Navigational System GPS Global Positioning System
EFCS Electronic Flight Control System
In the flowchart illustrated in Fig 7, a central processor 37 is provided with inputs 38 to 41 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 Pτ. 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 β. Input 41 provides information as to the fuel used.
As to the outputs, output 42 includes displays of (Pτ-Ps) , CAS, IAS, EAS, and TAS. The outputs 43, also in display form, comprise a , a and β 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 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 , 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 aircraft can be derived in a similar manner from the position of the stagnation point on the fin structure 18 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 system (sensor array and electric cable) is less than that of a conventional pitot-head static-port 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 by 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 typically vary from 3.1 lb/in2 at sea level (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/α-vane system it has the following advantages: (i) Reduction in drag.
(ii) Less prone to accidental damage
(iii) Performs the functions of pitot-tube and the α- vane systems with a single type of system, (iv) Built-in redundancy provides greater accuracy, reliability and safety, (v) Provides aircraft attitude information, (vi) Provides information on in-flight twist of wing and fin structures, (vii) Easier system integration - 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-flight trimming system to improve cruise efficiency.
The flush-mounted sensor arrays are installed at several chosen locations on an aircraft 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 advantage of the proposed system is that all sensor array output signals are electronic, and storage 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 other 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.

Claims

1. 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 predetermined variations in an aerodynamic state or 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.
2. A component according to claim 1, wherein the component has a leading edge profile formed by a frontal profile surface which extends outwardly and rearwardly from a predetermined reference plane and wherein the predetermined face region occupied by the sensor array lies in the frontal profile surface.
3. A component according to claim 2, wherein the the sensor array so extends over the frontal profile surface as to provide pressure sensitive elements on the frontal profile surface on each side of the reference plane.
4. A component according to claim 2 or 3, wherein the frontal profile surface of the component is such as to create on the surface an aerodynamic stagnation pressure, the position of which is subject to variation over the surface in response to the predetermined variations in the aerodynamic state or states of the component in relation to the medium and wherein the sensor array so extends over the frontal profile surface that the pressure sensitive elements are responsive to the variations in the position of the stagnation pressure.
5. A component according to any of claims 1 to 4 wherein the pressure sensitive element's are so located as to form a column of pressure sensitive elements in the face region.
6. A component according to claim 5, wherein the pressure sensitive elements of the array are so located as to form one or more further columns of pressure sensitive elements in the face region.
7. A component according to claim 6, wherein each element of the first column of elements forms with each corresponding element of the other column or each of the other columns of elements a row of elements extending over the face region in a direction transverse to the columns of elements.
8. A component according to claim 7, wherein 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.
9. A component according to claim 8 as appendent to claim 6 or 7 wherein the columns of elements are juxtaposed in the array.
10. A component according to any of claims 5 to 9 wherein the elements are juxtaposed in the column or each column.
11. A component according to any of claims 5 to 10, wherein 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 plane.
12. A component according to any of claims 2 to 11, wherein 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.
13. A component according to any of claims 2 to 12, wherein the component has a component profile which includes the leading edge profile and a trailing edge profile, wherein the predetermined reference plane passes through the centres of curvature of the leading edge profile and the trailing edge profile and wherein the component profile 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.
14. A component according to claim 13, wherein the component is an aircraft main supporting surface component.
15. A component according to claim 14, wherein 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 an advance movement of the component relative to the medium.
16. A component according to claim 15, wherein the signal generating means generates in response to the output signals an angle of incidence rate signal representing a time rate of change of the angle of incidence of the component with respect to the medium.
17. A component according to any of claims 2 to 12, wherein the component is a vertically arranged aircraft control surface component providing directional stability of the aircraft and having a component profile which includes the leading edge profile and a trailing edge profile, wherein the predetermined reference plane passes through the centres of curvature of the leading edge profile and the trailing edge profile and wherein the component profile is 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.
18. A component according to claim 17, wherein the signal generating means generates in response to the output signals an angle of sideslip signal representing the angle of side slip of the component as measured between the reference plane and the direction of advance movement of the component relative to the medium.
19. A component according to any of claims 13 to 18, wherein the component is an elongate component which 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 provide under predetermined load conditions and during relative advance movement with respect to the medium angles of incidence which vary along the transverse direction, wherein the sensor array comprises 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 first predetermined location and wherein the signal generating means generates in response to the output signals from the pressure sensitive elements of the two sub-arrays incidence signals representing the angle of 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.
20. A component according to claim 19, wherein the component is an aircraft main supporting surface or control surface component to be cantilevered from an aircraft body and wherein the first sub-array is located in a root region of the component and the second sub- array is located in a tip region of the component.
21. An aircraft including one or more components according to any of claims 1 to 20.
22. An aircraft according to claim 21, and including a main supporting' surface component according to claim 14, 15 or 16.
23. An aircraft according to claim 21 and including a vertically arranged control surface component according to claim 17 or 18 .
24. An aircraft according to claim 21 including a main supporting surface component according to claim 14, 15 or 16 and a vertically arranged control surface component according to claim 17 or 18.
25. An aircraft component substantially as hereinbefore described with reference to Figs 1 to 4 or 5 and 6 of the accompanying drawings.
26. An aircraft substantially as hereinbefore described with Figs 1 to 6 and including an air data handling system substantially as hereinbefore described with reference to Fig 7.
EP94904225A 1993-01-08 1994-01-05 Aerodynamic pressure sensor systems Withdrawn EP0629166A1 (en)

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PCT/GB1994/000016 WO1994015832A1 (en) 1993-01-08 1994-01-05 Aerodynamic pressure sensor systems

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Families Citing this family (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB9518800D0 (en) * 1995-09-14 1995-11-15 Gkn Westland Helicopters Ltd Method & apparatus for determining the airspeed of rotary wing aircraft
US5928309A (en) * 1996-02-05 1999-07-27 Korver; Kelvin Navigation/guidance system for a land-based vehicle
US5986547A (en) * 1997-03-03 1999-11-16 Korver; Kelvin Apparatus and method for improving the safety of railroad systems
DE10001813C2 (en) * 2000-01-18 2003-10-30 Eads Deutschland Gmbh Measuring system for determining air data of an aircraft and a method for determining the air data
FR2807737B1 (en) * 2000-04-14 2002-07-12 Aerospatiale Matra Airbus PRESSURE TAKING DEVICE AND ITS IMPLEMENTATION METHOD ON AN AIRCRAFT FUSELAGE PANEL
US6697752B1 (en) 2000-05-19 2004-02-24 K&L Technologies, Inc. System, apparatus and method for testing navigation or guidance equipment
US6452542B1 (en) * 2001-03-02 2002-09-17 Rosemount Aerospace Inc. Integrated flight management system
DE102010052905B4 (en) 2010-12-01 2014-08-28 Deutsches Zentrum für Luft- und Raumfahrt e.V. Sensor device for measuring a direction of flow and evaluation for it
DE102011005355A1 (en) * 2011-03-10 2012-09-13 Airbus Operations Gmbh Component assembly on an aircraft structure and method for installing a component in an aircraft structure
DE102012104366A1 (en) * 2012-05-21 2013-11-21 Deutsches Zentrum für Luft- und Raumfahrt e.V. Anströmrichtungssensoreinrichtung and Anströmrichtungserfassungseinrichtung
FR3021116B1 (en) * 2014-05-13 2018-12-07 Airbus Operations MEASURING SYSTEM FOR MEASURING THE SPEED OF AN AIRCRAFT
FR3074298B1 (en) * 2017-11-30 2020-02-07 Airbus Operations ASSEMBLY COMPRISING AN AERODYNAMIC PROFILE AND A SYSTEM FOR DETERMINING CHARACTERISTICS OF AN INCIDENTAL AIRFLOW ON A LEADING EDGE OF THE AERODYNAMIC PROFILE
US10823753B2 (en) * 2018-12-14 2020-11-03 Rosemount Aerospace Inc. Air data probe with optical pressure integration

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1579638A (en) * 1977-06-24 1980-11-19 Secr Defence Airstream pressure sensing probes
US4350314A (en) * 1980-07-10 1982-09-21 The Board Of Trustees Of Western Michigan University Wing mounted stall condition detector
DE3337978C2 (en) * 1983-10-19 1987-01-15 Deutsche Forschungs- und Versuchsanstalt für Luft- und Raumfahrt e.V., 5000 Köln Device for measuring a pressure and a temporal pressure curve
DE3544144A1 (en) * 1985-12-13 1987-06-25 Eduard Prof Dr Ing Igenbergs METHOD AND DEVICE FOR DETERMINING A FLOW ANGLE ON VEHICLES
US5083279A (en) * 1990-05-09 1992-01-21 Honeywell, Inc. Canard based high angle of attack air data sensor

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of WO9415832A1 *

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WO1994015832A1 (en) 1994-07-21
AU5836594A (en) 1994-08-15
JPH07507144A (en) 1995-08-03
GB2274170A (en) 1994-07-13

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