WO1992001905A1

WO1992001905A1 - Aircraft instrumentation systems

Info

Publication number: WO1992001905A1
Application number: PCT/GB1990/001106
Authority: WO
Inventors: John Richard Hall
Original assignee: The Secretary Of State For Defence In Her Britanic Majesty's Government Of The United Kingdom Of Great Britain And Northern Ireland
Priority date: 1990-07-19
Filing date: 1990-07-19
Publication date: 1992-02-06
Also published as: JPH05509153A; EP0545920A1

Abstract

An aircraft instrumentation system providing a head-up display which displays modified indications of flight path and other parameters, very substantial modifications being made to achieve a display which is useful in all man÷uvres even in an agile fast-jet aircraft. The modifications reduce the transient movements of the display indicia which are due to the increased incidence needed to instigate flight path changes. One modification provides an aircraft symbol constrained to stay on a vertical centreline of the display unit, with a displacement depending on bank angle to preserve horizon correlation as far as possible. A comparatively small symbol (vv diamond) with unlimited motion indicates the flight path direction and a pitch ladder indicates modified climb-dive angle when read against the aircraft symbol. The positions of the aircraft symbol, vv diamond, and pitch ladder are all modified by Q functions derived from pitch attitude or body pitch rate to offset the effect of transient variations of pitch in flight path man÷uvres. The Q functions are calculated by algorithms involving a variable time constant τ which is approximately matched to the time constant of the heave mode response of the aircraft, which is a function of mass, airspeed, wing area, and lift-curve slope coefficient. A linear approximation is used, with an experimental optimisation, to derive an appropriate expression for τ. Peripheral scales showing attitude and optional other parameters are positioned in a fixed relationship to the variable position of the aircraft symbol.

Description

AIRCRAFT INSTRUMENTATION SYSTEMS

This invention relates to aircraft instrumentation systems and particularly to aircraft instrumentation systems including head-up displays. It is now common for aircraft especially military fast jet aircraft to be fitted with head-up displays which provide analogue and/or digital indications of the attitude, flight path and other vital parameters of the aircraft's situation by projecting illuminated indicia on a semi-transparent screen through which the pilot may also see a view ahead.

It is conventional for head-up displays to show the flight path direction by the position of an aircraft symbol, to show bank angle and either pitch or climb/dive angle by a pitch ladder display, and to show various other parameters by peripheral indications.

Early head-up displays showed pitch angles, but it is more useful to show the climb-or-dive angle of the actual flight path, which is particularly necessary and useful for landing guidance. However there are considerable difficulties in providing actual flight path indications on a display suitable for use as the primary instrumentation for all motions of flight, particularly for an agile, aerobatic or combat aircraft.

In order to provide the accuracy required for low-level flight and landing approach guidance a head-up display must show sensitive and accurate indications of actual flight path direction. When an aircraft makes a substantial manoeuvre, to climb, dive, bank, turn or roll the aircraft's pitch and its attitude relative to the direction of its flight path will change very rapidly. Comparatively large changes in the aircraft's attitude will be instigated by the pilot, to initiate the manoeuvre, and the flight path direction as seen in the head-up display may change substantially with vertical movements due to the incidence needed for manoeuvring and lateral movements due to sideslip when rolling. On a head-up display where the aircraft symbol moves to show the actual direction of flight relative to the longitudinal fuselage datum or fore-and-aft direction of the aircraft, these effects cause the symbol to move rapidly and it may go out of the field of view or be stopped at a limiting position on the edge of the field, where it shows only that the flight path direction is beyond a set limit.

In any claim, dive, bank or roll manoeuvre the pitch ladder will move up or down and/or swing round rapidly. In such circumstances digits identifying specific bars of the pitch ladder will often move too rapidly to be read.

In general the rapid motions of all indicia caused by these effects may drastically reduce the effectiveness of the display both for promoting the pilot's spatial awareness and for assisting any precise control or adjustment, and they can be very disconcerting and may tend to promote a state of disorientation in which the pilot may misinterpret his position. This can be highly dangerous.

Any discontinuity or sudden jump or apparent disappearance of a vital indication may cause the pilot to doubt the reliability of the instrumentation. It is therefore very important to avoid as far as possible any conditions which could cause sudden jumps in the position of any indicia. Such conditions may arise for instance when the aircraft pitches or rolls through a 90 degree pitch or bank angle, or where for any reason the drive signals controlling the display have to be switched from one drive law to another, or where the aircraft symbol reaches or leaves a limiting condition. This invention relates to a new system for positioning the indicia including the aircraft symbol, the pitch ladder, peripheral scales and a special velocity vector symbol, so that they provide better indications for guidance and control generally without discontinuous or disconcerting movements. A related patent application to be filed on the same day as this application discloses various special features of the preferred form of pitch ladder display.

In most head-up displays the peripheral scales are written at fixed positions relative to the field of view of the display. When flying in strong crosswinds or in conditions which produce high incidence or sideslip, with a conventional display in which the aircraft symbol moves to follow the actual direction of the flight path it can often move far enough to overwrite or become confused with one of the peripheral indications, which may make them unreadable or cause dangerous confusion. One suggested solution is to make the peripheral indications move with the aircraft symbol, but when this is done a scale or value may be moved off the edge of the field of view and its information lost, unless limits are applied, leading to discontinuity of information and conditions where the symbol shows only that the limit has been exceeded.

It is also important to maintain horizon correlation as far as possible and to ensure that when the horizon is visible in the field of view, the elevation of the aircraft symbol above the actual horizon as seen shall accurately show the actual pitch or climb-dive angle of the aircraft. It may not be possible to ensure this at high angles of incidence or bank but it is important to indicate when horizon correlation is lost and to recover it as quickly as possible when the relevant angles return to a lower value.

A further very important requirement is that when the pilot returns his attention to the display after having any distraction or task requiring him to look elsewhere or after any violent manoeuvre, the display will immediately and accurately show him all necessary information and present it in some way which avoids any possibilities of misinterpretation.

It is an object of the invention to provide an aircraft instrumentation system which avoids the overwriting or loss of peripheral indications, maintains horizon correlation over a fairly wide range of conditions, and includes features planned to reduce or avoid possible misinterpretation.

According to the invention there is provided an aircraft instrumentation system including a head-up display, means for receiving signals representing parameters and/or measurements of the aircraft's position, and means for providing and controlling the position of an aircraft symbol on the head-up display, characterised in that it is arranged to position the aircraft symbol on the centreline of the display field at a displacement from a datum point, the displacement being generally of the form Y_acs = Y_A + Q_A where Y_A is a function appropriate for maintaining horizon correlation over a moderate range of bank angles, and Q_A is a function which follows transient variations in the aircraft's pitch or attitude and offsets and reduces undesirable transient movements of the aircraft symbol.

Preferably the displacement Y_acs is subject to an overriding requirement that Y_acs shall be limited to keep the aircraft symbol within the field of view of the display and the symbol shall be distinctively altered when its position is limited by this condition.

The function Q_A is selected and calculated to remove most but not all of the unwanted motions of the display indicia which are due to the transient changes in angle of attack which are needed to cause a change of flight path. Its calculation involves a time constant which varies according to a relationship which is matched to the characteristics of the aircraft as hereinafter described. This makes the aircraft symbol move steadily in response to changes instigated by the pilot, and it moves smoothly to the position representing the climb/dive angle reached by the manoeuvre. Preferably the system also has means for providing a comparatively small symbol to show the actual direction of the flight path. This symbol may be a diamond and it will be called the velocity vector or w diamond. To position it correctly allowing for the bank angle ∅ of the aircraft, and to remove most of the unwanted transient motion associated with changes of the aircraft's course, it is placed at a position in the display field representing the instantaneous relative direction of flight modified by a displacement Q_v upwards with respect to the display plane where Q_v is a function which follows transient variations in the aircraft's pitch or attitude.

The displacements of the w diamond are unllimited and it may at times go outside the field of view of the display.

The system will also have means for providing and positioning a pitch ladder display to give the pilot clear indications of bank angle and climb-dive angle.

The pitch ladder is a pattern of bars, each drawn (subject to inaccuracies, resulting from the sensors and instrumentation) parallel to the actual horizon and forming a scale against which the position of the aircraft symbol indicates a measure of the climb or dive angle of the aircraft. A central bar in the pattern represents zero climb/dive angle ie level flight, and it should overlay the actual horizon as seen through the display whenever the horizon comes within the field of view of the display. It follows that the pattern must be tilted, with respect to the display centreline, by minus the bank angle of the aircraft, and the pattern centre so placed that the aircraft symbol position gives the desired climb or dive angle measurement when read against the scale formed by the pitch bars.

In systems according to the present invention it is important that unwanted transient motion of the pitch ladder due to changes in the angle of attack associated with changes of flight path shall be reduced or suppressed and therefore it is positioned to indicate ɤ + Q_L cos ∅ when read against the aircraft symbol, where ɤ = climb or dive angle and Q_L is a function which follows transient variations of the aircraft's pitch or attitude. Hence the centre of the pitch ladder pattern will be placed at a point defined by head-up display coordinates X_plc = sin∅.f( ɤ+ Q_L cos ∅)

Y_plc = Y_acs - cos ∅.f( ɤ + Q_L cos ∅)

The centre of the pitch ladder pattern may be off the field of view; only the portion of the pitch ladder pattern which falls within a prescribed window area of the display will actually be displayed. In these equations f( ) represents the relationship between displacement from the centre of the pitch ladder pattern and the actual pitch. This is preferably a nonlinear relationship as detailed in the aforesaid copending application.

There are several alternative functions which may be used for maintaining horizon correlation: 0 6 Cx

c^o

In these equations, α _f is a signal representing the angle of attack of the aircraft, suitably smoothed to reduce noise; ω is the angle subtended by the aircraft upward velocity component in a plane orthogonal to the longitudinal fuselage datum direction; it will usually be negative, as explained later; and

∅ is the bank angle of the aircraft conventionally considered positive when it represents a clockwise rotation around the longitudinal datum.

Of these alternatives, (i) which appears simplest, has several disadvantages. The signal α _f is derived from an aerodynamic sensor which is subject to turbulence and tends to provide a noisy signal , and it does not take into account any effects of winds . The alternatives using ω are preferred because ω is available from inertial navigation equipment; it is a smooth and highly reliable signal, and it takes into account the effects of vertical and horizontal winds.

Alternative (iii) may be used with confidence only in an aircraft which is never flown with a bank angle∅ outside a range from -∅_m to +∅_m where∅_m is less than 90 degrees. This is because sec 0 has a discontinuity at∅ = ±90º , and the extreme change in values of Y which occurs as the aircraft rolls through∅ = ±90º causes the aircraft symbol to flash rapidly up or down on the display; such movement is disconcerting; it can lead to spatial disorientation and may be misinterpreted as a sign of system error. The alternative (ii) is preferred for any aircraft which may be rolled through ∅ = ±90º , and it has been fully tested and found very satisfactory in all circumstances.

The functions (iv) and (v) are possible alternatives whose feasibility and consequences have not yet been fully studied. The function (iv) makes Y_A change linearly from ω sec∅₁ to ω sec (180º -∅₁) ; the discontinuity at these values may cause difficulties. If possible, ∅₁ should be chosen so that ω sec∅₁ will probably be slightly larger than the maximum displacement within the field of view of the head-up display. The first form of function (v) shows its derivation and behaviour; where cos ∅ has a small value it becomes approximately ωKc os ∅ but where cos ∅ is nearly one it becomes a good approximation to ω sec∅, It has the advantage of being a smooth and continuous function at all values of∅. The second form given is clearly preferable for computation. The constant k may have any value; it determines the range of values of∅ over which the function is close to sec ∅, and its steepness near ∅ = 90°.

These alternatives should not be taken as an exhaustive or limiting list; persons skilled in the art having read this specification may well derive other functions which may be more convenient or possibly more accurate for achieving the desired usefulness.

The preferred requirement that the displacement shall be limited to keep the aircraft symbol within the field of view should be remembered. In a typical current head-up display the field of view may extend in elevation from about +5 to -15 degrees relative to the pilot's sight line along the longitudinal datum direction. The aircraft symbol displacement Y_acs may therefore be limited to a range from about +2 to -12 degrees.

However it is difficult to draw any conclusions about when Y_acs will reach the limiting conditions, because ω is not independent of∅.

The functions Q_a, Q_v and Q_L depend on the pitch rate or attitude rate of the aircraft. For many flight conditions it is highly preferable or necessary to use a Q function calculated from the body pitch rate of the aircraft measured with respect to reference axes in the aircraft; this variable is called q. For near-level flight close to the ground it is necessary to use a Q function derived from the pitch angle θ of the aircraft measured with respect to the earth. Hence it is highly desirable to compute two functions continuously, Q1 which is derived from θ and Q2 which is derived from q. These functions are calculated by data processing algorithms which are mathematically defined as operations having the transfer functions of the forms given in the equations

r

where Ʈ is a time constant matched to the aircraft's characteristics as hereinafter described; Ʈo is a much shorter fixed time constant which may be o.04 seconds; S is the Laplace transform operator, and |θ| represents the modulus of θ,ie its value regardless of sign. In both expressions G = gain will be not greater than 1.0 and preferably about 0.7.

In practice all sensor signals will be sampled and digitised with sampling intervals ΔT, and there are various well-known algorithms in finite-difference form which give results approximating to the desired transfer functions. The preferred finite-difference form of the calculations for Q1 and Q2 uses the equations:

W

where P1, P2 and W1 are intermediate variables, Δt = sampling interval and exp = exponential function.

The function Q2 is most effective in reducing unwanted display movement especially when high bank angles and rolling or pulling manoeuvres are used, but it will tend to give inaccurate horizon correlation in such manoeuvres, which could be dangerous in flight close to the ground. The function Q1 maintains better horizon correlation but is less effective in inhibiting unwanted display movements at high bank angles. A third function Q3 may be derived from Q1 and Q2. This is equal to Q1 when -10 < θ < 10º and equal to Q2 when

|θ| >30º, and between these ranges it is a linear blend which changes smoothly from Q1 to Q2 as |θ| varies:

30⁴ >θ >10°.

The values Q_A, Q_V and Q_L to be used are selected according to the type of aircraft and its flight situation as follows:

For an agile combat aircraft operating away from the ground, A_A = Q2 is preferred.

For all aircraft in take-off or landing approach

manoeuvres Q_A must be either Q3 or Q1; automatic switching means should be arranged to prevent Q2 being used when the undercarriage is down.

For other cases Q_A may be either Q2 or Q3.

In all cases Q_v is preferably = Q3 or Q1. Q_v must never be made = Q2 when |θ|< 10°.

Similarly Q_L is preferably = Q3 or Q1 and it must never be made = Q2 when | θ| <10°.

In the expressions for calculating the Q variables the time constant should, as far as possible, be made equal to the time constant of the heave mode response of the aircraft. For this reason it should vary as

where V_T is the airspeed and ϭ is the relative air density ratio. To get an approximate match with a simple linear expression, suitable for cases where mass, wing area and lift curve are almost constant we set

where A and B are constants to be determined by experiment as hereinafter described. Ʈ may be limited to 0.2<Ʈ<6 seconds.

Embodiments of the invention will now be described in greater detail with reference to the accompanying Figures, of which:

Figure 1 is an isometric sketch illustrating the geometry of an aircraft's flight path and definitions of terms used in the description;

Figure 2 is an orthogonal view of lines in the plane of the head-up display;

Figures 3a and 3b are graphs showing parameters of a flight plotted against time;

Figure 4 is a sketch showing the pilot's sight-line to the display in a side view;

Figure 5 is a block diagram indicating the equipment used in an embodiment of the invention;

Figures 6a and 6b show preferred forms for the aircraft symbol; and

Figure 7 shows the preferred form and arrangement of peripheral indications on the display.

Figure 1 shows a line OH representing the longitudinal fuselage datum or fore-and-aft direction of an aircraft in flight, the projection of this line on a horizontal plane is called the heading and is shown as a line ON. The angle between ON and OH, in a vertical plane with respect to the earth, is the pitch angle θ of the aircraft. Because of incidence (angle of attack), sideslip, vertical winds and crosswinds the actual flight path direction is along a line OP. The projection of the flight path OP on the horizontal plane is called the track, OT, and the angle between heading ON and track OT, in the horizontal plane, will be called∈ . The angle, in a vertical plane, between the flight path OP and the track OT is the climb-or-dive angle ɤ and it will usually be significantly less than θ.

Inertial navigation equipment in the aircraft will be arranged to measure the aircraft's velocity components in three orthogonal directions and to provide signals representing their relative magnitudes. As shown in Figure 1, H, F, P, T, N, E are points in a plane orthogonal to the longitudinal fuselage datum direction. V_x is the velocity component along the datum direction OH. V_y is the velocity component in the direction of a horizontal line with respect to earth. V_z is the velocity component in the direction of the intersection of the heading plane (OHEN, a vertical plane with respect to earth) with the plane HFPTNE. Being orthogonal to OH the plane is tilted at angle θ to the vertical.

Let 0 be the current position of the aircraft and P be its position at time δt later. Then OH = V_x δt

HF = V_y δt

HE = V_z δt

and the actual flight path FPA passes through point P which is the resultant of vectors HE and HF. OE is the flight path elevation (FPE) which is the projection of the actual flight path on to the heading plane. The inertial navigation equipment will provide signals representing either the ratio of velocity or alternatively the related angle

and also the ratio altern¬

atively η = p where η is the angle < HOF, not shown on

Figure 1 for the sake of clarity. Because of the angle of attack needed to maintain course and achieve lift, the angle ω will usually have a negative value.

It will be assumed that the indicia of the head-up display are formed as images in a plane orthogonal to the longitudinal fuselage datum. If in any case this is not precisely true, the consequences can be considered if necessary as a modification to the following presentation. Without loss of generality we can take the plane containing H, F, P, T, N, E as the display plane, take the point 0 as the average pilot's eye position, and take the distance OH as a unit distance.

Figure 2 shows lines in the head-up display plane looking along the direction of the aircraft's longitudinal datum OH which therefore appears in Figure 2 as a point H. The heading plane (vertical with respect to earth) is seen edge-on as the line HEN and the horizon is seen as a horizontal line through NT. In general the aircraft may have a bank angle ∅ which is a rotation or roll about the longitudinal datum. Consequently, the centre of the head-up display, which would be vertical in level unbanked flight, is tilted at an angle ∅ to the vertical plane OHEN as indicated by the line HC. Displacements of positions in the head-up display will naturally be specified with regard to the aircraft equipment by rectangular coordinates X and Y where X indicates displacement transverse from the line HC, positive to the right, and Y indicates displacement from H parallel to HC, positive upwards. The dashed line FOV indicates the field of view of the head-up display and it may be noted that it typically extends in elevation over a range from about Y = +5º to Y = -15º. (It is convenient to measure these displacements in terms of the angle subtended at the pilot's eye position, point 0 in Figure 1; if OH is taken as a unit length, and the angles are converted into radian measure the displacements from H are substantially equal to the angles; strictly the displacement would be the tangent of the angle subtended).

It is conventional on a head-up display to show the direction of flight, relative to the aircraft, by producing an indicium at the point P where the actual flight path direction intersects the head-up display plane. The inertial navigation equipment will give to and η which determine the vertical and transverse displacements of P from H in the plane HFPTNE. In manoeuvring flight the Q_v function adds an offsetting displacement in the direction of the Y coordinate of the display, and allowing for the head-up display being rotated by -∅ it follows that the display coordinates of P will be given by:

The actual direction of flight, indicated by P, is the resultant of a number of factors, several of which cannot be directly measured by instrumentation in the aircraft. The angle of attack which is the attitude of the aircraft relative to its direction of travel through the mass of air in which it is flying may reasonably be represented by a vector oe extending from H towards C. The effects of sideslip are represented by a vector β transverse to the line HC. The effects of crosswinds and any vertical drift of the air mass are represented by vectors Ψ_w and ɤ _w respectively. Obviously all these vectors are variable and will not always have the relative values shown in the Figure.

In a conventional head-up display, an aircraft symbol would be generated around the point P. As previously explained in a system according to the present invention it is desired to position the aircraft symbol always on the centreline of the display, but so as to preserve horizon correlation as farr as possible. It is therefore desired to position the aircraft symbol on the centreline HC but at a point near S whose elevation above the horizon is substantially the same as the elevation of the point P. This elevation represents the climbdive angle ɤ in steady flight. The desired alignment may be substantially achieved by placing the aircraft symbol at a position given by any of the alternative expressions for Y_A given in the introduction. Remembering that HE and EP will be equal to ω and η respectively, the terms in

these expressions can clearly be seen to correspond to various distances on the drawing. For instance in the alternative (ii) the first term represents the projection of HE onto

HC, and the second term represents a part of the displacement due to the angle of attack. As noted in the introduction the displacement of the aircraft symbol Y_acs is limited to ensure that it remains within the field of view. For instance a typical display unit may have a field of view extending from about +5º to -15º and on such a display the aircraft symbol displacement could be llimited to the range from +2º to -12º. When this limitation occurs the symbol is distinctively altered. Figure 6a shows the conventional form for the aircraft symbol in normal or unlimited condition, and figure 6b shows a form for the symbol showing a fin 60 to indicate that it is in a limited condition.

When so limited the position of the aircraft symbol does not show the climb/dive angle, but the position of the pitch ladder relative to the aircraft symbol does continue to show the climb/dive angle accurately in steady near-level flight. In manoeuvring flight, as explained with reference to figure 3b, the addition of the Q term makes the relative position of the pitch ladder show the trend of the climb/dive angle towards the result of the manoeuvre. As the w diamond is unlimited, while it remains within the field of view it also provides a useful and accurate indication of the flight path, and it retains horizon correlation.

Figure 3a is a graph plotting the altitude of an aircraft against time, and Figure 3b is a graph plotting the pitch and climb-dive angles of aircraft on similar flight paths to that shown in Figure 3a.

On the lefthand side of these graphs the aircraft is in a gradual climb with climb angle ɤ ₁, angle of attack α₁ and pitch θ₁. On the righthand side the aircraft is in a steeper steady climb with climb single ɤ₂, angle of attack a₂, and pitch θ₂. To move from the gradual climb to the steeper climb the pilot causes a substantial transient increase in the angle of attack; the flight path follows with a considerable lag until the pilot eases the pitch angle back to a value which just maintains the desired steeper climb.

Figure 3b is a graph showing the variations of θ and in this manoeuvre. In a conventional head-up display, if the aircraft symbol is positioned to show climb-dive angle with horizon correlation, the considerable increase in the angle of attack moves the symbol well down the display and possibly off the field, or to a limiting position where it is not very helpful to the pilot. Because the head-up display is fixed in the aircraft, its datum point inevitably follows the pitch angle variation, and the downward displacement of the aircraft symbol follows the difference between the θ curve and the ɤ curve.

In the present invention the Q term is applied to compensate for and considerably reduce this effect. It makes the aircraft symbol position follow the curve ɤ + Q_A. This is further illustrated by Figure 4 which is a sketch of lines in the plane containing the head-up display centreline HC and the pilot's eye position 0. As shown the position of the aircraft symbol A is determined by the displacement Y (negative and acting downwards) offset by the term Q. The position of the w diamond is similarly modified.

If the gain G is set to 1.0 and the time constant in the calculation of Q is matched to the characteristics of the aircraft the curve ɤ + Q_Ashould closely follow the shape of the θ curve. However in practice it has been found better to set the gain G at about 0.7 so as to attenuate but not eliminate the effects of the incidence changes which control the manoeuvre.With this setting the aircraft symbol makes a more gradual movement which the pilot can use as a very helpful guide to achieving the new desired condition.

The pitch ladder is a pattern of bars which are always parallel to the horizon. They form a scale of climb-or-dive angles which has a linear one-to-one relationship with the actual climb-or-dive angle ɤ for angles close to zero, for instance within the range from +5° to -5°, but with an increasingly compressed relationship as the angle moves towards the limits of +90° and -90°. These limits are indicated by distinctive symbols rather than bars. The pitch ladder pattern will be so positioned that if the aircraft symbol is read as a pointer against the scale formed by the bars of the pitch ladder it will indicate ɤ + Q_L cos ∅ where ɤ is the current climb-or-dive angle. The angle of the bars, and the angle of the axis of symmetry of the pattern will provide the pilot with very clear indications of the bank angle. A more detailed description of the preferred form of pitch ladder is given in the copending application previously mentioned.

Figure 5 is a block diagram showing diagrammatically the equipment of an aircraft instrumentation system. It comprises a data processor 51 having inputs 52 connected to receive signals from various standard sensors (not shown). Outputs from the data processor 51 control a head-up display unit 53. Also provided are means 54 for setting parameters to be used in calculations within the data processor 51.

An experimental method for optimising the parameters of the Q calculations to match the characteristics of a particular aircraft type will now be described. The pilot must be able to alter parameters G, A, B and δk (where δΑ represents a chosen small change in A) during flight trials in an aircraft fitted with a prototype system, or in a simulator which accurately reproduces the aircraft's flight characteristics. The prototype system must have a facility for generating an experimental cross on the display at a controllable position.

Step 1: Set G = 1.0, A = 0, A = 0, B = 300

In steady flight at about the maximum normal speed for the aircraft, position the experimental cross in the centre of the aircraft symbol and also set V_T, m and σ to their appropriate values for the flight conditions.

Step 2: Increase the flight path angle by about 5 and note the relative movement of the aircraft symbol with respect to the cross. As the cross is effectively fixed with respect to the aircraft it will be following the variations of pitch angle (curve θ in Figure 3b) but below it, starting coincident with ɤ₁.

The aircraft symbol may be temporarily depressed below the cross - in this case increase δA. If in any case the aircraft symbol rises temporarily above the cross ok should be decreased. Note values of δA, V_T and σ when it appears that a reasonable match has been achieved.

Step 3: Repeat step 2 for various flight conditions (particularly various initial flight angles). It will not be possible to achieve a perfect match in all cases but the aim is to achieve a compromise which is reasonably well matched over a range of flight angles, keeping V_T and constant as far as possible. After several iterations reduce G to 0.7, check the effects by further tests and select the δA which appears best for this value of the product V ϭ which must be noted as (V ϭ )₁.

Step 4: Now put the aircraft into steady flight in a nearminimum normal speed and repeat steps 1-3 to get a selected value δA₂ suitable for use when V_T ϭ has the new value

(v ϭ)₂.

Step 5: Calculate

and reset the parameters A and B to the values A new and

B new. Reset δA to zero. Repeat steps 1 to 5 with these values A new and B new in place of A and B. The process should be repeated until the magnitudes of δA₁ and δA₂ are less than 0.01 and 0.02 respectively. Satisfactory operation should be checked over the full range of operating conditions.

The values of A and B may then be fixed for all identical aircraft. However if the system is fitted in an aircraft whose mass may vary considerably between unloaded and fully-loaded conditions it may be desirable to provide means for selecting values of A and B appropriate to the load condition of the aircraft. Similarly if the aircraft's wing area or left-curveslope coefficient may be varied significantly, appropriate changes should be made preferably by interlocked switching so that they are always made when the aircraft configuration is changed. In some such cases, a different equation for Ʈ might be useful to achieve better matching, and details of the optimising method would be correspondingly altered.

The system may also provide indications of airspeed V_T, heading (compass angle), track direction, altitude, vertical speed, angle of attack, engine revolutions per minute and Mach number by peripheral indications in a fixed format which is moved with the aircraft symbol, so that they always appear in a constant position with respect to the aircraft symbol.

Figure 7 shows the preferred form for the peripheral indications. In this example, the aircraft symbol is seen in the centre. At the bottom left corner is the engine rpm indication in this example indicating 85.5% of a prescribed rpm rating. The line anddots above the rpm indication constitute a scale against which the angle of attack α is shown by an arrowhead, and a vertical line which extends from the zero mark to the arrow position. The digital indication, in this example 0.78, shows Mach number. Airspeed (in this case 500 knots) is shown by a counter-pointer system in which a radial pointer moves round within a circle of dots making one revolution per 100 knots. The small triangle outside the circle can be positioned to indicate any desired speed.

The digits at the top are part of a moving-tape indication of heading (compass angle). The vertical line on the centre of the format acts as a pointer against which the moving-tape display shows the heading, in this case 160 from North. The tape scale is compressed by a factor of 5 to 1; that is to say the distance representing 10 degrees on the scale subtends an angle of 2 degrees at the pilot's eye position. The arrowhead under the scale points to the compass bearing of the track.

Altitude is shown by a counter-pointer display at the top right corner. In this case the pointer makes one revolution per 1000 feet, and the triangle outside the circle of dots is a marker which can be positioned at any desired position around the circle.

Vertical speed is shown by the position of an arrowhead against the vertical scale formed by the lines on the righthand side. The graduations of this scale are evenly spaced but they represent non-linear speed increments; the central, longer line represents zero and the other lines represent ±500, ±1000, ±2000, and -4000 ft/min respectively. Of these peripheral indications, altitude is essential; digital speed, heading and track almost essential, but the others may be regarded as optional and would only be displayed if desired.

The data processor 51 will be a highly reliable computer programmed to carry out the following actions in a repeating cycle:

Receive and store updated sample values of all sensor inputs.

Calculate Ʈ using set values for A and B and sensor inputs for V_T and ϭ .

Calculate Q1, Q2 and Q3 from sensor inputs of q and θ using the calculated value of Ʈ.

Calculate Q_A, Q_V and Q_L.

Calculate Y_a.

Calculate Y_acs, X_p, Y_p, X_plc, Y_plc.

Generate graphics signals to display the peripheral indications around the aircraft symbol and position them at positions determined by Y_acs. Generate signals to display the w diamond and position it at X_p, Y_p.

Select graphics signals to generate the appropriate part of the pitch ladder pattern and position it appropriately according to the values of X_plc, Y_plc.

Claims

1. An aircraft instrumentation system including a head-up display, means for receiving signals representing parameters and measurements of the aircraft's position, and means for providing and controlling the position of an aircraft symbol on the head-up display, characterised in that it is arranged to position the aircraft symbol substantially on the centreline of the display field at a displacement from a datum point, the displacement being of the form Y_acs = Y_A + Q_A where Y_A is a function appropriate for maintaining horizon correlation over a moderate range of bank angles, and Q_A is a function which follows transient variations in the aircraft's pitch or attitude and offsets and reduces undesirable transient movements of the aircraft symbol.

2. An aircraft instrumentation system as claimed in claim 1 and wherein the displacement Y_acs is limited to keep the aircraft symbol within the field of view of the display and the aircraft symbol is distinctively altered when its displacement is limited by this condition.

3. An aircraft instrumentation system as claimed in claim 1 and also comprising means for providing and controlling the position of a comparatively small symbol on the display to show the flight path direction modified by a function Qv which follows transient variations in the aircraft's pitch or attitude and offsets and reduces undesirable transient movements of this symbol.

4. An aircraft instrumentation system as claimed in claim 1 and also comprising means for providing and positioning a pitch ladder display with a displacement and orientation such that the position of the aircrft symbol relative to the pitch ladder substantially indicates ɤ + Q_L cos ∅ where ɤ is the climb or dive angle of the flight path, ∅ is the bank angle, and Q_L is a function which follows transient variations of the aircraft's pitch or attitude and offsets and reduces undesirable transient movements of the pitch ladder.

5. An aircraft instrumentation system as claimed in claim 1 and also comprising means for providing and positioning further indications of various measurements always with a fixed relationship to the position of the aircraft symbol.

6. An aircraft instrumentation system as claimed in claim 1 and wherein the function Y_A is substantially Y_A =ω cos ∅ - α_f sin²∅ where ω is the angle between longitudinal fuselage datum (LFD) and flight path elevation (FPE) ∅ is the bank angle, and α_f is a smoothed signal prepresenting angle of attack, noting that ω is considered negative when FPE is below LFD and ∅ is considered positive when it represents a clockwise rotation about LFD.

7. An aircraft instrumentation system as claimed in claim 1 and wherein the function Y_A has any of the alternative forms disclosed in the specification.

8. An aircraft instrumentation system as claimed in claim 1 and also comprising means for calculating one or more Q functions according to algorithms involving a variable time constant Ʈ which is matched to heave mode response characteristics of the aircraft.

9. An aircraft instrumentation system as claimed in claim 8 and also comprising means for calculating the time constant Ʈ according to an equation substantially of the form

^w

where A and B are constants determined by experimental trials, V_T is the true airspeed, and ϭ is the relative air density ratio.

10. An aircraft instrumentation system as claimed in claim 9 and also comprising means for selecting constants A and B appropriate to the loading or configuration of the aircraft.