WINDSHEAR DETECTION AND WARNING SYSTEM

Abstract: Windshear warning apparatus for an aircraft, which apparatus separately compares a plurality of signals derived from horizontal and vertical inertial acceleration and air mass acceleration components to indicate incipient windshear. The compared signals are substantially equal under normal wind conditions but are significantly different under windshear conditions. A signal proportional to the product of the magnitude and rate of change of the total or resultant inertial acceleration of the aircraft is used to provide a warning of a significant windshear condition.

Description: The present invention relates generally to aircraft flight performance computer systems and more particularly to a system for detecting and/or providing a warning of impending dangerous windshear conditions and to provide an avoidance command through the pitch director and throttle systems to enable the flight crew to avoid such conditions in a timely and safe manner. Windshear is a weather condition which results in rapidly changing wind velocity and direction and in terms of aircraft flight performance is, of course, most dangerous during take-offs and landing approaches. A number of attempts and proposals have been made to provide the flight crew of an aircraft with a warning of impending windshear conditions. One method presently in use is based on ground observations of wind conditions at sporadic locations around an airport. These wind determinations have been shown to be inadequate during several occurrences which have resulted in serious accidents. Other methods utilise on-board apparatus for detecting ground speed and comparing the ground speed to a measure of the airspeed of the aircraft and may include some sensing of vertical inertial acceleration. U.S. Patent Specifications Nos. 4,0l2,7l3 and 4,079,905 disclose arrangements typical of such systems. A prior system previously described by one of the present inventors is based on the energy residing in the aircraft at any instant as it proceeds along a flight path and is disclosed in British Patent Specification No.2,094.006A. A deficiency of the prior on-board systems is that while they may be effective in determining horizontal wind variations, they are inadequate for adequately determining vertical wind variations such as downdrafts and updrafts. Analyses of accidents that have been caused by severe wind variations near the ground indicate that they are composed of rapidly changing combinations of both wind components. These result in rapid changes in the magnitude and direction of the total acceleration vector which determines the aircraft's flight path in the vertical plane. A further deficiency of prior art systems that utilise indicated airspeed (IAS) or calibrated airspeed (CAS) is that they do not take into account changes in the density of the air mass at varying temperature and altitude, thereby resulting in errors in the apparent wind rate. A typical windshear situation is graphically illustrated in Figure l of the accompanying drawings. An aircraft approaching a runway, 3, attempts to fly along a glide slope 4, defined as the straight line between P₁ and P₂. If a wind downburst exists along the glide slope, as illustrated in Figure l, the initial effect on the aircraft occurs at a point 0₁ on the glide slope, where it encounters an increasing headwind. This causes an increase in lift and a deceleration relative to the ground. The acceleration vector, designated A VH , is the resultant of the inertial vertical acceleration component a V and the inertial horizontal acceleration component a H . At point 0₁, the vector A VH points upwards and backwards with respect to the aircraft's path relative to the ground. Thus, there is a significant change in the magnitude of the resultant acceleration vector as the aircraft travels between P₁, where there is no appreciable acceleration vector, and 0₁ where there now exists a sizeable acceleration vector. The increase in lift at point 0₁ causes the flight path 5 of the aircraft to go above the glide slope 4. There is also an increase in airspeed at 0₁ due to the inertia of the aircraft carrying it into the headwind and in spite of the fact that the aircraft is being decelerated relative to the ground. The pilot's normal reaction under these conditions is to reduce thrust and pitch the aircraft down to reduce airspeed and to realign the aircraft to its intended glide path 4. As the aircraft passes the centre of the wind burst column, the headwind changes to a tailwind. The tailwind causes a decrease in airspeed, again due to its inertia, carrying it with the tailwind and a forward acceleration relative to the earth. The decrease in airspeed now results in a loss of lift at point 0₂. The combined actions of the reduction of thrust at 0₁, the pitch down manoeuvre at 5, forward acceleration at 0₂ and the loss of lift at 0₂ results in a larger resultant acceleration vector at 0₂ which is directed forwardly and downwardly. If not corrected rapidly by the pilot by increasing thrust and pitching the aircraft nose up to increase its angle of attack, the aircraft will impact the ground violently at point 6. In the windshear example discussed above, it will be noted that during the transition from point 0₁ to point 0₂, the direction of the total acceleration vector A VH has undergone a large angular rotation on the order of 180°. EP-A-0 125 087 discloses a windshear warning apparatus which separately compares a plurality of signals derived from horizontal and vertical inertial acceleration and air mass acceleration components to indicate incipient windshear. The compared signals are substantially equal under normal wind conditions but are significantly different under windshear conditions. A signal proportional to the product of the magnitude and rate of change of the total or resultant inertial acceleration of the aircraft is used to provide a warning of a significant windshear condition. Timing circuits provide for activating a warning system after a windshear has persisted for a predetermined time. However, it is known that depending on the characteristics of a particular aircraft (i.e., power-to-weight ratio) and the magnitude of the windshear, that safe flight may advantageously be sustained over a time period which is variable, and which may exceed the fixed time period of EP-A-0 125 087. Hence, unnecessary alerts and alarms will be generated for windshear encounters during which the aircraft is capable of sustaining safe flight. Under normal flight conditions, that is no wind or essentially steady state winds, the inertial and air mass accelerations of an aircraft will be essentially equal, whilst in rapidly changing winds they will not be essentially equal. Therefore, in accordance with the present invention, which is defined in the appended claims, the vertical and horizontal components of the aircraft's inertial acceleration are compared with the vertical and horizontal components of the aircraft's acceleration relative to the sustaining air mass and any discrepancy above a predetermined value is used to initiate a warning that a windshear condition exists. The rate of change of any discrepancy, together with the rate of rotation of the aircraft's resultant acceleration vector, is monitored in both sense and magnitude so as to provide a measure of the severity of the windshear condition. When the magnitude and direction of any one of such parameters exceed prescribed limits, the warning is changed to indicate to the pilot that evasive action is required. The present invention provides flight commands through the flight director and autothrottle systems to pitch up to a maximum safe angle of attack and to increase throttle to a maximum safe power setting. Windshear detection and warning apparatus in accordance with the invention will now be described in greater detail, by way of example, with reference to the accompanying drawings, in which:- Figure l is a diagram already briefly discussed which illustrates the basic problem of an aircraft encountering a downburst windshear condition during an approach to a runway, Figure 2 is a diagram which illustrates the characteristics of the acceleration vector in the vertical plane and its relationship to aircraft movement in the air mass, Figure 3 is a diagram which indicates the effects of normal and longitudinal acceleration on the configuration of a liquid level toroidal sensor which is mounted on the longitudinal axis of the aircraft, Figure 4 is a block diagram which illustrates a preferred embodiment of the windshear detection and warning system, Figures 5A and 5B present a preferred embodiment of the angle of attack display used in the windshear detection and warning configuration shown in Figure 3, Figure 6 is a block diagram which illustrates a preferred embodiment of the apparatus which generates the alert outputs of the windshear computer and detector of Figure 4, Figure 7 is a block diagram of the alpha function computer of Figure 4, Figure 8 is a block diagram of the data processor of Figure 6, Figure 9 is a diagram of the switching logic used to control the two level warning annunciator of Figure 4. Figure l0 is a block diagram which illustrates a windshear detection circuit useful in the circuit of Figure 6, and Figure ll is a timing diagram useful for understanding the operation of the circuit of Figure l0. Referring again to Figure l, the wind downburst l consists of a vertical column of air which is usually generated by severe thunderstorms and moves downwards towards the ground 2 where it is deflected by the ground outwards from the centre of the column in all directions. Whilst this column is illustrated as centred over the glide slope approach path to the runway, its effects are the same on the runway take-off path or if the column is laterally displaced from the flight path where there are large components of the wind in the direction of flight of the aircraft. A very significant characteristic of the aircraft's resultant or total acceleration vector A VH between points 0₁ and 0₂ is a rotation of the direction of the vector as the aircraft enters the disturbance and leaves the disturbance. It is, therefore, an object of this invention to detect not only the change in magnitude of the resultant acceleration vector A VH , but also to detect the rate of change of its direction and to use these characteristics to provide a warning to the flight crew. It is a further object to provide thrust and pitch guidance to enable the flight crew to avoid the dangerous condition that exists during a windshear condition, such as at point 0₂ of Figure l. Figure 2 illustrates the characteristics of the A VH acceleration vector referred to in Figure l. The A VH vector 7 is the acceleration in the vertical plane of the aircraft, considered to be a single mass at point O. The vertical plane is defined as the plane which contains the line 9 that extends from point O to the centre of the earth and the aircraft longitudinal axis 10, designated as the line from O to X. The horizontal plane 11 is orthogonal to the vertical plane and also contains the aircraft point mass at O. The intersecting planes 8 and 11 form an earth reference which consists of mutually orthogonal lines 12, 13, 14. Line 12 extends between points O and H; line 13 extends between points O and V; line 14 extends between points O and L. The Angle ϑ is the angle between reference line 12 and the aircraft longitudinal axis 10. Angle ϑ is defined as the pitch angle of the aircraft. The resultant or total acceleration vector 7 consists of mutually orthogonal vector components 15 and 16. Component 15, designated a H , is the inertial acceleration along reference line 12; component 16, designated a V , is the inertial acceleration along reference line 13. The rate of rotation of the acceleration vector 7, as would occur in the situation previously considered in Figure 1, is the time rate of change of angle ρ , which is the angle between vector 7 and inertial reference line 12. The rotational rate of 7, designated ρ ˙ , is depicted as the magnitude of the vector 17 which is orthogonal to the vertical plane 8 and passes through the aircraft point mass 0. It is useful to consider the product of the rate vector 17 (ρ ˙) and the magnitude of the vector 7|A VH |. This product, designated |A VH |ρ ˙ , is the direction of which is colinear with the vector 17 in accordance with well known principles of the product of a vector and a scalar quantity. The use of this product assures that it is nominally near zero for small values of A VH (in the absence of vertical wind gusts) regardless of the value of ρ ˙ and likewise is near a null for small values of (when the rotation rate of the total acceleration vector is essentially zero) regardless of the value of A VH . Thus, significant values of A VH x ρ ˙ require both a significant magnitude of the acceleration vector and a significant rotation rate of the total acceleration vector 7, both of which together are indicative of a dangerous windshear condition. Accordingly, while neither A VH nor ρ ˙ by itself may be representative of a dangerous windshear, when they both occur at the same time, it indicates that a dangerous windshear exists. The aircraft longitudinal axis l0 has a movement relative to the air mass. This movement can be conveniently considered to be comprised of two orthogonal velocity components l9 and 20, existing along lines l2 and l3, respectively. True airspeed is generally designated V T and is measured along the flight path of the aircraft by a conventional air data computer. It is sufficiently accurate for the purpose of the present invention to consider that V T coincides with vector l9 since the angle between the aircraft flight path and reference line l2 is usually small for large aircraft such as transport aircraft. Vertical speed, designated h, is also calculated by a conventional air data computer and is derived from the rate of change of the atmospheric static pressure existing around the aircraft. If the aircraft is not being subjected to windshear conditions, the time rate of change of true airspeed l9, designated V ˙ T , essentially equals the horizontal inertial acceleration component l5. Similarly, the time rate of change of vertical speed 29, designated , essentially equals the vertical inertial acceleration component l6. However, these two relationships, i.e. V ˙ T = A H and = a V , are disturbed during windshear conditions, such as described in Figure l, and the magnitude of the disturbance is indicative of the strength of the windshear. As previously discussed, windshear conditions also result in a rotation of the acceleration vector 7, the rate of such rotation also being indicative of the strength or severity of the windshear. The apparatus of the present invention measures and utilises disturbances in the foregoing relationships as criteria for detecting the magnitude and direction of any windshear that may. surround the aircraft. These disturbances are designated A through E and are summarised as follows: A = - a v (1) B = V T - a H (2) C = d / dt ( - a v ) (3) D = d dt ( V T - a H ) (4) E = |A VH | ρ (5) In order to use available on-board sensors, it is convenient to express equation 5, in terms of a V and a H in equation 12 below, Its derivation is as follows, referring to Figure 2: A 2 / VH = a 2 / V + a 2 / H (6) Solving for ρ ˙ in equation 9 and substituting equation 8 therein: Now substituting equation 6 into equation 11 and simplifying:- The above derivation of equation 12 provides a measurement of the disturbance parameter |A VH |ρ ˙ from direct measurements of linear accelerations acting on the aircraft and available from the same horizontal and vertical accelerometers strapped to the airframe as used in equations 1 through 4. A preferred embodiment of the accelerometer configuration for use in determining the a V and a H components of the total acceleration vector A VH is illustrated in Figure 3. In order to improve the clarity of the description, Figure 3 has been drawn to illustrate the special case where the bank angle of the aircraft is zero; that is, wings level flight which is the usual approach and take-off attitude at lower altitudes. The relationships derived for a V and A H , however, do include the effects of bank angle in accordance with well known principles. The aircraft bank angle is designated  in this discussion. The accelerometer configuration comprises a conventional force balance linear accelerometer 36, mounted with its sensing axis aligned with the Z axis 21 of the aircraft and a toroidal liquid level, or acceleration sensor 22, mounted in the X-Z plane of the aircraft, having an electrical null when the liquid surface 23 is parallel with the longitudinal X axis l0 of the aircraft. The toroid accelerometer may be of the type disclosed in the present Applicants' U.S. Patent Specifications Nos, 4,028,8l5; 3,823,486; 3,604,275; and 3,l7l,2l3. Under the conditions shown in Figure 2, the aircraft's X axis l0 has an attitude ϑ with respect to the horizontal plane ll which is normal to the vertical line 9 (see also Figure 2,), which extends from the centre of the toroid to the centre of the earth. The angle is measured in a conventional manner by a vertical gyro, also mounted in the aircraft. The Z reference line 2l is orthogonal to the X reference line l0, both of which are fixed to the vertical and longitudinal axes, respectively, of the fuselage of the aircraft. The normal accelerometer 36 is mounted with its sensing axis parallel to the aircraft Z axis 2l and responds to the apparent acceleration component along the Z axis in the X-Z plane of the aircraft. The apparent acceleration component 24, designated a Z in the diagram, is the sum of the true normal acceleration component 25, designated and the component 26, designed g Z , of the earth gravity vector that exists in the longitudinal plane of the aircraft X-Z. The earth gravity vector g is, of course, always aligned with the vertical reference line 9. The gravity vector is resolved into orthogonal components, projected onto the X, Y and Z axes of the aircraft. The lateral of Y axis has been omitted from Figure 3 in order to improve its clarity. It can be considered, however, to be a line which extends from the centre of the toroid at right angles to the plane of the diagram of Figure 3. The components of the earth gravity vector are well known and can be expressed by the following general relationships which include the effect of bank angle or non-level flight: g X = g sin ϑ l3 g Y = g cos ϑ sin  l4 g Z = g cos ϑ cos  l5 where ϑ and  are the aircraft's pitch and roll angles respectively and measured by the vertical gyro in the aircraft in conventional fashion. The apparent acceleration vector 27 in the longitudinal plane of the aircraft is designat A XZ . For the case where bank angle  is zero, it is the vector sum of the true acceleration vector 7, designated A VH , and the gravity vector 28, designated g XZ , existing in the XZ plane. The components of vector 27 in aircraft Z and X axes are 24 and 29, designated a Z and a X , respectively. Using relationships l3 and l5, the value of these vectors are expressed as a x = g x + = g sin ϑ + l6 a z = g z + = g cos ϑ cos  + l7 where the earth gravity components g Z and g X are vectors 26 and 3l, respectively, and where and are vectors 30 and 25 respectively, which are the components of the true acceleration vector 7 along the X and Z axes, respectively, of the aircraft. The angle λ in the diagram is the angle between the acceleration vector 27 and the Z axis 24 of the aircraft, it being equal to the angle between the aircraft X axis l0 and the surface 23 of the fluid in the toroid accelerometer 22 since the liquid surface will always be normal to the apparent acceleration vector 27. Therefore, it follows that the electrical output of the toroid accelerometer 22 is proportional to the angle λ because the toroid is positioned in the aircraft to have an electrical null when the liquid surface 23 is parallel with the X axis l0 of the aircraft. The angle λ is related to the components of the apparent accleration vector 27 as follows: Using the relationships l3 through l8, the inertial vertical acceleration component 32, designated a V , and the inertial horizontal acceleration component 33, designated a H , can be expressed in a general form (which also includes the effect of bank angle for completeness): a v = cos ϑ cos  + cos ϑ sin  + sin ϑ l9 a H = cos ϑ - sin ϑ cos  20 where: = a z - g cos ϑ cos  2l = a y - g cos ϑ sin  22 = a x - g sin ϑ 23 However, in a coordinated turn, it can be assumed that: a y = O 24 = - g cos ϑ sin  25 Therefore, relationships l9 and 20 can now be expressed: a v = a z (cos ϑ cos  + tan λ sin ϑ ) - g 26 a H = a z (tan λ cos ϑ - sin ϑ cos  ) - g sin ϑ cos ϑ sin ² 27 The relationships 26 and 27 involve direct measurements from the normal accelerometer 36 (a z ), the toroidal liquid level 22 (λ) and a vertical gyro 34 (ϑ,). Referring now to Figure 4, a block diagram of a preferred embodiment of the windshear detection, warning and correction system is illustrated. This system may be an integral part of an aircraft performance management system (PMS) and preferably constitutes a separate digital computer subroutine of the PMS. The vertical gyro 34, toroidal accelerometer 35 and normal accelerometer 36, having respective signal outputs 37, 38, 39, 40, are processed by an acceleration component computer 4l to provide a V and a H output signals 42, 43 in accordance with equations 26 and 27, previously defined. The component computer or data processor 4l for solving equations 26 and 27 is completely conventional and may be readily implemented by those skilled in this art using well known digital computer component conversion subroutines. The outputs 42 and 43 of the component computer 4l are the vertical and horizontal accelerations of the aircraft relative to the earth coordinates, i.e. relative to the ground. Outputs 42, 43 in conjunction with a true airspeed output 45, designated V T , and vertical speed output 46, designated h, of an air data computer 47, together with flap position output 48 (designated δ FL ) of a flap position sensor 49 are processed by a windshear computer and detector 44 to derive the disturbance relationships l through 5, defined previously. The details of the windshear computer and detector and warning signal generator 44 will be described hereinbelow with reference to Figures 6 and 7. As will be described there, the disturbance relationships are combined in a manner to provide a two-level alert signal. The first level represents a moderate windshear condition and is manifested as a steady alert signal 50 which illuminates a warning light on a windshear annuniciator 52. The second level represents a severe windshear condition and is manifested as a flashing illumination of the warning light by an alert signal 5l. A corresponding audible warning may also preferably be provided. The pilot will acknowledge the alert and warning at his discretion and operate a palm switch 54 which, in most large aircraft, is located on the throttle control levers and normally used to initiate a missed approach manoeuvre. The palm switch 54 latches a relay 55, having contactors 56 and 57. The function of contactor 46 is to transfer the autothrottle system 58 to operate at maximum thrust command 59 in lieu of the normal approach thrust command 60, which was the situation shown in Figure l. The function of contactor 57 is to transfer the pitch flight director 6l to operate on an angle of attack director signal 62 in lieu of the normal approach director signal 63, which was the situation shown in Figure l. Thus, in accordance with the present invention, not only is a warning given to the pilot when the aircraft experiences a windshear condition, it also provides guidance information to the pilot to enable him to manoeuvre the aircraft promptly and safely to avoid a dangerous situation. Upon actuation of the palm switch 54, an alpha function computer 64, to be described more fully below, supplies a command signal which will direct the pilot to pitch the aircraft to its maximum safe angle of attack. Normally, this angle of attack is approximately that of the stick shaker angle of attack, that is just below the stall angle of attack. The alpha function computer also provides an angle of attack deviation signal to a fast-slow indicator 68 to provide an analogue display which shows the actual aircraft situation relative to the stick shaker angle of attack. It will now be appreciated that the aircraft will be commanded to its maximum lift configuration and avoid the windshear problem discussed with respect to Figure l. It is important to note that by issuing flight director command and autothrottle commands automatically, the pilot will be guided to follow an abnormal procedure rather than his normal instincts as described previously. The alpha function computer 64 is used to process a signal 65 from a local flow alpha vane 66 and the flap position sensor output 48 to generate a true angle of attack value. The angle of attack signal 67 operates into a fast-slow indicator 68, as shown in Figures 5A and 5B. The alpha function generator is illustrated in block diagram format in Figure 9. As shown, it is responsive to a signal on a lead 65 proportional to the aircraft angle of attack α V , usually provided by a conventional vane sensor 66. It is also responsive to a signal on lead 58 proportional to the position of the flaps δ FL . The details of the angle of attack function computer 64 are shown in Figure 7. Basically, the computer 64 provides a signal proportional to the error between the actual angle of attack of the aircraft and a centre reference angle of attack for the particular mode of operation of the aircraft as shown in Figure 5B for display, together with a peak angle of attack reference for use as an angle of attack command for the flight director. The angle of attack vane signal on lead 65 is passed through a suitable gain 200 and provides on a lead 20l a signal proportional to the true angle of attack or alpha of the aircraft. The flap angle signal on lead 48 is supplied to a source of reference angle of attack signal depending upon the flight mode. It will be recalled in the discussion of Figure 5B above that these reference angle of attack signals are a function of flap position and whether the aircraft is in an all engine TOGA mode, or in an approach mode, and that these reference alpha signals correspond to a centred position of the fast/slow indicator 68. The discretes TOGA, ENG OUT, TOGA, and APPR are supplied conventionally from the mode initiation switches which in turn supply reference alpha voltages on a lead 2l6 in accordance with the curves shown in Figure 5B. These reference voltages are identified as αCENT since they represent the centre position of the fast/slow indicator 68. The alpha reference signals are compared with the actual alpha signal on a lead 20l at a junction 2l7 to provide an alpha error signal Δα on a lead 230. This is the signal supplied to the fast/slow indicator 68 via lead 67. The alpha error signal on the lead 230 is supplied also to the flight director pitch command cue after being conditioned for flight director operation. This simply involves the addition of a damping term from a rate taker 2l0 and conventional limiting by means of a limiter 222. The final flight director alpha command signal appears on lead 62 which energises a conventional flight director visual cue. The fast/slow indicator 68 shown on Figure 5A comprises a cue 69 which moves vertically against a fixed scale having an upper index 70, marked F, a centre index 7l, and a lower index 72, marked S. The characteristics of the angle of attack signal 67 (Figure 4) are such that the moving cue 69 will be positioned against the F index 70 when the true fuselage angle of attack of the aircraft is zero degrees. The cue will be positioned at the S index 72 when the true fuselage angle of attack is at a value corresponding to the stick shaker threshold. As stated above, the stick shaker is designed to provide an alert to the flight crew that the aircraft is nearing a stall condition. When the moving cue 69 is adjacent the centre index 7l, it indicates proper angle of attack for three separate conditions, designated 74, 75, 76 in the graphical plot of Figure 5B which represents reference fuselage angle of attack values as functions of flap position. These are the stick shaker reference 72, the zero angle of attack reference 70, and the three centre conditions which are all-engine take-off/go-around (TOGA) 75, the failed engine TOGA 74, and the normal approach condition 76. The centre values 74, 75, 76 serve as a valuable cross check on the validity of the signals since the flight crew can compare the airspeeds that exist when the angle of attack display is centred under normal conditions. Figure 6 is a preferred embodiment of the windshear computer and detector 44 of Figure 4. The sensor input signals to the windshear computer are 42, 43, 45, 46 as depicted in Figure 4 and repeated in Figure 6. Filters 69, 70, 71, 72 operate on the sensor signals 42, 43, 45, 46 to reduce the spurious noise that is a general characteristic of accelerometer and air data signals. The time constants of filters 69, 70, 71, 72 are all preferably identical in order to maintain a proper dynamic relationship between the signals. The filtered a V and a H signals 42 and 43 are subjected to washouts filters 73 and 74 in order to remove any steady state stand-off signals which are also characteristic of accelerometer signals. The time constants of the washouts filters 73 and 74 are again preferably identical in order to maintain a proper dynamic relationship between the a V and a H signals. The washed out and filtered a V and a H signals are processed by a data processor 75 to generate an output signal 76 which represents the solution to the mathematical relationship 12, discussed above, which is the vector 18 (Figure 2) and is the vector product of A VH and ρ ˙. A block diagram of the data processor 75 is shown in Figure 8. It will be appreciated that the data processor conventionally implements equation 12 and a detailed description thereof would be superfluous. Normal extensions and retractions of aircraft flaps will result in sizeable acceleration vectors which must be compensated in order to avoid nuisance trips of the alert of warning signals. This is obtained by a variable gain block 77 which responds to the signal on lead 48 proportional to movement of the flap position sensor 49, Figure 4. This is processed through a rate taker 78 and absolute value detector 79 to generate a variable gain characteristic shown by the graph 80. Graph 80 shows that movement of the flaps results in a reduction in the gain of signal 76 which gradually recovers with time. A typical time period between the movement of the flaps and the recovery of substantially the full signal may be on the order of 60 seconds. The output of the variable gain 77 is filtered by smoother filter 81 before entrance into level detector E where a high logic output is provided on a terminal 83 when the output of the filter 81 exceeds a preset negative value. Level detector 82 also provides a similar high logic signal for a positive signal. The logic signals on terminals 83, 84 reflect disturbance relationship 5 above. The filtered vertical speed signal 46 is operated on by a washout filter 89 and a rate taker 85 to generate an air data derived acceleration signal 86, designated , which is the aircraft vertical acceleration relative to the air mass. This is compared with the filtered inertial acceleration (a v ) signal 87 at a summing junction 88 and then filtered at 90. The output from filter 90 is entered into 91, level detector A, where a high logic output is provided on a terminal 92 when the output of filter 90 exceeds a preset positive or negative value. The logic signal on the terminal 92 reflects the disturbance relationship 1 above. The filtered true airspeed signal is operated on by a washout filter 97 and a rate taker 93 to generate an air data derived acceleration signal 94, designated V ˙ T , which is the aircraft forward acceleration relative to the air mass. This is compared with the filtered inertial acceleration (a H ) signal 95 at a summing junction 96 and then filtered at 98. The output from the filter 98 is entered into 99, level detector B, where a high logic output is provided on a terminal 100 when the output of filter 98 exceeds a preset positive or negative value. The logic signal on the terminal 100 reflects the disturbance relationship 2 above. A signal 101 represents the processed signal - a V and this signal is further processed by a rate taker 102 and filter 103 prior to entrance into 104, level detector C, where high logic outputs are provided on terminals 105 and 106. Terminal 105 represents a negative disturbance; terminal 106 represents a positive disturbance in accordance with relationship 3 above, both greater than a predetermined threshold value. A signal 107 represents the processed signal V ˙ T - a H and this signal is further processed by a rate taker 108 and filter 109 prior to entrance into 110, level detector D, where high logic outputs are provided on terminals 111 and 112. Terminal 111 represents a negative disturbance; terminal 112 represents a positive disturbance in accordance with relationship 4 above, both greater than a predetermined threshold value. Two levels of alert are provided; one representing a moderate windshear condition, the other representing a severe windshear condition. Referring to Figure 9, as discussed previously, logic levels indicative of windshear appear on terminals 83, 94, 111, 112, 105, 106, 100 and 92 of Figure 6. The logic levels from terminal B, 100, and terminal A, 92 are supplied to a conventional logic OR gate 123 via leads 119 and 120, respectively. Thus, the output of the OR gate 123 will be a high logic signal if either or both leads 119 and 120 has a high logic signal. The output of the gate l23 is impressed on a counter l24 which will output a high logic level signal if the output of the gate l23 remains at a high logic level for a predetermined length of time, for example three seconds. Otherwise, the output of the counter l24 is maintained at a low logic signal. The output of the counter l24 appears at a junction l25, one lead from which supplies a timer l26 which will ouput a high logic signal for a predetermined amount of time, for example 20 seconds. If the output of the counter l24 is ever at a high logic level, even momentarily, the output of the timer l26 appears on a lead l28 which is impressed as one input to conventional logic AND gates l3l, l32, l33 and l34. The logic signal from terminal 83 appears on a lead ll3 and is impressed as one input to logic AND gate l3l. The output of the gate l3l appearing on a lead l35 can be a high logic level signal if, and only if, the logic levels on leads ll3 and l28 are high logic level signals. The logic signal levels appearing at terminals lll and l05 are impressed on a conventional logic OR gate l2l. The output of the gate l2l on a lead l30 will thus be a high logic signal if either or both leads ll5 and ll7 has a high logic signal. The logic signal level on the lead l30 is impressed on a conventional logic AND gate l32. The output of the gate l32 appearing on a lead l36 can thus be a high logic level signal if, and only if, the logic levels on leads l30 and l28 are high logic level signals. The logic levels appearing on leads l35 and l36, discussed above, are impressed on a conventional logic OR gate l37. The output of the gate l37 appears at a junction l4l, one lead from which is impressed on a timer l43 which will output a high level logic signal for a predetermined time, for example l0 seconds, if a high logic level signal appears at junction l4l, even momentarily. If, at the completion of the predetermined time, a high logic level signal remains on the lead l4l, the timer l43 will again output a high logic level signal for the predetermined time. Otherwise, the output of the timer l43 will become a low logic level signal. The output of the timer l43 appears at a junction l44 and thence on the lead 5l which is supplied to the windshear annunciator 52. A high logic level signal on lead 5l causes a flashing windshear annunciation. The logic level signal on the terminal 84 appears on the lead ll4 and is thence supplied as one input to a conventional logic AND gate l33. The output of the gate l33 appears on a lead l38 and will be a high logic level signal if, and only if, the logic level signals on leads ll4 and l28 are high logic level signals. The logic level signals from the terminals ll2 and l06 appear on leads ll6 and ll8, respectively, and are thence supplied as inputs to a conventional logic OR gate l22. The output of the gate l22 appears on a lead l29. The logic level signal on the lead l29 will be a high logic level signal if either or both leads ll7 and ll8 has a high logic level signal. The logic level signal on the lead l29 is supplied as one input to a conventional logic AND gate l34. The output of the gate l34, appearing on lead l39, will be a high logic level signal if, and only if, the logic level signals on leads l29 and l28 are high logic level signals. A conventional logic OR gate l40 receives as inputs the signals on leads l38, l39 and l27. The output of the gate l40 will be a high logic level signal if any one, or any combination, of the inputs on the aforementioned leads is a high logic level signal. The output of the gate l40 appears on a lead l48 and is thence impressed on a timer l49 which operates in a manner similar to timer l43, discussed above. A high logic level signal output from the timer l49, on lead 50, will cause a steady windshear annunciation. The logic level appearing at the junction l44 and on the lead l45 is impressed on the timer l49 such that a high logic level signal will cause an instantaneous low logic level signal on lead 50 regardless of the value of the logic level signal on lead l48 or the amount of elapsed time of the timer l49. This action allows the flashing windshear annunciation to take precedence over the steady windshear annunciation should both events occur simultaneously. The windshear detection logic provides the following features: (A) the moderate windshear alert is generated when the input to the detector A or the input to the detector B rise above positive or negative preset trip levels for a predetermined amount of time. The moderate windshear alert remains active until the input to the detector A and the input to the detector B both drop below the trip level. (B) The detectors C, D and E are armed to operate for a time interval which starts when the moderate windshear alert is actuated and ends a predetermined amount of time after the inputs to the detectors A and B have fallen below the preset trip levels. (C) When armed, positive polarity trips of the detectors C, D and E will also operate the moderate windshear alert annunciation. (D) When armed, negative polarity trips of the detectors C, D, and E will operate the severe windshear alert annunciation. Clearly, it is desirous to minimise or eliminate nuisance or false warnings of windshear to the flight crew. The important aspects of windshear are both the magnitude and duration of the shear. For example, common turbulence may create a change in wind speed sufficient to cause large excursions in aircraft true airspeed, but the duration of the change is so short as to pose no serious threat to safety of flight. Conversely, windshear can not only cause large airspeed excursions but can endure sufficiently long enough to endanger the aircraft. Hence, an important part of windshear detection is consideration of both the magnitude and length of time of the windshear. A further preferred embodiment of the present invention incorporates the determination of both the magnitude of the shear and the allowable time the shear can be tolerated at that magnitude. A description of the technique used follows. Referring again to Figure 6, the action of the detector 9l is replaced by the circuit of Figure l0. The signal on the lead lOl of Figure 6 is supplied to a threshold detector l50 operable to output a logical ONE on a lead l5l if the signal on the lead lOl exceeds a predetermined value, for example O.O2 g, whether positive or negative in sign, where g has a value of 32 feet per second per second (9.75 metres per second per second). Detector l5O outputs a logical ZERO for values less than the predetermined value. The output of the threshold detector l5O is supplied to a timer l52. The output of the timer l52 appears on a lead l53 and is a signal proportional in magnitude to the elapsed time since the signal on the lead l5l was triggered to become a logical ONE. The lead l5l also supplies a lead l54 and thereby provides the reset function of the timer l52, the circle terminating the lead l54 representing a logical NOT function. By the described action, the timer l52 will output a continous signal proportional to elapsed time on a lead l53 whenever the signal on the lead l5l is a logical ONE and will output a null whenever the signal on the lead l5l is a logical ZERO. Simultaneously, the signal on the lead lOl, representing the longitudinal wind rate, appears on a lead 155 and thereby is applied as the denominator to a conventional divider 156. A constant value K, which is representative of the maximum allowable time the windshear can be tolerated and has, for example, a value of 0.237 9-seconds, appears on a lead 157 and is supplied as the numerator for the divider 156. The divider 156 produces the quotient of the value of the signals on the lead 155 and 157, which quotient is applied to a lead 158. The signal on the lead 158 is representative of the allowable time that the windshear may be endured at the current measured magnitude. It will be seen that the signal has units of seconds when the magnitude of the windshear is measured in gravitational units (G). The signal on the lead 158 thereby defines curve 208 in Figure 11. The lead 158 is coupled to a conventional absolute value taker 159 which operable to provide a positive value signal on a lead 160 equal in magnitude to the value on the lead 158. The output of the absolute value taker 159 appears on a lead 160 and is applied to a conventional summation device 161. The summation device 161 receives the signals on the lead 153 and 160 and produces the algebraic sum in the sense of a numerical difference on a lead 162. The signal on the lead 162 thus represents the difference between the actual elapsed time of the wind rate condition on the lead 153 and the computed allowable time which appears on the lead 160, this signal being applied on the lead 162 to the polarity detector 163. The polarity detector 163 operates in such a fashion as to output a logic ONE on a lead 164 to terminal A (92 of rig. 6) whenever the signal on the lead 162 is positive in sign; that is, whenever the actual elapsed time exceeds the computed allowable time. For a null or negative signal on the lead 162, the output of polarity detector 163 will be a logic ZERO. The logic signal on the lead 164 and terminal A reflects the vertical windshear disturbance relationship of equation (l), discussed above. This signal may be used to activate a visible and/or audible alarm, as in the present invention, or to enable a windshear guidance system as in EP-A-0 235 964. It will be clear to one skilled in the art that a similar circuit may be employed in place of the level detector 99 of Figure 6, so that both vertical and longitudinal windshear components may be detected whilst minimising false alarms. In that event, the counter l24 and timer l26 become redundant and may be deleted from the circuit of Figure 9. Figure ll is illustrative of the operation of the level detector 9l of Figure l0, where the ordinate axis 203 is wind rate in gravitational units (g = 32 ft/sec/sec -9.75 metres/sec/sec) and the abscissa 202 is actual time in seconds. Assume at some point A the wind rate, line 204, exceeds the trip value of the threshold detector l50. At this point the timer l52 is activated, represented by the vertical line 206. As the wind rate continues to exceed the value of the threshold detector, the elapsed time from the line 206 increases. When the elapsed time exceeds the computed allowable time on the line l60 of Figure l0, which is a function of the magnitude of the wind rate and is represented, for example, by point B of Figure ll, a logic ONE will appear at terminal A of Figure l0, indicating the presence of a wind disturbance. Point B of Figure ll is one of a locus of points 208 of computed allowable time as appears on the lead l60 of Figure l0. It will be appreciated from the foregoing that the level detector 9l of Figure l0 provides the following actions: l) The annunciation of a windshear condition is dependent on both the magnitude and duration of the wind rate. 2) For large wind rates, the allowable time before annunciation is relatively short; conversely, for small wind rates above the threshold level, annunciation does not occur for a much longer time. 3) Wind rates that are above the threshold level for elapsed times less than the computed allowable will not result in an annunciation. 4) Very high magnitude wind rates that do not exist for long periods of time, such as can be caused by atmospheric turbulence, do not cause false annunciations since they do not exceed the computed allowable time.

Claims: Apparatus for generating in an aircraft a windshear warning control signal (162), of the type having first detector means (44) for sensing the difference between aircraft inertial acceleration and acceleration with respect to the supporting airmass, and providing therefrom a difference signal consisting of a vertical (101) component, and a component (107) along the longitudinal axis of the aircraft, second detector means (150) responsive to said difference signal for providing a windshear rate signal (151) representative of a magnitude of windshear rate exceeding a predetermined value corresponding to a threshold of detection, first timer means (152) responsive to the windshear rate signal exceeding the predetermined value for providing a first timing signal (153) representative of a time duration of the rate signal, characterised in that the apparatus comprises means (156,159) for generating a second timing signal (160) representative of a maximum allowable duration that safe flight can be sustained at said windshear rate, and summing means (161) for combining the first timing signal (153) and the second timing signal (160) to obtain an algebraic difference thereof and generate the windshear warning control signal (162) when the first timing signal exceeds the second timing signal, so that the warning signal is generated after a time delay inversely proportional to the magnitude of the windshear rate signal. Apparatus according to claim 1, characterised in that it comprises : timing signal means (K) for providing a signal representative of the maximum allowable time a windshear can be tolerated, and divider means (156) responsive to said timing signal means (K) and said vertical or longitudinal component for deriving said second timing signal. Apparatus according to claim 1 or 2, characterised in that it further comprises third detector means (163) responsive to the windshear warning control signal for generating a windshear annunciation signal (120) when the first timing signal has a value exceeding the second timing signal. Apparatus according to any of the preceding claims, characterised in that: the first detector means (150) provides a signal having a first logic condition when said predetermined value of the magnitude of windshear rate is exceeded, and a second logic condition when the predetermined value is not exceeded, in that the first timer means (152) is responsive to the first and second logic signals and indicative of a time interval corresponding to a change in state between the second and first logic conditions, and in that the divider means (156) for generating the second timing signal provides an output (158) indicative of a time duration comprised of the quotient of the signal generated by said timing signal means (K) and a signal corresponding to said vertical or longitudinal component. Apparatus according to any of the preceding claims, characterised in that the windshear rate signal comprises vertical components of a windshear disturbance. Apparatus according to any of the preceding claims, characterised in that the windshear rate signal comprises longitudinal components of a windshear disturbance. Apparatus according to any of the preceding claims, characterised in that the windshear rate signal comprises vertical and longitudinal components of a total acceleration vector representative of the resultant vectorial sum of the vertical and horizontal inertial acceleration components of the aircraft with respect to the supporting airmass. A method of generating a windshear warning control signal in an aircraft, characterised in that it comprises the steps of: providing a signal corresponding to a rate component of the windshear disturbance, detecting the rate component and providing an output signal when the rate exceeds a predetermined threshold, applying the output signal to a timer for providing a signal indicative of a duration thereof, providing a signal representative of an aircraft performance parameter indicative of the maximum allowable time the windshear can be tolerated, applying the rate component and the parameter signal to a divider for providing a quotient signal thereof, applying the quotient signal to a means for deriving a signal indicative of the absolute value thereof, combining the duration signal and the absolute magnitude signal and deriving an algebraic difference thereof, and generating said windshear warning control signal in response to a given polarity of the algebraic difference indicating that the value of the duration signal exceeds the absolute value of the quotient signal, so that the warning signal is generated after a time delay inversely proportional to the magnitude of said rate component.

Inventors: MILLER, HARRY, ZWEIFEL, TERRY L.
Applicants: HONEYWELL INC.
Application year: 1987
Main classification: G05D1
IPC codes: G01P15, G01W1, G01P5, G05D1, G01C23, B64D45