GB2094006A - Windshear detection and warning system - Google Patents
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- 230000002441 reversible effect Effects 0.000 claims abstract description 12
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- 230000001419 dependent effect Effects 0.000 claims description 7
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/04—Control of altitude or depth
- G05D1/06—Rate of change of altitude or depth
- G05D1/0607—Rate of change of altitude or depth specially adapted for aircraft
- G05D1/0615—Rate of change of altitude or depth specially adapted for aircraft to counteract a perturbation, e.g. gust of wind
- G05D1/063—Rate of change of altitude or depth specially adapted for aircraft to counteract a perturbation, e.g. gust of wind by acting on the motors
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- G—PHYSICS
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- G06G—ANALOGUE COMPUTERS
- G06G7/00—Devices in which the computing operation is performed by varying electric or magnetic quantities
- G06G7/48—Analogue computers for specific processes, systems or devices, e.g. simulators
- G06G7/70—Analogue computers for specific processes, systems or devices, e.g. simulators for vehicles, e.g. to determine permissible loading of ships, centre of gravity, necessary fuel
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Abstract
A windshear detection and warning system based on evaluating the effects of the rate of change of the total reversible energy that resides in a moving aircraft, which energy is the sum of its potential and kinetic energy. The sources of energy change comprise thrust minus drag multiplied by the velocity vector and variations of the wind. When the wind does not vary, the thrust minus drag divided by weight is exactly related to the potential flight path angle of the aircraft which is the sum of actual flight path angle and the acceleration along the flight path. This relationship is disrupted in the presence of wind variations. Means are provided to detect the magnitude of this disruption and use it as a basis for detecting the presence of windshear. <IMAGE>
Description
SPECIFICATION
Windshear detection and warning system
The present invention relates generally to aircraft flight performance computer systems and more particularly to a method and a system for providing a warning and/or flight control signal indicative of impending dangerous windshear conditions.
Windshear is a weather condition which results in a change in wind velocity and/or direction (usually experienced at varying elevations) and in terms of aircraft flight performance is most dangerous during landing approaches along predetermined flight paths defined by ground-fixed references, but it may also be bothersome during other flight regimes, such as take-off, climb, descent and go-around where predetermined ground reference flight paths (such as those defined by a flight management system) are desired to be followed. A number of attempts and proposals have been made in the past to provide the pilot of an aircraft during a landing approach with a warning of impending windshear conditions.Most of these have involved some means for detecting changes in ground speed such as by using a ground based reference, for example distance measuring equipment (DME), or by using an airspeed transducer and measuring the rate of change of its output augmented by longitudinal acceleration to provide an inertial component related to the earth. Other proposals additionally employed vertical accelerometers to provide measures of craft vertical motion produced by the effects of windshear. Thus, prior art windshear detection and warning systems proposed using a direct measure of ground speed from a ground based reference or such prior art systems were dependent upon an airpseed sensor as a primary windshear information source. The foregoing techniques are generally direct or "brute force" approaches to providing windshear detection and warning as standalone systems.
According to one aspect of the present invention there is provided a method of determining the magnitude of a windshear condition encountered by an airborne aircraft comprising the steps of:
a) establishing from the total reversible energy contained in the aircraft under unchanging wind conditions a balance between the ratio of the aircraft thrust minus dragto its gross weight and the potential flight path angle of the aircraft, and then,
b) determining the increase or decrease in the total reversible energy contained in the aircraft induced by changing wind conditions by detecting any change in said balance.
According to another aspect of the present invention there is provided apparatus for detecting the magnitude of a windshear condition encountered by an airborne aircraft comprising means responsive to the thrust, drag and weight of the aircraft for providing a first component signal corresponding to the ratio of the thrust minus drag to the weight, means responsive to the longitudinal and normal acceleration of the aircraft and a function of the aircraft angle of attack for providing a second component signal corresponding to the potential flight-path angle of the aircraft and means for combining said first and second component signals in such a manner as to provide a resultant signal corresponding to any difference therebetween.
According to a still further aspect of the present invention there is provided apparatus for detecting the existence of a windshear condition encountered by an airborne aircraft comprising accelerometer means for providing a first signal proportional to the acceleration of the aircraft along its longitudinal axis and a second signal proportional to the acceleration of the aircraft along its vertical axis, means for providing a third signal proportional to the angle of attack of the aircraft, computation means responsive to said first, second and third signals for providing a component measure of the potential flight path angle of the aircraft and responsive to said second and third signals for providing a component measure of the drag characteristic of the aircraft, means for providing a component measure of the thrustto-weight ratio of the aircraft, and combining means responsive to the algebraic sum of all of said component measures for providing an output signal corresponding to the magnitude of the windshear encountered by the aircraft.
The present invention thus provides a more fudamental approach to providing detection and indication of impending windshear conditions; one based on the fundamental airborne aerodynamic characteristics of the particular aircraft type involved. It is based on the total reversible energy residing in the aircraft at any instant as it proceeds along a desired or predetermined flight path and is particularly valuable as the aircraft proceeds along a predetermined approach path to a landing runway where unexpected severe windshear effects may be disastrous.
The reversible energy residing in an aircraft is the sum of its potential energy and its kinetic energy.
The rate of change of this total energy per unit of aircraft weight divided by the velocity of the aircraft determines the characteristics of the flight path along which the aircraft will travel; i.e., its potential flight path angle. The sources of reversible power input to or output from the aircraft are: a) the product of the difference between thrust and drag times velocity and b) wind variations. Under zero wind change conditions, the ratio Qf thrust minus drag to aircraft weight and potential flight path angle are in balance. However, if the aircraft encounters any wind change, this energy equation is unbalanced.
The present windshear detection and warning computer system effectively measures this energy unbalance and provides a warning dependent upon the severity of the unbalance; that is, indicative of impending potentially dangerous windshear conditions.
A windshear detection and warning system for an aircraft in accordance with the present invention will now be described in greater detail, by way of example, with reference to the accompanying drawings, in which:
Figures la and ib together represent a block diagram of the system,
Figure 2 is a diagram of the effective forces acting
on an aircraft in its longitudinal plane and the resultant aircraft flight path, and useful in understanding the derivation of certain of the relationships perti
nent to the invention; and
Figure 3 is a nomograph illustrating the derivation of the angle of attack dependent function illustrated in Figure 4.
While the present invention may be implemented using an analog or digital computer based stand alone product system, it may more advantageously be incorporated in an overall digital computer based aircraft flight performance andlor flight management system for controlling or regulating the entire vertical flight path profile mission of the aircraft in the most economic and safest manner Included in such an overall system, the carrying out of the pres entinvention can take advantage of the many common system inputs it requires as well as many common flight and aerodynamic parameters of a particular aircraft type included in the overall computer memory data base.
Before discussing a preferred embodiment ofthe windshear detection and warning system of the present invention, the fundamental mathematical relationships upon which the invention is based will be described.
The reversible energy that resides in an aircraft is the sum of its potential energy and its kinetic energy: E=Wh+mV2 (1) wherein
E = total reversible energy
W = aircraft weight
H = aircraft altitude
m = aircraft mass
v = aircraft velocity
which reduces to V2
E=(h±)W (2)
g The rate of change of the reversible energy content, i.e. its power per unit of weight is: E = (h + VV) = (h + V) V (3)
W g g Since sin y = h where,' is the aircraft's flight path
V angle,
E = (siny + V)V (4) g There are two sources of reversible power input to or power output from the aircraft:
a) (T - D)V; where T = Thrust and D = Drag
b) wind variations
The balance between (T-C) and potential flight V
path angle (sin y + V ) is derived as follows, reference g being made to Figure 2. From this diagram of the forces acting on an airborne aircraft it is evident that T cos α - D - W sin γ = WV (5)
g or = W (x cos a - z sin a) g where x and z are the true acceleration compo- nents along the x and z axes of the aircraft.
Accelerometers mounted on the x and z axes will
measure these components mixed with components
of the earth's gravityfield, namely g sin #, and g cos ss cos y, where 6 is pitch angle and # is roll angle.
Thus:
x = ax - g sin # z = az - cos # cos# The flight path angle (y) is related to angle of attack (a), pitch (ss), and roll (#) by the following well known
relationship:
sin y = cos asian ss - sin acos # cos # Since ais normally a relatively small angle, it can be assumed that cos a is approximately equal to unity.
Equation (5) may therefore be rewritten as T - D = V + sin γ = ax y gsina (6)
W g- g Sln g based on no wind variations.
If the left-hand side of this relationship does not
equal the right-hand side, the magnitude of the dif
ference represents the effect due to wind variations.
Afunction, A, can now be defined which expresses
the unbalance. Thus: T - D ~ ax + aZsin a (7)
W g g This relationship representing the mathematical
basis of the present invention may be implemented
using on-board sensors of thrust, drag, weight, vertical and longitudinal acceleration and angle of
attack, and the computations required thereby performed by analogue or digital computer apparatus to derive a measure A of the magnitude of the windshear. However, while the terms T, W; ax, q, and sin a
are readily available from sensors, the drag term D is
not readily capable of being measured directly. This term of the basic equation (7) may be derived by further analysis of the aerodynamic characteristics
of the particular aircraft in which the invention is
installed (as obtained from the aircraft manufacturer).
D = CD q S (8)
where CD = aircraft'scoefficientof
drag
q = dynamic pressure
S = aricraftwingarea; L = (aZ)W = CLqS (9) g where CL = aircraft's coefficient of lift.
Solving equation (9) for qS and substituting in (8) D = (CD) /-ZwW (10)
g Now substituting this definition of D into the basic
equation (7) and simplifying yields a result which is
readily implemented computationally as will be described below:
T aX aZ CD A = - - - - - (- - sino) (11) W g g CL
Referring now to Figure 1, there is disclosed in
block diagram format a preferred embodiment of
computer apparatus comprising a combination of
interconnected discrete hardware elements which cooperate to provide an output signal indicative of the magnitude of changing winds being encountered by the aircraft.Also included are further ele
ments for assessing the magnitude of such wind
shear relative to what is determined from a practical
standpoint, to constitute an impending dangerous or
hazardous flight condition and to supply a warning thereof to the human pilot. Also provided are ele
ments for automatically biasing the aircraft automatic throttle control system to reduce throttle activity
during such turbulent flight conditions. The inter
connected hardware elements may be embodied as
part of a complete aircraft performance manage
ment system or as a stand alone windshear detector
and warning system and in either configuration the
computations and certain dependent parameter
generation functions defined by these elements may
be performed by either analog or digital mechan
isms.
Referring again to Figure 1, the sources of the var
ious input parameters to the system are shown
along the left-hand side of Figure la, the elements
performing the computation functions are in the centre of the combined Figures la and 1b, and the
output elements are at the right-hand side of Figure lb. If the computations are to be performed digitally, the central elements may comprise part of a conventional program controlled digital computer including conventional I/O section, processor section, memory section and control section, all readily implemented
by those skilled in the digital computer art in accordance with the teachings of the present invention.
Similarly, the present invention may be implemented by conventional analogue elements for performing the indicated arithmetic and certain
dependent parameter generation functions.
Conventional linear accelerometers 10 and 11
mounted on the aircraft so as to sense craft accelerations along its longitudinal (x) axis and vertical (z)
axis provide electrical signals proportional to a2 and az, respectively, and therefore constitute a means for
providing signals proportional to the longitudinal and vertical acceleration of the aircraft. Flap position sensor 12, which may comprise one or more synchros connected to measure flap deflection, provide a resultant electrical signal proportional to flap position and thus constitutes a means for providing a signal proportional to flap position.An angle of attack sensor 13, such as a conventional vane orser- voed airstream probe, provides a signal proportional to the aircraft's airborne angle of attack and constitutes a means for providing a signal proportional to the aircraft's angle of attack. The latter two signals are used to generate drag and left related parameters dependent thereon as will be described below.
Also, as will be further described below, a measure of the thrust of a turbojet engine can be obtained by parameters of engine pressure ratio (EPR), Mach number and static pressure ratio. It can also be obtained by measurements of engine fan speed (N1),
Mach number, static pressure ratio and air tempera- ture ratio. Thus, for an aircraft with three such engines, EPR sensors 14,15, and 16 provide electrical signals proportional to each engine's pressure ratio while a conventional air data computer 17 supplies electrical signals proportional to Mach, true airpseed and pressure altitude. It will be appreciated by those skilled in this art that the thrust parameter may be provided for any type of engine.Thus for a turbofan engine, fan speed (N1), may be used instead of EPR; for a turboprop engine, turboshaft speed and prop angle may be the functions employed. Depending upon engine type, other parameters such as total air temperature, etc., may be required to generate the thrust signal used in the present invention. In the embodiment illustrated in Figure 1, the EPR, Mach and pressure altitutde signals are supplied to a thrust computation element 18, the latter constituting a means for supplying a signal proportional to the thrust imparted to the aircraft. This element may comprise a thrust computer of the type disclosed in
Figure 8 of British Patent Specification No. 1,595,686.
The aircraft gross weight parameter may be provided by a proportional electrical signal derived from an electrical signal generator 19 manually set by the pilot in accordance with the existing gross weight of the aircraft which is conventionally available to the flight crew from independent on-board sources. Preferably, however, the aircraft's gross weight parameter may be derived continuously and automatically from a gross weight computer system 20 of the type disclosed in the above British Patent Specification
No. 1,595,686. Also, if the windshear warning system of the present invention is incorporated in an overall performance management system, this weight parameter computation is part of such system and a signal proportional to weight is readily available.It should be noted also that one of the basic terms of the windshear measuring relationship, equation (7) above, is the thrust-to-weight ratio T and that the W apparatus of Figure 5 of the above British Patent
Specification No. 1,595,686 provides a measure of,
or a signal proportional to, this ratio. Therefore, in
Figure 1 of the present application that apparatus
including the gross weight computer 20 and thrust
computer 18 may be of the type illustrated in that
patent which, together with a conventional analogue
or digital divider 21 responsive to their outputs con
stitutes a means for providing a signal proportional
to the ratio of the aircraft's thrust to its weight.
The acceleration terms of the basic windshear
measuring relationship, equation (7), require modifi
cation by the gravity constantg, specifically by
means for determining the magnitude of the air
craft's longitudinal and vertical acceleration relative
to the earth's gravitational force. As illustrated in
Figure 1, the elements for accomplishing this func
tion comprise conventional analogue or digital
divider means 22, 23, respectively, responsive to the
output signals of longitudinal orx-axis accelerome
ter 10 and vertical orz-axis accelerometer 11 and a
fixed signal 24 proportional to the magnitude of the
gravity constant.Thus the x-axis accelerometer 10,
divider 22 and gravity constant source 24 constitute
a means for providing a signal proportional to the
longitudinal acceleration of the aircraft relative to the
earth's gravity field while the z-axis accelerometer
11, divider 23 and gravity signal source 24 constitute
a means for providing a signal proportional to the
vertical acceleration of the aircraft divided by the earth's gravity field. These signals a2 and az are rep
9 9
resented on connections or leads 25 and 26 from the
dividers 22 and 23, respectively.
The relationship ( CD - sin a) of equation (7) above
CL
is a predetermined function of angle of attack and
flap position and a signal proportional to the actual
physical value of this relationship is derived from the
computer element 30 of Figure 1.
The computational techniques for deriving a
measure of this term may be similar as those dis
closed in the reference British Patent Specification No.1,595,686. Figure 3 is a nomographwhich demonstrates that this expression is a function of the
aircraft's angle of attack and flap position; that is, the
right-hand portion of Figure 3 is a family of curves
representing the actual values of CL as a function of
angle of attack for various flap angle settings, while
the left-hand portion is a family of curves represent
ing the actual values of Cn for the same flap angle
settings. These curves are available from the
designer of the particular aircraft involved.Using the
data provided by Figure 3, Figure 4 is constructed
which provides the value of the required expression ~ sin a) as a function of angle of attack and flap
L position. To those skilled in the art of analogue com putertechnology, the values of the required expres
sion illustrated in Figure 4 for the flap angles illus
trated may be determined continuously or in dis
crete steps by suitable potentiometer or resistor
networks set or selected in accordance with the
actual or measured aircraft angle of attack. To those
skilled in the art of digital computers, the generation
of the value of the desired expression is readily
determined by storing the required values for the
illustrated flap angles in a ROM or other memory
device and conventionally addressing the memory
as a function of angle of attack.Thus the computer
30 comprises a means for providing a signal propor
tional to the aircraft coefficient of drag-to-lift minus
the sine of its angle and such signal is represented
on connection or lead 31 of Figure 1.
From the foregoing, it will be appreciated that the
elements of Figure 1 thus far described provide
physical measures proportional to all of the terms of the windshear relationship (equation (7) above), and
it now remains to combine these measures in accordance with arithmetic requirements of this relationship. For this purpose, the thrust-to-weight ratio component term which appears as a signal or physi- - cal measure on connection or lead 32 is supplied to one input of a summing device 33 while the compo nent signal or physical measure of the longitudinal axis acceleration term a9x on connection or lead 25 g is supplied to one input of a summing device 34.The
component signal or physical measure representing
the last term is provided on the output connection 35
of a multiplying device 26 to which is supplied the vertical acceleration term az on connection 26, and g the component signal or measure representing the
value of ( CD - sin e) on connection 31. This signal or
CL measure is supplied to the summing device 34 where it is summed with the other input thereto to provide a resultant measure of the physical value of g + a, ~ sin a).At this point in the description of g g a preferred embodiment of the present invention, it will be noted that the drag component signal D of the relation T - D is included in the latter term, viz
W
CD az
D=(CL) (g)
g while the potential flight path angle component
signal p az - az sin α g g
is also included in this term. It will be appreciated
that these terms are readily accomplished in a cost
effective manner by electromechanical or digital
implementations, as noted above, but other compu
tational techniques may be devised by those skilled in the computer art separately to generate the T W D
W term and the potential flight path angle term p upon
which the thesis of the present invention is based.
The resultant component signal or measure on
lead or connection 35, including drag component (CCDL) (a,) and a signal (CCD ) (g) and a potential flight path angle com (2x a2
ponent signal (ax - az sin α), is supplied to the other
g g input of the summing device 33 where it is sub
tracted from the measure of WT so that the output of
device 33 is a signal or physical measure of A in accordance with equation (6), i.e. the magnitude of
the windshear being encountered by the aircraft. It
will be appreciated by those skilled in the art that the
order in which the various terms are combined is
immaterial.
Having provided a measure of the magnitude of
the windshear being encountered, it is now desirable to establish the threshold value thereof which is considered to be excessive, that is to exceed the operational capabilities ofthe autothrottle system and therefore to pose a threat to the safety of the aircraft; for example, a windshear of 10 knots per
100 feet of altitude. In orderto establish a realistic threshold value, it is convenient to consider the landing approach situation where the aircraft is being controlled to follow a predetermined landing path such as an ILS glide slope, with the automatic throttle control system maintaining a predetermined approach airpseed. It should be pointed out here that in conventional autothrottle control systems, such as disclosed in British Patent Specification No.
1,374,101, the throttle servo is controlled primarily in accordance with an airspeed term and an airspeed rate term, the latter of which may include filtered inertial terms, such as longitudinal acceleration, to improve system response and stability. However the inclusion of such inertial terms in the autothrottle system tends to compromise the system's performance in a windshear situation. For example, the autothrottle system response to an encountered windshear may actually aggravate the effect of windshear on the aircraft's flight path rather than to ameliorate it. It has been found that a well designed autothrottle system can perform adequately in windshear conditions of about 6 knots per 100 feet or lower.Therefore, if windshears in excess of 8 knots per 100 feet of altitude are encountered, the pilot should be warned so that he may continue the landing manually or execute a go-around manoeuvre.
Returning to the glide slope approach situation, the aircraft's flight path angle y is the ILS glide slope angle, normally about -2.87 degrees. Therefore, assuming still air or an unchanging wind condition,
T-D the value of ( W ) is such asto result in a potential flight path angle of -arc sin 2.87 = -0.05. Thus from equation (6) above T - D = 0.05 (12)
W
At this flight path angle the vertical descent rate in
feet per second is V sin ,', where V is true airspeed.
The elapsed time to descent 100 feet, therefore is
100
V sin 9 seconds. If a windshear exists, expressed in
knots per 100 feet, the wind change, expressed as
feet per second per 100 feet, is 1.688 times the windshear. Therefore, the acceleration along the flight path resulting from the energy input to the air
craft due to the changing wind is
V 1.688 x V sin y x windshear
g 100 x 32.2
Note that this acceleration is a function of true airspeed V.For example, assuming a typical approach velocity V = 250 ft/sec. and sin ,' = 0.05 and a windshear of -8 kts/1 00 ft: v = 1.688 x 250 x (-0.05) (-8) = 0.52 (14)
g 100 X 32.2
Thus, in a windshear of -8 kts/100 feet, an approach speed of 148 kts and a glide slope of -2.87 degrees, V . ~ a2 ~ asina (15)
V +sin
g g g or 0.052 - 0.05 = 0.002 and from equations (6), (12) and (15)
A = -0.05 - 0.002 = -0.052 (16)
Now in an opposite windshear condition, +8 kts/100 feet, the value of equation (15) would be -0.102 and
A = +0.052.
Returning now to Figure 1, and having above established a typical value of A above which a warning should be given to the pilot, the absolute value of A on connection or lead 40 from a conventional absolute value detector 41 is applied to a conventional comparator network 42 to which is also applied a reference signal corresponding to the threshold value of Afrom a reference signal generator 43. Therefore, if the absolute value of A exceeds the reference value of A, a warning signal on lead 47 results, which signal may be used to provide a visual and an audible warning to the pilot of an impending dangerous windshear condition. It will
be noted from equation (13) that the threshold bias is
a function of the aircraft's true airspeed V since the value of A corresponding to a windshear of 10 kts/1 00 feet will increase or decrease as the true airs
peed increases or decreases. For this purpose a signal proportional to true airspeed V is convention
ally supplied from the air data computer 17 via lead or connection 48 to the threshold bias generator 43 to vary this threshold value as a function of true airs
peed.
The windshear magnitude signal A may also be
supplied to the aircraft autothrottle system through a conventional wash-out circuit 45 and rectifier 46. The
resultant signal on lead 49 is used in the autothrottle
system to compensate for the aggravating effects of
its intertial terms as described above.
Claims (11)
1. A method of determining the magnitude of a
windshear condition encountered by an airborne
aircraft comprising the step of:
a) establishing from the total reversible energy
contained in the aircraft under unchanging wind
conditions a balance between the ratio of the aircraft
thrust minus drag to its gross weight and the poten
tial flight path angle of the aircraft, and then,
b) determining the increase or decrease in the
total reversible energy contained in the aircraft
induced by changing wind conditions by detecting
any change in said balance.
2. Apparatus for detecting the magnitude of a
windshear condition encountered by an airborne
aircraft comprising means responsive to the thrust,
drag and weight of the aircraft for providing a first
component signal corresponding to the ratio of the
thrust minus drag to the weight, means responsive
to the longitudinal and normal acceleration of the
aircraft and a function of the aircraft angle of attack for providing a second component signal corres
ponding to the potential flight path angle of the aircraft, and means for combining said first and second component signals in such a manner as to provide a
resultant signal corresponding to any difference therebetween.
3. Apparatus according to claim 2, wherein the means for providing said first component signal comprises means responsive to an operation characteristic of the aircraft engines at an existing airspeed and pressure altitude for providing a signal proportional to the thrust imparted to the aircraft by the engine, means responsive to the existing weight of the aircraft for providing a signal proportional to the aircraft weight, means responsive to said normal acceleration and angle of attack responsive means for providing a component signal proportional to the drag of the aircraft, and means responsive to the thrust and drag signals and the weight signal for providing a signal proportional to the ratio thereof.
4. Apparatus according to claim 3 and further comprising means responsive to the position of the aircraft flaps for modifying the drag signal providing means in accordance with flap position.
5. Apparatus according to any of claims 2 to 4, wherein the second component signal providing means includes means responsive to the position of the aircraft flaps for modifying the flight path angle signal in accordance with flap position.
6. Apparatus according to any of claims 2 to 5, and further comprising means for providing a bias signal corresponding to a predetermined magnitude of windshear, and means responsive to the combining means and the bias signal for providing a warning signal when the output of the combining means exceeds the bias signal.
7. Apparatus according to claim 6 and further including means responsive to the aircraft airpseed for varying the value of the bias signal.
8. Apparatus according to any of claims 2 to 7, wherein the means responsive to the drag and flight path angle of the aircraft comprises means for storing predetermined values of the ratio of the coefficient of drag to the coefficient of lift minus the sine of the aircraft angle of attack as a function If the aircraft angle of attack, means responsive to the actual aircraft angle of attack for providing a measure of the actual value of said predetermined function, means responsive to the aircraft's normal acceleration and the actual predetermined function measure for providing a measure of the product thereof in the form of a first component signal proportional to the drag of the aircraft, and means responsive to said product measure and the aircraft longitudinal acceleration for providing a second component signal proportional to the flight path angle ofthe aircraft.
9. Apparatus according to claim 8, wherein the stored predetermined values are dependent upon the position of the aircraft flaps, and means responsive to the position of the flaps for providing the
measure of the actual value of said predetermined function corresponding to the actual position of the flaps.
10. Apparatus for detecting the existence of a windshear condition encountered by an airborne aircraft comprising accelerometer means for providing a first signal proportional to the acceleration of the aircraft along its longitudinal axis and a second signal proportional to the acceleration of the aircraft along its vertical axis, means for providing a third signal proportional to the angle of attack of the aircraft, computation means responsive to said first, second and third signals for providing a component measure of the potential flight path angle of the aircraft and responsive to said second and third signals for providing a component measure of the drag characteristic of the aircraft, means for providing a component measure of the thrust-to-weight ratio of the aircraft, and combining means responsive to the algebraic sum of all of said component measures for providing an output signal corresponding to the magnitude of the windshear encountered by the aircraft.
11. Apparatus for detecting the magnitude of a windshear condition encountered by an aircraft substantially as herein particularly described with reference to the accompanying drawings.
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US23928981A | 1981-03-02 | 1981-03-02 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| GB2094006A true GB2094006A (en) | 1982-09-08 |
Family
ID=22901498
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB8205793A Withdrawn GB2094006A (en) | 1981-03-02 | 1982-02-26 | Windshear detection and warning system |
Country Status (5)
| Country | Link |
|---|---|
| JP (1) | JPS57191198A (en) |
| DE (1) | DE3207478A1 (en) |
| FR (1) | FR2509469A1 (en) |
| GB (1) | GB2094006A (en) |
| IT (1) | IT1156246B (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0120855A1 (en) * | 1982-09-30 | 1984-10-10 | Boeing Co | Total energy based flight control system. |
| EP0125087A2 (en) * | 1983-05-06 | 1984-11-14 | Honeywell Inc. | Windshear detection and warning system |
| EP0235962A2 (en) * | 1986-02-28 | 1987-09-09 | Honeywell Inc. | Windshear detection and warning system |
| EP0235963A2 (en) * | 1986-02-28 | 1987-09-09 | Honeywell Inc. | Vertical windshear detection for aircraft |
| CN104318067A (en) * | 2014-09-29 | 2015-01-28 | 中国商用飞机有限责任公司 | Wind shear detection method and device based on energy management |
| CN116612669A (en) * | 2023-05-05 | 2023-08-18 | 江苏省气象信息中心(江苏省气象档案馆) | Intelligent aviation real-time meteorological data analysis and early warning method and equipment |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2817748B2 (en) * | 1992-01-08 | 1998-10-30 | ヤマハ株式会社 | Electronic musical instrument |
| FR2891802B1 (en) * | 2005-10-11 | 2009-03-13 | Airbus France Sas | METHOD AND DEVICE FOR ATTENUATING THE EFFECTS OF VERTICAL TURBULENCE ON AN AIRCRAFT |
| JP7330293B2 (en) * | 2020-06-29 | 2023-08-21 | 楽天グループ株式会社 | CONTROL DEVICE, INFORMATION SYSTEM AND METHOD |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3691356A (en) * | 1970-12-10 | 1972-09-12 | Sperry Rand Corp | Speed command and throttle control system for aircraft |
| FR2334110A1 (en) * | 1975-12-03 | 1977-07-01 | Equip Navigation Aerienne | FAST WIND GRADIENT DETECTION METHOD AND SYSTEM |
-
1982
- 1982-02-23 JP JP57028035A patent/JPS57191198A/en active Pending
- 1982-02-26 GB GB8205793A patent/GB2094006A/en not_active Withdrawn
- 1982-03-01 IT IT47885/82A patent/IT1156246B/en active
- 1982-03-01 FR FR8203355A patent/FR2509469A1/en not_active Withdrawn
- 1982-03-02 DE DE19823207478 patent/DE3207478A1/en not_active Withdrawn
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0120855A1 (en) * | 1982-09-30 | 1984-10-10 | Boeing Co | Total energy based flight control system. |
| EP0120855A4 (en) * | 1982-09-30 | 1985-07-01 | Boeing Co | Total energy based flight control system. |
| EP0125087A2 (en) * | 1983-05-06 | 1984-11-14 | Honeywell Inc. | Windshear detection and warning system |
| EP0125087A3 (en) * | 1983-05-06 | 1985-05-08 | Sperry Corporation | Windshear detection and warning system |
| EP0235962A2 (en) * | 1986-02-28 | 1987-09-09 | Honeywell Inc. | Windshear detection and warning system |
| EP0235963A2 (en) * | 1986-02-28 | 1987-09-09 | Honeywell Inc. | Vertical windshear detection for aircraft |
| EP0235962A3 (en) * | 1986-02-28 | 1988-09-07 | Honeywell Inc. | Windshear detection and warning system |
| EP0235963A3 (en) * | 1986-02-28 | 1988-09-07 | Honeywell Inc. | Vertical windshear detection for aircraft |
| CN104318067A (en) * | 2014-09-29 | 2015-01-28 | 中国商用飞机有限责任公司 | Wind shear detection method and device based on energy management |
| CN116612669A (en) * | 2023-05-05 | 2023-08-18 | 江苏省气象信息中心(江苏省气象档案馆) | Intelligent aviation real-time meteorological data analysis and early warning method and equipment |
Also Published As
| Publication number | Publication date |
|---|---|
| DE3207478A1 (en) | 1982-09-16 |
| IT1156246B (en) | 1987-01-28 |
| IT8247885A0 (en) | 1982-03-01 |
| JPS57191198A (en) | 1982-11-24 |
| FR2509469A1 (en) | 1983-01-14 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| WAP | Application withdrawn, taken to be withdrawn or refused ** after publication under section 16(1) |