DYNAMIC LIMITATION OF MONOBLOCK FLIGHT CONTROL
SURFACES INCLINATIONS DURING STALL SUSCEPTIBILITY
CONDITIONS
FIELD OF THE DISCLOSED TECHNIQUE
The disclosed technique generally relates to V-tail aircrafts with
automatically controlled monoblock flight control surfaces.
BACKGROUND OF THE DISCLOSED TECHNIQUE
Conventional aircrafts are usually designed in a T-tail
configuration, in which there are three tail stabilizing surfaces at the rear
of the aircraft, with two horizontally oriented stabilizers mounted on either
side of a vertically oriented stabilizer, resembling the shape of the letter
"T" when viewed from the front or rear. An alternative configuration is the
"V-tail", also known as a "butterfly tail", where the three tail stabilizers (two
horizontal and one vertical) are replaced with two slanted stabilizers,
resembling the shape of the letter "V" when viewed from the front or rear.
The movable flight control surfaces differ between these two types of
aircrafts. Whereas a T-tail aircraft includes "rudders" and "elevators", for
separately controlling the yaw and pitch motions, respectively, a V-tail
aircraft includes "ruddervators", which control the yaw and pitch motions
jointly.
In a T-tail aircraft, the rudders are mounted on the trailing edges
on either side of the vertical stabilizer (or "fin"), and the elevators are
mounted on the trailing edges of each of the two horizontal stabilizers (or
"tailplanes"). In a V-tail aircraft, there are two ruddervators mounted on
the trailing edge of the left and right tail stabilizers, respectively. A T-tail
aircraft pitches down by tilting both elevators downwards, resulting in
lower pressure above each tailplane and higher pressure below, causing
the tailplanes to lift and the aircraft to nose-down. Correspondingly, when
both elevators are raised, the pressure is reduced below the tailplanes
and raised above them, causing the aircraft to tail-down and nose-up. A
V-tail aircraft pitches down by tilting the left ruddervator downward and to
the left and tilting the right ruddervator downward and to the right,
producing an overall tail lifting force while the resultant left and right yaw
forces cancel each other out. Correspondingly, a V-tail aircraft pitches up
by raising the left ruddervator upward and to the right and tilting the right
ruddervator upward and to the left, producing an overall downward force
on the tail stabilizers while the resultant left and right yaw forces cancel
each other out.
A T-tail aircraft yaws to the right by tilting both rudders to the
right, resulting in lower pressure on the left side of the fin and higher
pressure to the right, causing the tail to move left and the aircraft to noseright.
Correspondingly, when both rudders are tilted to the left, the
pressure is reduced on the right side of the fin and raised on the left side,
causing the tail to move right and the aircraft to nose-left. A V-tail aircraft
yaws to the right by tilting the left ruddervator upward and to the right
while tilting the right ruddervator downward and to the right, resulting in an
overall tail-rightward force (causing the aircraft to nose-right) while the
resultant up and down pitch forces cancel each other out.
Correspondingly, a V-tail aircraft yaws to the left by tilting the left
ruddervator downward and to the left while tilting the right ruddervator
upward and to the left, resulting in an overall tail-leftward force (causing
the aircraft to nose-left) while the resultant up and down pitch forces
cancel each other out.
In general, a V-tail aircraft has less weight and produces less
drag relative to a T-tail aircraft, but requires a more complex control
system to handle the flight control surfaces and also suffers reduced
directional dynamic stability.
In some aircrafts, the flight control surfaces are integrally formed
together with the respective tail stabilizer surfaces, rather than being
formed as a separate movable trailing edge. Such a design is also
referred to as a "monoblock" configuration.
Aircrafts generally have multiple control surfaces, each of which
may incline or tilt about a different rotational axis, for controlling different
types of aircraft motion. Reference is now made to Figures 1A, 1B, and
1C. Figure 1A is a rear view schematic illustration of a V-tail aircraft 10
ruddervator, referenced 14, in a centered position about a first rotational
axis, referenced 18. Figure 1B is a rear view schematic illustration of the
ruddervator 14 of Figure 1A rotated in a clockwise direction. Figure 1C is
a rear view schematic illustration of the ruddervator 14 of Figure 1A
rotated in a counterclockwise direction.
Reference is now made to Figures 2A, 2B and 2C. Figure 2A is
a top view schematic illustration of a V-tail aircraft, referenced 20, with
ruddervators, referenced 22 and 24, in a centered position about a second
rotational axis, referenced 26. Figure 2B is a top view schematic
illustration of the V-tail aircraft 20 of Figure 2A with ruddervators 22, 24
rotated in a first direction. In particular, both ruddervators 22, 24 are tilted
toward the rear of aircraft 20 (i.e., when viewed from the top of aircraft 20,
right ruddervator 22 is tilted clockwise and left ruddervator 24 is tilted
counterclockwise). Figure 2C is a top view schematic illustration of the
V-tail aircraft 10 of Figure 2A with ruddervators 22, 24 rotated in a second
direction. In particular, both ruddervators 22, 24 are tilted toward the front
of aircraft 20 (i.e., when viewed from the top of aircraft 20, right
ruddervator 22 is tilted counterclockwise and left ruddervator 24 is tilted
clockwise).
Reference is now made to Figures 3A, 3B and 3C. Figure 3A is
a rear view schematic illustration of a V-tail aircraft, referenced 30, with
ruddervators, referenced 32 and 34, in a centered position about a third
rotational axis, referenced 36. Figure 3B is a rear view schematic
illustration of the V-tail aircraft 30 of Figure 3A with ruddervators 32, 34
rotated in a first direction. In particular, both ruddervators 32, 34 are tilted
upwards (i.e., when viewed from the rear of aircraft 30, left ruddervator 32
is tilted clockwise and right ruddervator 34 is tilted counterclockwise).
Figure 3C is a rear view schematic illustration of the V-tail aircraft 30 of
Figure 3A with ruddervators 32, 34 rotated in a second direction. In
particular, both ruddervators 32, 34 are tilted downwards (i.e., when
viewed from the rear of aircraft 30, left ruddervator 32 is tilted
counterclockwise and right ruddervator 34 is tilted clockwise).
The "angle of attack (AOA)" of an aircraft refers to the acute
angle between the chord of the airfoil (i.e., aircraft wing) and the direction
of undisturbed relative airflow, which is essentially the angle between the
direction of the aircraft wing and the direction it is travelling. The "angle of
sideslip (AOS)" refers to the angle between the aircraft centerline and the
relative wind, which can be considered the directional AOA of the aircraft.
An aircraft will experience stall if the aircraft exceeds a value known as the
"critical angle of attack", resulting in a rapid decrease in lift caused by a
separation of airflow from the wing surface. In a stall, the wing cannot
generate adequate lift to sustain level flight. The lift coefficient generally
increases as a function of AOA up until a maximum point, after which it
decreases dramatically. This maximum lift coefficient point corresponds
to the critical AOA. A stall may occur at any pitch attitude or any airspeed,
but usually occurs when the airspeed is reduced below what is known as
the "unaccelerated stall speed".
Each fixed-wing aircraft has a specific unique critical AOA at
which a stall would occur. This value is usually static and predefined prior
to the flight, such that the pilot and aircraft control systems can avoid
reaching the critical AOA and thus avoid entering into a stall. The actual
value of the critical AOA is dependent on various parameters associated
with the design of the aircraft (e.g., wing profile, planform, wing aspect
ratio), but is typically in the range of 8°-20°. These parameters may be
influenced by the weather conditions. In particular, the temperature and
humidity in the flight environment may result in the formation of ice and
other forms of frozen precipitation on the surfaces of the wings, which in
turn would affect the predefined critical AOA value, usually to further limit
the critical AOA. Reference is now made to Figure 4, which is a graph,
generally referenced 50, showing the effect of accumulated ice on the lift
coefficient of a V-tail aircraft as a function of the angle of attack. The
y-axis of graph 50 represents the lift coefficient (CL) , while the x-axis of
graph 50 represents the angle of attack (a) in degrees. Graph 50 depicts
the lift coefficient as a function of the angle of attack for V-tails with
varying degrees of accumulated ice on their surfaces. Curve 52
represents a "clean V-tail", i.e., one with no accumulated ice, while curves
54, 56 and 58, respectively represent V-tails with accumulated ice at a
thickness of increasing 5% chord-wise intervals.
Some aircrafts are equipped with mechanisms for ice removal
from the wings, but these mechanisms are not always completely reliable
or totally effective, and may still leave a certain amount of ice.
Furthermore, the weather conditions tend to change in real-time during the
actual flight, and cannot be forecasted ahead of time with 100% reliability.
It is possible to completely refrain from implementing flights during
weather conditions that would result in ice accumulation on the wing
surfaces, or to modify the flight route to mitigate the effect of these
weather conditions, although these approaches are not always feasible or
practical. Safety considerations should be taken into account in defining
the particular critical AOA that will be utilized during the flight. In severe
weather conditions such as rain, snow and ice, the aircraft must reduce
loss of aerodynamic characteristics to a tolerable level and increase its
aerodynamic safety margin. Unmanned aerial vehicle (UAV) aircrafts are
particularly sensitive to icy weather conditions, as such aircrafts are
typically not equipped with mechanisms and resources for dealing with
such a scenario.
U.S. Patent No. 5,826,834 to Potter et al, entitled "Self adaptive
limiter for automatic control of approach and landing", is directed to a fail
passive flight control system for controlling the approach and landing of an
aircraft. The control system includes a pitch limiter in communication with
an autopilot. The limiter computes an estimated flight path angle based
on vertical speed data and horizontal speed data of the aircraft. The
limiter continuously computes a nominal flight path angle from the
estimated flight path angle during a tracking phase of the
approach/landing, until a predetermined altitude is reached and the
nominal flight path angle is latched. The limiter continuously computes a
nominal vertical speed based on the nominal flight path angle and
horizontal speed data, and further continuously computes a vertical speed
limit from the nominal vertical speed and altitude data. The limiter
computes a pitch limit value from the vertical speed limit, the vertical
speed, and aircraft pitch data. The autopilot limits the aircraft pitch to the
pitch limit value, thus preventing the aircraft from pitching down
excessively and descending below certification terrain clearance
requirements.
U.S. Patent No. 6,253,1 26 to Palmer, entitled "Method and
apparatus for flight parameter monitoring and control", is directed to the
monitoring of aircraft flight parameters, particularly air pressures acting on
various surfaces of the aircraft. According to one aspect, the skin of the
aircraft is provided with small openings or ports that are connected by an
air pressure conduit to pressure sensors. The ports are sensitive to air
pressure changes associated with flight at different speeds. The ports are
also provided with means to deter extraneous matter (e.g., water, vapor,
lubrication and deicing fluids, particulates), means to prevent icing of the
port, and means to decontaminate the port (e.g., a port heater and a sump
volume). The air pressures are measured, recorded and stored during a
first flight condition, and subsequently during a second flight condition.
The measurements are compared, and utilized for deducing aerodynamic
performance data (e.g., correct angle of attack and margin to stall) and
determining how to control the aircraft accordingly.
U.S. Patent Application Publication No. 2009/0062973 to
Caldeira et al, entitled "Stall, buffeting, low speed and high attitude
projection system", is directed to an aircraft flight control system for
providing further safety controls. The aircraft control surfaces may be
actuated to deploy to a certain position by a pilot interceptor (pilot input
device) command. The control system monitors a set of flight parameters
(e.g., angle-of-attack, angle-of-attack rate, airspeed, airspeed rate, flap
position, gear position, pitch attitude, pitch rate, height above ground, ice
detection) and processes the data to determine if the aircraft is operating
inside a permitted envelope. If the aircraft is close to the envelope limits,
the control system may bypass the pilot interceptor command to
automatically position the control surfaces. The control system may
protect the aircraft from scenarios such as low speeds, high attitude, stalls
and buffetings.
Abzug, "V-Tail Stalling at Combined Angles of Attack and
Sideslip Information", J. Aircraft, Vol.36, No.4 : Engineering Notes, 1999,
pp.729-731 , discloses the calculation of the V-tail panel geometric angle
of attack (AOA) and sideslip angle (AOS) as a function of six variables:
aircraft AOA (a), aircraft AOS () , V-tail average downwash angle () ,
V-tail average sidewash angle (o), V-tail dihedral angle () , and V-tail
incidence angle () for an all-moving V-tail. The calculations are valid for
large AOA and AOS values to support studies of possible panel stalling.
In a sample calculation of a landing approach for a 30 ° dihedral V-tail, the
left panel would reach a stall point at an AOA of - 12° which is obtained at
a right sideslip angle of 17°. The critical AOS for panel stall was found to
be reduced by 3 degrees when the assumed sidewash angle is increased
from 20% to 50% of the AOS. The critical AOS for panel stall was found
to be reduced by 5 degrees when the downwash factor (0) is increased
from 4 to 8. Induction from the opposite panel was found to reduce the
local panel AOA of a V-tail in sideslip below those for the same AOA (i.e.,
raising the panel AOA at which a stall would occur), relative to the same
V-tail without sideslip. Conversely, panel crossflow on a V-tail in sideslip
lowers the panel AOA at which a stall would occur, relative to the same
V-tail without sideslip.
SUMMARY OF THE DISCLOSED TECHNIQUE
In accordance with one aspect of the disclosed technique, there
is thus provided an aircraft that includes monoblock flight control surfaces
(FCS) and a controller for dynamically limiting the inclinations of the flight
control surfaces during a stall susceptibility condition. The controller
obtains the real-time calibrated airspeed of the aircraft, obtains the real
time angle of attack (AOA) of the aircraft, obtains the real-time sideslip
angle (AOS) of the aircraft, and calculates the real-time local AOA and
AOS of the flight control surfaces, based on the obtained aircraft
parameters. The controller dynamically limits the inclination of each of the
flight control surfaces relative to the critical values according to the
calculated real-time local AOA and AOS of each of the flight control
surfaces. The aircraft may be an unmanned aerial vehicle (UAV). The
aircraft may be a V-tail aircraft. The flight control surfaces may include at
least a left tail stabilizer and a right tail stabilizer, which are independently
dynamically limited. The aircraft may further include temperature and
precipitation sensors for detecting current weather conditions in the
aircraft environment. The stall susceptibility condition may include icy
conditions.
In accordance with another aspect of the disclosed technique,
there is thus provided a method for dynamically limiting the inclinations of
the monoblock flight control surfaces (FCS) of an aircraft. The method
includes the procedure of activating dynamic limitation of the flight control
surfaces if a stall susceptibility condition is detected in the current
environment of the aircraft. The method further includes the procedures
of obtaining the real-time calibrated airspeed of the aircraft, obtaining the
real-time angle of attack (AOA) of the aircraft, obtaining the real-time
sideslip angle (AOS) of the aircraft, and calculating the real-time local
AOA and AOS of the flight control surfaces, based on the obtained aircraft
parameters. The method further includes the procedure of dynamically
limiting the inclination of each of the flight control surfaces relative to the
critical values according to the calculated real-time local AOA and AOS of
each of the flight control surfaces. The real-time calibrated airspeed of the
aircraft may be measured using an aircraft measurement apparatus.
Alternatively, the real-time calibrated airspeed of the aircraft may be
estimated if the measured airspeed data is deemed unsuitable. The
estimation may be based on the density, engine RPM, measured throttle,
measured pitch angle, and measured x-axis acceleration of the aircraft,
using an open loop state-space model. The real-time AOA of the aircraft
may be estimated based on the measured z-axis acceleration and
calibrated airspeed of the aircraft. The real-time AOS of the aircraft may
be estimated based on the measured y-axis acceleration, true airspeed,
calibrated airspeed, yaw rate, and rudder angle of the aircraft. The
real-time local AOA and AOS of the flight control surfaces may be
calculated by estimating the local AOA and AOS based on previously
calculated average FCS downwash angle and FCS sidewash angle,
known FCS dihedral angle, and airspeed velocity components in the wind
coordinate system axes. The method may further include the procedure
of detecting current weather conditions in the aircraft environment using
temperature and precipitation sensors. The aircraft may be an unmanned
aerial vehicle (UAV). The aircraft may be a V-tail aircraft. The stall
susceptibility condition may include icy conditions.
BRIEF DESCRIPTION OF THE DRAWINGS
The disclosed technique will be understood and appreciated
more fully from the following detailed description taken in conjunction with
the drawings in which:
Figure 1A is a rear view schematic illustration of a V-tail aircraft
ruddervator in a centered position about a first rotational axis;
Figure 1B is a rear view schematic illustration of the ruddervator
of Figure 1A rotated in a clockwise direction;
Figure 1C is a rear view schematic illustration of the ruddervator
of Figure 1A rotated in a counterclockwise direction;
Figure 2A is a top view schematic illustration of a V-tail aircraft
with ruddervators in a centered position about a second rotational axis;
Figure 2B is a top view schematic illustration of the V-tail aircraft
of Figure 2A with ruddervators rotated in a first direction;
Figure 2C is a top view schematic illustration of the V-tail aircraft
of Figure 2A with ruddervators rotated in a second direction;
Figure 3A is a rear view schematic illustration of a V-tail aircraft
with ruddervators in a centered position about a third rotational axis;
Figure 3B is a rear view schematic illustration of the V-tail
aircraft of Figure 3A with ruddervators rotated in a first direction;
Figure 3C is a rear view schematic illustration of the V-tail
aircraft of Figure 3A with ruddervators rotated in a second direction;
Figure 4 is a graph showing the effect of accumulated ice on the
lift coefficient of a V-tail aircraft as a function of the angle of attack;
Figure 5 is a block diagram of an unmanned aerial vehicle
(UAV) with a V-tail configuration, constructed and operative in accordance
with an embodiment of the disclosed technique;
Figure 6 is a block diagram of a method for dynamically limiting
the inclinations of monoblock flight control surfaces of an aircraft,
operative in accordance with an embodiment of the disclosed technique;
Figure 7 is a schematic illustration of a complementary filter
architecture used for calculating an estimation of the aircraft sideslip angle
(AOS), operative in accordance with an embodiment of the disclosed
technique;
Figure 8 is a schematic illustration of a process architecture for
estimating the aircraft sideslip angle (AOS), operative in accordance with
an embodiment of the disclosed technique; and
Figure 9 is a plurality of graphs showing the result of a
simulation of the dynamic V-tail angle limitation of the disclosed technique
in conjunction with a Hermes® 450 UAV.
DETAILED DESCRIPTION OF THE EMBODIMENTS
The disclosed technique overcomes the disadvantages of the
prior art by providing a method for dynamically limiting the inclinations of
the monoblock flight control surfaces of an aircraft during flight in a stall
susceptibility scenario, such as icy weather conditions. The disclosed
technique improves upon the usage of a static limitation on the control
surfaces inclination for preventing stalling, by providing a real-time
dynamic limitation for the degree of inclination of each of the control
surfaces according to aerodynamic constraints. The real-time tracking
and subsequent dynamic limitation applied to each individual flight control
surface ultimately provides the aircraft with enhanced maneuvering ability.
The method includes detecting current weather conditions and activating a
dynamic limitation of the aircraft flight control surfaces if a stall
susceptibility condition is detected. The real-time calibrated airspeed of
the aircraft is measured or estimated. The real-time angle of attack (AOA)
of the aircraft is measured or estimated. The real-time sideslip angle
(AOS) of the aircraft is measured or estimated. The real-time local AOA
and AOS of the control surfaces are calculated from aircraft parameters.
The inclination of the control surfaces are dynamically limited based on
the calculated local AOA and AOS, where the limitation is applied
independently to each control surface (e.g., a left tail stabilizer and a right
tail stabilizer) in accordance with the parameters associated with that
specific control surface. The disclosed technique is particularly applicable
to unmanned aerial vehicle (UAV), specifically aircrafts implementing
automatic flight control, but is generally applicable to other types of
aircrafts as well. Additionally, the disclosed technique is particularly
applicable to aircrafts with a V-tail configuration, but is generally
applicable to other types of aircrafts (e.g., T-tail aircrafts) as well.
The term "monoblock" control surfaces, and any variations
thereof, as used herein refers to flight control surfaces that are integrally
formed together with the respective wing or tail stabilizer surface, i.e.,
rather than being formed as a separate movable trailing edge. Namely,
the wing or tail in its entirety is also formed and operates as a unitary flight
control surface. For example, in a V-tail aircraft with monoblock control
surfaces, the left tail stabilizer and left ruddervator is integrated into a
single movable surface (rotatable about a first control axis), while the right
tail stabilizer and right ruddervator is similarly integrated into a single
movable surface (rotatable about a second control axis), for controlling the
aircraft pitch and yaw motions.
The terms "tilting" and "inclining", and any variations thereof, are
used herein interchangeably to refer to the operation of adjusting the
alignment of the entire control surface plane in relation to a given
reference plane or axis, or alternatively, rotating the entire control surface
plane about a given reference plane/axis, i.e., such that the plane of the
control surface defines a particular inclination angle with respect to the
reference plane/axis. Such an operation does not result in any
manipulation of the form or shape of the control surface itself (e.g., a
twisting or deformation thereof).
The term "icy conditions", and any variations thereof, as used
herein refers to weather conditions that result in the accumulation of
frozen precipitation on the aircraft wings and other aircraft surfaces, which
influences the real-time critical angle of attack (AOA) of the aircraft
(beyond which the aircraft would enter into a stall).
The term "stall susceptibility condition", and any variations
thereof, as used herein refers to any situation that causes the aircraft
critical AOA to change and/or any situation that increases the likelihood of
the aircraft entering into a stall. One type of stall susceptibility condition is
icy conditions (as defined above).
The following abbreviations will be used hereinbelow:
UAV = Unmanned Aerial Vehicle
A/V = Aerial Vehicle
= A/V Y-axis acceleration (body coordinates)
= A/V Z-axis acceleration (body coordinates)
b = A/V Span
L = Lift
m = Mass
p =A/V roll rate
r = A/V yaw rate
S = A/V reference area
= Lift coefficient
= Lift coefficient at zero angle of attack
= Lift coefficient slope
= Y-axis total force coefficient (body coordinates)
= Y-axis force coefficient at trim
= Y-axis force coefficient due to sideslip angle
= Y-axis force coefficient due to aileron angle
= Y-axis force coefficient due to rudder angle
= Y-axis force coefficient due to roll rate
= Y-axis force coefficient due to yaw rate
= Calibrated airspeed
= True airspeed
a = AOA = Angle of attack
= AOS = Sideslip angle
= Air density at sea level
= V-tail average downwash angle
= A/V track angle (angle between north and ground velocity)
oa = V-tail average sidewash angle
= V-tail control angle
S
= aileron angle
= rudder angle
= V-tail dihedral angle
Reference is now made to Figure 5, which is a block diagram of
an aircraft, generally referenced 100, constructed and operative in
accordance with an embodiment of the disclosed technique. Aircraft 100
includes a controller 102, temperature and precipitation sensors 104, flight
parameters measurement apparatuses (FPMA) 106, flight control surfaces
(FCS) actuators 108, and flight control surfaces 110. FCS 110 includes a
left tail stabilizer 112 and a right tail stabilizer 114. FPMA 106 includes at
least one pitot tube 116. Controller 102 is coupled with temperature and
precipitation sensors 104, with FPMA 106, and with FCS actuators 108.
FCS actuators 108 are further coupled with left tail stabilizer 112 and with
right tail stabilizer 114.
Aircraft 100 is preferably a UAV, for example a Hermes® series
type UAV (such as Hermes®450 or Hermes® 900), which is a UAV that is
classified as a class 1 (small, light) according to Rockwell RPV Flying
Qualities Design Criteria. The longitudinal and lateral stabilization and
control of a Hermes® 450 UAV is performed using two monoblock V-tail
control surfaces. Accordingly, aircraft 100 is preferably a V-tail aircraft,
with monoblock flight control surfaces.
Reference is now made to Figure 6, which is a block diagram of
a method for dynamically limiting the inclinations of monoblock flight
control surfaces of an aircraft, operative in accordance with an
embodiment of the disclosed technique. In procedure 152, current
weather conditions are detected using temperature and precipitation
sensors. Referring to Figure 5, temperature and precipitation sensors 104
detect various weather parameters (e.g., temperature, precipitation level),
to provide an indication of the current weather conditions in the
environment in which aircraft 100 is currently situated. Controller 102 may
optionally determine an updated real-time critical AOA for aircraft 100 that
is appropriate for the detected real-time weather conditions.
In procedure 154, dynamic limitation of the aircraft flight control
surfaces is activated if a stall susceptibility condition is detected.
Referring to Figure 5, if the current weather parameters detected by
temperature and precipitation sensors 104 provide an indication of icy
conditions in the environment (e.g., if the detected parameters exceed
some predefined threshold levels or meet some predefined criteria), than
aircraft 100 activates the implementation of dynamic limitation of flight
control surfaces 110 in accordance with the real-time conditions. It is
appreciated that the dynamic limitation may generally be activated upon
detection of other types of stall susceptibility conditions as well.
In procedure 156, the real-time calibrated airspeed of the aircraft
is obtained. Procedure 156 may be implemented via procedure 158, in
which the real-time calibrated airspeed of the aircraft is measured using
an airspeed measurement apparatus, or alternatively via procedure 160,
in which the real-time calibrated airspeed of the aircraft is estimated from
other flight parameters if the measured airspeed value is disqualified.
Referring to Figure 5, pitot tube 116 acquires measurements of the real
time calibrated airspeed of aircraft 100. A pitot tube is a pressure
measurement instrument that measures fluid flow velocity, and is
commonly used to determine the airspeed of an aircraft. Controller 102
receives the acquired calibrated airspeed measurements. If the
measurements acquired by pitot tube 116 is deemed unsuitable or
insufficiently reliable, then these measurements are disqualified and an
updated real-time calibrated airspeed value is estimated. In certain
situations, particularly during icy conditions, a pitot tube does not provide
reliable measurements. For example, if aircraft 100 includes multiple pitot
tubes 116 and there are significant discrepancies between the airspeed
measurements provided by each of the pitot tubes 116, then these
measurements are deemed unsuitable and an estimation process is
carried out instead. Controller 102 estimates the calibrated airspeed
using an open loop state-space observer with a heuristic correction
component (based on flight test data). The open loop state-space model
utilizes the following flight parameters as input: density, engine RPM,
measured throttle, measured pitch angle, and measured X-axis
acceleration. These flight parameters are measured by FPMA 106 and
transferred to controller 102. The density is calculated using a fully
redundant static pressure sensor (e.g., included in FPMA 106), which was
tested and shown to perform with high reliability under icy conditions.
In procedure 162, the real-time angle of attack (AOA) of the
aircraft is obtained. Procedure 162 may be implemented via procedure
164, in which the aircraft AOA is measured using an AOA measurement
apparatus, or alternatively via procedure 166, in which the aircraft AOA is
estimated from other flight parameters if the measured AOA value is
disqualified. Referring to Figure 5, an AOA sensor or other suitable
sensor of FPMA 106 acquires measurements of the real-time AOA of
aircraft 100. Pitot tube 116 may be utilized to obtain the AOA
measurements. If the measurements acquired by FPMA 106 is deemed
unsuitable or insufficiently reliable (which is likely to occur during icy
conditions), then these measurements are disqualified and an updated
real-time AOA value is estimated. In particular, controller 102 directly
calculates the AOA of aircraft 100 using the following equations:
i = . -S =
2} .< K C = , ¾ . • .
1 — "¾ ¾
Equations ( 1) and (2) were derived from the aerodynamic
database of a Hermes® 450 UAV extracted from wind tunnel tests and
validated using flight tests. If for example aircraft 100 is a Hermes® 450
type UAV, the AOA can be estimated using the following equation:
a = - ,9 4 1 -¾ -
Assuming a precise aerodynamic database, the only source of
error in this estimation arises from delays and errors in the Z-axis
acceleration and the calibrated airspeed measurements/estimates.
In procedure 168, the real-time sideslip angle (AOS) of the
aircraft is obtained. Procedure 168 may be implemented via procedure
170, in which the aircraft AOS is measured using an AOS measurement
apparatus, or alternatively via procedure 172, in which the aircraft AOS is
estimated from other flight parameters if the measured AOS value is
disqualified. Referring to Figure 5, an AOS sensor or other suitable
sensor of FPMA 106 acquires measurements of the real-time AOS of
aircraft 100. Pitot tube 116 may be utilized to obtain the AOS
measurements. If the measurements acquired by FPMA 106 is deemed
unsuitable or insufficiently reliable (which is likely to occur during icy
conditions), then these measurements are disqualified and an updated
real-time AOS value is estimated. In particular, controller 102 calculates
the AOS of aircraft 100 indirectly, based on a calculated approximation of
the AOS derivative and a simplified approximation of the AOS itself.
Direct calculation of the AOS of the aircraft is problematic since certain
stability derivatives , c S ) are functions of the AOS, as is evident from
the following force equation along the aircraft y-axis:
3) = - ¾- - p + - a G¾ S ÷
' .
Reference is now made to Figure 7, which is a schematic
illustration of a complementary filter architecture, generally referenced
180, used for calculating an estimation of the aircraft sideslip angle (AOS),
operative in accordance with an embodiment of the disclosed technique.
The output of complementary filter architecture 180 is the summation of a
low frequency response component and a high frequency response
component. The low frequency response component is a simplified
approximation of the AOS, which may be very accurate initially but which
deviates over time. The high frequency response component corresponds
to the integration over an accurate approximation of the AOS derivative,
which would be accurate only as time progresses. A correct summation of
these two components (i.e., by a correct selection of ) provides an
adequate estimation of the AOS value.
A simplified but precise approximation of the AOS derivative is
calculated using the following equation (assuming « J )
8 = — p . - « os s ) = —— —r
TAS
To obtain the simplified approximation of the AOS, equation (3)
above is simplified into the following crude approximation (assuming
where >CYSr and CY are calculated by recording several UAV
flights in various configurations (e.g., weight and center of gravity) with a
pitot tube that measures AOS, and then using optimization techniques to
derive a set of discrete values for C which minimizes the
overall difference between the measured AOS and the estimated AOS. A
filtered FCS issued rudder command may be used to eliminate the
estimation consolidation upon the measured rudder angle (Br).
Equations (4) and (5) above are then applied within
complementary filter architecture 180 to provide an accurate AOS
estimation, as the low-pass filtered AOS approximation yields appropriate
AOS values with a steady state error that is nullified by the high-pass
filtered integration of the AOS derivative. Reference is now made to
Figure 8, which is a schematic illustration of a process architecture,
generally referenced 190, for estimating the aircraft sideslip angle (AOS),
operative in accordance with an embodiment of the disclosed technique.
Referring back to Figure 6, in procedure 174, the real-time local
angle of attack (AOA) and sideslip angle (AOS) of the flight control
surfaces are calculated, based on the obtained aircraft parameters.
Referring to Figure 5, controller 102 calculates the local AOA and the local
AOS of left tail stabilizer 112 and right tail stabilizer 114. An FCS
coordinate system is defined as follows: the x-axis is positive toward the
forward direction in the FCS chord plane, the z-axis is positive normal to
the FCS chord plane, and the y-axis is positive toward the right.
Subsequently, the following four successive rotations may be
implemented to transform from the wind coordinate system (at which the
aircraft AOA and AOS have been obtained) to the FCS coordinate system,
neglecting the effects of the A/V aerial velocity:
1) Rotation by +;
2) Rotation by -;
3) Rotation by ;
4) Rotation by ;
where:
= the average downwash angle at the respective tail stabilizer;
oa = the average sidewash angle at the respective tail stabilizer;
= the dihedral of the FCS plane (positive for left panel); and
= the FCS control angle (for a monoblock tail stabilizer).
In matrix form, these transformations can be represented as follow:
The FCS local AOA and AOS can now be derived using the
following relations:
using previously calculated average FCS downwash angle and
FCS sidewash angle values, using the known FCS dihedral angle, and
using airspeed velocity components in the wind coordinate system axes.
It is appreciated that the calculated local AOA and AOS values may be
asymmetrical for each side of the FCS, e.g., the values for left tail
stabilizer 112 may be different than those for right tail stabilizer 114.
In procedure 176, the inclination of each control surface is
dynamically limited according to the calculated real-time local control
surface AOA and AOS. Referring to Figure 5, controller 102 sends
signals to FCS actuators 108 to limit the degree of tilting left tail stabilizer
112 and right tail stabilizer 114, such that the calculated real-time local
AOA and AOS values of these control surfaces are sufficiently far from the
critical AOA and AOS values (i.e., in accordance with aerodynamic safety
constraints). It is appreciated that these dynamic limitations are applied
asymmetrically to each relevant FCS in accordance with the parameters
associated with that specific control surface (e.g., the limitation applied to
left tail stabilizer 112 may different than the limitation applied to right tail
stabilizer 114).
The disclosed technique was tested in a simulation environment
(flight tests proven) with an aerodynamic database based upon several ice
wind tunnel runs. Reference is now made to Figure 9, which is a plurality
of graphs showing the result of a simulation of the dynamic V-tail angle
limitation of the disclosed technique in conjunction with a Hermes® 450
UAV. A simulation of a 1 second singlet in the pitch channel while the
Hermes® 450 performs a coordinated turn with 10 roll angle is shown in
Figure 9 (height = 8Kft, calibrated airspeed = 60kn; in the presence of
medium turbulence, according to the Dryden turbulence model). As can
be seen in Figure 9, the dynamic V-tail angle limitation of the disclosed
technique limits the V-tail angle in such a way that the local AOA remains
far from its stall limit (stall AOA = 10.5 , maximum AOA without limiter =
8.9 at the left V-tail, maximum AOA with limiter = 6.5 at the right V-tail).
Moreover the UAV dynamic responses (as can be seen from the
Hermes® 450 AOA and AOS) are clearly more satisfactory with the
dynamic V-tail angle limitation of the disclosed technique.
It will be appreciated by persons skilled in the art that the
disclosed technique is not limited to what has been particularly shown and
described hereinabove.
CLAIMS
1. An aircraft comprising monoblock flight control surfaces and a
controller for dynamically limiting the inclinations of said flight control
surfaces in said aircraft during a stall susceptibility condition, said
controller operative to obtain the real-time calibrated airspeed of said
aircraft, to obtain the real-time angle of attack (AOA) of said aircraft,
to obtain the real-time sideslip angle (AOS) of said aircraft, to
calculate the real-time local AOA and AOS of said flight control
surfaces based on the obtained aircraft parameters, and to
dynamically limit the inclination of each of said flight control surfaces
relative to the critical values according to the calculated real-time
local AOA and AOS of each of said flight control surfaces.
2. The aircraft of claim 1, wherein said aircraft is an unmanned aerial
vehicle (UAV).
3. The aircraft of claim 1, wherein said aircraft is a V-tail aircraft.
4. The aircraft of claim 1, wherein said flight control surfaces comprises
at least a left tail stabilizer and a right tail stabilizer, which are
independently dynamically limited.
5. The aircraft of claim 1, further comprising temperature and
precipitation sensors coupled with said controller, said temperature
and precipitation sensors operative for detecting current weather
conditions in said aircraft environment.
6. The aircraft of claim 1, wherein said stall susceptibility condition
comprises icy conditions.
7. A method for dynamically limiting the inclinations of the monoblock
flight control surfaces of an aircraft, the method comprising the
procedures of:
activating dynamic limitation of said flight control surfaces if a
stall susceptibility condition is detected in the current environment of
said aircraft;
obtaining the real-time calibrated airspeed of said aircraft;
obtaining the real-time angle of attack (AOA) of said aircraft;
obtaining the real-time sideslip angle (AOS) of said aircraft;
calculating the real-time local AOA and AOS of said flight control
surfaces, based on the obtained aircraft parameters; and
dynamically limiting the inclination of each of said flight control
surfaces relative to the critical values according to the calculated
real-time local AOA and AOS of each of said flight control surfaces.
8. The method of claim 7, wherein said procedure of obtaining the
real-time calibrated airspeed of said aircraft comprises measuring the
real-time calibrated airspeed of said aircraft using an aircraft
measurement apparatus.
9. The method of claim 7, wherein said procedure of obtaining the
real-time calibrated airspeed of said aircraft comprises estimating the
real-time calibrated airspeed of said aircraft if measured airspeed
data is deemed unsuitable.
10. The method of claim 9, wherein said estimating is based on the
density, engine RPM, measured throttle, measured pitch angle, and
measured x-axis acceleration of said aircraft, using an open loop
state-space model.
11. The method of claim 7, wherein said procedure of obtaining the
real-time AOA of said aircraft comprises estimating said AOA based
on the measured z-axis acceleration and calibrated airspeed of said
aircraft.
12. The method of claim 7, wherein said procedure of obtaining the
real-time AOS of said aircraft comprises estimating said AOS based
on the measured y-axis acceleration, true airspeed, calibrated
airspeed, yaw rate, and rudder angle of said aircraft.
13. The method of claim 7, wherein said procedure of calculating the
real-time local AOA and AOS of said flight control surfaces comprises
estimating said local AOA and AOS based on previously calculated
average FCS downwash angle and FCS sidewash angle, known FCS
dihedral angle, and airspeed velocity components in the wind
coordinate system axes.
14. The method of claim 7, further comprising the procedure of detecting
current weather conditions in said aircraft environment using
temperature and precipitation sensors.
15. The method of claim 7, wherein said aircraft is an unmanned aerial
vehicle (UAV).
16. The method of claim 7, wherein said aircraft is a V-tail aircraft.
17. The method of claim 7, wherein said stall susceptibility condition
comprises icy conditions.