Method and apparatus for determining the airspeed of rotary wing aircraft

A method and apparatus for determining the airspeed of a rotary wing aircraft having a sustaining rotor with a plurality of radially extending rotor blades for rotation about a substantially vertical axis comprises the steps of measuring an airflow sideslip angle established during rotation of each rotor blade by the vector of a radial flow component due to the relative wind vector and the rotational velocity perpendicular to a blade feathering axis, producing a signal representative of the sideslip angle, comparing the signal in a processor containing information on a known relationship between the sideslip angle and airspeed, providing an output signal representative of the airspeed and wind direction to a display which may display both the airspeed and wind direction.

BACKGROUND TO THE INVENTION 
This invention relates to a method and apparatus for determining the 
airspeed of a rotary wing aircraft and the wind direction, and is 
particularly concerned with providing such information at low airspeed. 
A means for accurately determining the airspeed of a helicopter and wind 
direction relative thereto in low speed flight is highly desirable 
especially during certain phases of operation such as during rejected 
take-offs and when operating from the decks of ships and oil rig 
platforms. 
Such instrumentation with sufficient accuracy has thus far proved difficult 
to achieve. 
1. Description of the Prior Art 
Conventional fuselage mounted devices such as pitot and static pressure 
sensing devices are adversely affected especially at low airspeeds by the 
highly irregular flow close to the helicopter fuselage due to the 
interaction of many influences including the effects of main rotor wake. 
An alternative prior approach has been to relate the rotor control 
positions to the airspeed and direction of the helicopter. This has proved 
difficult because the centre of gravity position must be known, the 
position of control means such as a swash plate must be very accurately 
monitored and allowances must be made for control inputs and accelerations 
so that the steady airspeed can be deduced. 
2. Object of the Invention 
A main objective of this invention is to provide a method and apparatus for 
determining the airspeed of a rotary wing aircraft which overcomes or 
reduces the aforementioned problems. A further objective is to achieve the 
main objective using measurements of the airflow on the main sustaining 
rotor blades which can be related directly to airspeed and relative wind 
direction. 
SUMMARY OF THE INVENTION 
Accordingly, in one aspect, this invention provides a method of detecting 
the airspeed of a rotary wing aircraft having a sustaining rotor with a 
plurality of radially extending rotor blades for rotation about a 
substantially vertical axis, comprising the steps of measuring on at least 
one of the rotor blades of the sustaining rotor, an airflow sideslip angle 
established by rotation of the rotor, producing a signal representative of 
said sideslip angle, comparing said signal in processing means containing 
information on a known relationship between said sideslip angles and 
airspeed, and providing an output signal representative of the airspeed 
and wind direction to a display means. 
In another aspect the invention provides an airspeed indicating system for 
a rotary wing aircraft having a sustaining rotor comprising a plurality of 
radially extending rotor blades for rotation about a generally vertical 
axis comprising sensing means on at least one of the rotor blades for 
sensing an airflow sideslip angle established by interaction between a 
radial flow caused by a relative wind vector and a rotational velocity 
perpendicular to a blade feathering axis and generating a signal 
representative of the sideslip angle, transmission means for transmitting 
the signal to processing means on the aircraft fuselage, the processing 
means containing information on a known relationship between the measured 
sideslip angle and airspeed for providing an Output signal representative 
of the airspeed and relative wind direction to indicating means for 
displaying said airspeed and relative wind direction. 
The airflow sideslip angle .beta. may be established by the equation: 
EQU .beta.=arctan{Vcos(.PSI.+.theta.)/(xVT+Vsin(.PSI.+.theta.))} 
where: 
V is the velocity of the relative wind, 
.theta. is the relative direction of the wind (the wind azimuth), 
VT is the tip speed of the rotor blades, 
x is the radial location of the sideslip detectors on the blades, and 
.PSI. is the rotor blade azimuth angle.

DESCRIPTION OF THE PREFERRED EMBODIMENTS 
Referring now to FIG. 1 a helicopter to, illustrated schematically in plan, 
includes a fuselage 21 having a rearwardly extending tail boom 22 located 
symmetrically about a longitudinal centre line 26. The helicopter has a 
main sustaining rotor 23 comprising four rotor blades 23A, 23B, 23C and 
23D extending radially from a rotor head (not shown) for rotation about a 
generally vertical axis 24 in a generally horizontal plane in the 
direction of arrow 24A. An anti-torque rotor 25 is located at the end of 
the tail boom 22. 
The azimuthal position of the rotor blades in the rotational disc are 
usually identified from 0.degree. relative the longitudinal centreline at 
the rear of the tail boom 22 in the direction of rotation, and the 
0.degree., 90.degree., 180.degree.and 270.degree. azimuth positions are 
identified in FIG. 1. This identification arrangement will be retained 
throughout this specification. 
Normally, with the helicopter flying forward parallel to the longitudinal 
centreline 26, a rotor blade moving between 0.degree. and 180.degree. 
azimuth positions is termed the advancing blade and when moving between 
180.degree. and 360.degree., the retreating blade. In this specification 
however, these terms are applied in respect of rotation of the blades 
commencing at positions perpendicular to a relative wind vector (as 
hereinafter described). In contrast, wind azimuth angles are identified as 
0.degree. approaching the nose of the helicopter and run clockwise from 
that position. 
The method and apparatus of this invention is based on the realisation by 
the inventor that, as a helicopter translates in any direction relative to 
the wind, a cyclic airflow velocity component is generated along the 
radial dimension, i.e. the span, of the main rotor blades. 
The velocity vector diagrams concerned with this invention are illustrated 
in FIG. 1. 
The vector diagram at the top of the Figure is representative of a 
helicopter velocity 27 illustrated as a forward velocity along the 
longitudinal centreline 26 and a natural wind direction 28 which is 
resolved into a relative wind vector 29 as actually felt by the 
helicopter. The relative wind vector 29 extends at an angle .alpha. from 
the 0.degree. wind azimuth position and, for the purposes of description 
only, is 75.degree. as shown in FIG. 1. 
The vector diagrams shown in connection with rotor blades 23A and 23C are 
located at a predetermined radial location (as hereinafter discussed) and 
comprise a rotational velocity vector 30 located perpendicular to a blade 
feathering axis, and the relative wind vector 29. These vectors resolve 
into a vector 31 defining with vector 30, a sideslip angle .beta.. It will 
be understood that .beta. is at a maximum in respect of diametrically 
opposed blades 23A and 23C which are parallel to the relative wind vector 
29 (at blade azimuth positions as hereinbefore defined of 105.degree. and 
285.degree.), and that .beta. is zero (i.e. there is no sideslip angle) on 
diametrically opposed blades 23B and 23D where the rotational velocity 30 
and relative wind velocity 29 are parallel (at blade azimuth positions of 
15.degree. and 195.degree. respectively). 
Thus it will be clear that the actual azimuth position of maximum and zero 
sideslip angles .beta. will depend on the direction of the relative wind 
vector 29, and that they will always be approximately 90.degree. apart in 
the rotor rotational disc, as determined by equation 1. 
Accordingly each of the rotor blades 23A, 23B, 23C and 23D experiences a 
variation in sideslip angle .beta. as the blades rotate around the rotor 
azimuth, and the inventor has found that the airspeed and the relative 
wind vector can be obtained from measurements of the airflow sideslip 
angle .beta.. 
The relationship between the sideslip angle .beta. and the relative wind 
velocity is given by the equation: 
EQU .beta.=arctan{Vcos(.PSI.+.theta.)/(xVT+Vsin(.PSI.+.theta.))}(1) 
where: 
V is the velocity of the relative wind, 
.theta.is the relative direction of wind (the wind azimuth), 
VT is the tip speed of the rotor blades, 
x is the radial location of sideslip detectors on the blades, and 
.PSI. is the rotor blade azimuth angle. 
FIG. 2 is a graph plotting various maximum sideslip angles (.beta.MAX) 
against airspeed (knots) at a low end of the speed range, and clearly 
shows that the relationship is almost linear. Whilst not shown, this 
generally linear relationship was found to extend throughout the sideslip 
angles of interest in a helicopter up to a maximum of about 34 degrees 
that is representative of an airspeed of about 200 knots. 
FIG. 3 plots sideslip angle (.beta.) against rotor azimuth (.PSI.) with 
wind azimuth .theta.=0, rotor blade tip speed VT=670 feet per sec. and 
sensor location x=0.75, and shows resultant curves in respect of airspeeds 
of 50, 100 and 150 knots. The graph shows that maximum and minimum 
sideslip angles (i.e. the sideslip angles of opposite sign) occur 
approximately midway between the points where the curves cross the axis at 
zero sideslip angle. 
Since equation (1) involves division by a constant plus cyclic term, higher 
harmonics were found to be present in the sideslip angle .beta. and this 
was particularly evident as airspeeds increased above about 50 knots as 
shown in FIG. 3. However, as shown in FIG. 4 for the same design 
parameters the higher harmonics almost disappear and the sideslip angle 
.beta. tends to a simple cosine variation for speeds less than 50 knots. 
FIG. 4 includes curves representative of airspeeds of 10, 20, 30, 40 and 50 
knots and as would be expected with .theta. (wind azimuth)=0, the maximum 
sideslip angles (.beta.) are achieved at 0 and 180 degrees and zero 
sideslip angles occur at the cross over points at 90 and 270 degrees 
azimuth. 
The change in phase of the sideslip angle .beta. due to a change in the 
relative wind azimuth .theta. is shown in FIG. 5 for a relative wind 
velocity V (or airspeed) of 10 knots. The full line is representative of 
the 10 knot airspeed curve of FIG. 4 on a larger scale and the broken line 
represents the curve at the same wind velocity but with the relative wind 
azimuth .theta.=30 degrees. 
As shown, maximum sideslip angles now occur at 150 and 330 degrees rotor 
azimuth and zero sideslip angles occur at 60 and 240 degrees rotor 
azimuth. Thus, there is no variation in amplitude in respect of the 
sideslip angles only a change in phase. 
It will be noted that the maximum amplitude of .beta. in FIG. 5 is almost 2 
degrees which should be easily measurable by appropriate sensors and is 
much larger than any cyclic lag motion of the rotor blades themselves. It 
is anticipated that the blade sideslip angle should be able to be resolved 
to better than 0.5 degrees, in which case the airspeed should be able to 
be resolved within about 2 knots. 
As will now be clear, this invention relies on the measurement of the 
sideslip angle .beta. resulting from radial and tangential flow components 
effective on a rotating rotor blade, and this means that an appropriate 
sideslip detector must be positioned on the blade so as to be insensitive 
to blade angle of attack and away from tip effects. For these reasons it 
may be preferable that the sensor is located on the lower surface of the 
rotor blade and in a region between about a 50 percent blade chord and a 
95 percent blade chord where the local velocity is relatively insensitive 
to angle of attack, and the heat transfer coefficient is almost constant. 
The radial location of the sideslip sensor will be a compromise between the 
need to maximise the resultant sideslip angle by mounting the sensor 
inboard, and the need to minimise higher harmonics and avoid any 
interference effects of the fuselage by mounting it sufficiently far 
outboard. The sensor should therefore be located outboard of a 50 percent 
radial station and preferably at about a 75 percent radial station where 
the influence of the nose and tail boom should produce only a minimal 
disturbance to the sideslip angle, and the effect of trailed tip vortices 
intersecting with the sensing rotor blade will be of a transient nature 
that can be removed by filtering the signal. 
Any suitable sideslip angle sensor can be used although the design must be 
such that an electrical output (volts) is generally linearly proportional 
to the sideslip angle .beta. such as shown in the full line in FIG. 6. 
Some reduction in sensitivity at large sideslip angles may be acceptable 
and can be calibrated out, however, a loss of sensitivity at small 
sideslip angles such as indicated by the broken line is unacceptable since 
it would reduce the accuracy to which the airspeed can be resolved. 
Several alternative instrumentation systems are suitable for use in 
detecting the sideslip angle directly or, alternatively, the radial flow 
component from which the sideslip angle can be calculated. 
Two embodiments of a pressure sensing yaw meter similar to that commonly 
used in wind tunnel testing and adapted to the measurement of blade 
sideslip angle .beta. are illustrated in FIG. 7 in which the illustrated 
portion of blade 23 is shown inverted. Thus, either a small 
aerodynamically shaped blister 32 or a tapered vane 33 is attached 
chordwise to the blade 23 separating pressure ports 34 connected to a 
differential pressure transducer (not shown) for measuring the 
differential pressures due to sideslip. 
A further alternative system for detecting the sideslip angle which 
eliminates possible detrimental effects of centrifugal force is 
illustrated in FIG. 8 and consists of the use of hot film gauges 35 
positioned on opposed surfaces of an aerodynamic blister 36 to measure the 
differential heat transfer created by the airflow about the blister 36 on 
the surface of blade 23. The gauges may be manufactured integral with the 
aerodynamic blister 36 and coated with a protective layer to prevent 
contamination. 
In an unillustrated embodiment, temperature sensitive transistors could be 
used in similar locations to the gauges 35 of FIG. 8, to detect the 
difference in heat transfer on each side of the aerodynamic blister 36 
from which the sideslip angle .beta. can be calculated. 
The aerodynamic blister 36 and hot film gauges 35 or temperature sensitive 
transistors and necessary electrical wiring can be retro-fitted to 
existing rotor blades or incorporated during manufacture of new blades. 
Another type of hot film device capable of measuring both the direction and 
magnitude of the flow over a surface consists of twin hot film filaments 
37 having a length to width ratio of about 100 and mounted in the shape of 
an inverted `V` as shown in FIG. 9. Such a gauge detects the flow 
direction over a smooth surface by measuring differences in the thermal 
boundary layer, and the heat transfer from the gauges and hence their 
output follows a cosine relationship to the incident flow (with the cosine 
raised to a fractional power), from which a differential signal can be 
obtained. 
Such a hot film `V` gauge is considered particularly well suited for use in 
this invention since the twin gauges 37 can either be mounted directly on 
the blade surface, making retro-fit possible, or ultimately formed during 
the manufacture of the blade. An analysis of the characteristics of such 
devices suggests that all optimum configuration can be achieved which will 
both maximise sensitivity in the range of sideslip angles of interest and 
that the output will be linear within that range. 
Another method for detecting the radial flow along a rotor blade from which 
the sideslip angle .beta. can easily be calculated is the use of 
ultra-sonics. A possible installation is shown in FIG. 10, in which the 
rotor blade 23 is again shown inverted, as comprising two ultra-sonic 
transducers 38 (shown exaggerated in size for clarity) spaced-apart 
chordwise on the lower surface. 
This method has the advantage of not being affected by either centrifugal 
forces or contamination or bridge balancing requirements, and would be 
very robust. Since the transducers 38 are quite small, the system may be 
self contained with the radio transmission electronics included in a 
detector housing. 
Another possibility for detecting the direction of the airflow is the use 
of laser light. Small laser diodes could also be arranged to detect the 
radial flow velocity. 
The electrical signal from the blade sideslip detectors (of whatever form) 
is transmitted from the rotating blades to a signal processing unit P in 
the aircraft fuselage. This can be accomplished using slip rings but a 
system based on non-contact transmission using radio or other means is 
preferred. The transmission system can be small, lightweight, robust and 
inexpensive. The necessary electrical wiring can be built into new blades 
during manufacture, and sensor and transmission electronics could be 
integrated. 
The signal processing unit will contain information in suitable form 
concerning the relationship between sideslip angle .beta. and airspeed. 
Signal processing is straightforward since it is only necessary to filter 
the signal from the sideslip detectors and determine its phase and 
amplitude. At low speeds (as illustrated in FIG. 3) the signal closely 
follows a cosine wave with its phase and amplitude related directly to the 
relative wind direction and airspeed respectively. The phase angle can be 
established by detecting the zero sideslip angle crossing points in the 
rotor azimuth so as to minimise the effects of any higher harmonics 
present at high airspeeds. Since the amplitude and phase of the harmonics 
depends only upon the advance ratio, the processed signal can be readily 
calibrated for use also at high forward speeds. 
The resultant processed signals relating to both airspeed and relative wind 
direction can be incorporated into existing instrumentation in a cockpit 
for display to a pilot. Alternatively, a dedicated instrument as 
illustrated in FIG. 11 could be used. The instrument 39 includes digital 
displays 40 and 41 for displaying the relative wind azimuth and airspeed, 
and a central circular display screen 42 including a movable bar 43 
showing the direction of the relative wind from the longitudinal 
centreline of the helicopter as represented at 26. The length of the bar 
43 varies with airspeed so as to provide, in combination with concentric 
rings 44 representative of different airspeeds, a visual indication of 
airspeed to the pilot. 
While several embodiments have been described and illustrated it will be 
understood that many modifications may be made without departing from the 
scope of the invention. For example, the method and apparatus can be 
applied to rotary wing aircraft having any number of rotor blades. Any 
suitable detection means can be used and a plurality of sensors could be 
used on each blade to extend or optimise the range and sensitivity of the 
system. In such an arrangement, automatic switching between the plurality 
of sensors could be incorporated to cater for different phases of 
operation. In some installations, the or some of, the sensors may be 
located on an upper surface of the rotor blade.