He Whareleura-tini Kaihautu 0 Aotearoa
THE OPE N POLYTECHNIC OF NEW ZEALAND
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Bas/0 He//copier Aerodynamics 555—3— 10
CONTENTS %
Basic Helicopter Flight Aerodynamics Hovering
1 =
2
1
Transition from Hover to Forward Flight Translational Lift
3 4
Transition from Forward Flight to Hover Power Required Power Available
5 7 9
Forward Flight
12
Dissymetry of Lift
12
Limits of Forward Speed
15
Stability
17
Cyclic Control Forces Vortex Ring
18 19
Control on the Ground
23
Ground Resonance Taxying Blade Sailing
23 25 26
Centre of Gravity
26
Autorotation
28
Autorotative Force
29
Forward Speed
32
All-up Weight
3%
Altitude
35
Range and Endurance
E
x
35
Copyright This material is for the sole use of enrolled students and may not
be reproduced without the written authority of the Principal, TOPNZ. 555/3/10
'32
AIRCRAFT ENGINEERING
HELICOPTERS
ASSIGNMENT 10
BASIC HELICOPTER FLIGHT AERODYNAMICS
Hovering when the helicopter is at rest on the ground with rotor rev/min set at the normal take—off figure, the lift resulting without collective pitch is negligible. In this condition, the only effective force acting on the aircraft is that of gravity acting on the mass. The only reason that this unbalanced force does not produce movement is because the ground supplies an equal and opposite reaction. As collective pitch is applied and the rev/min kept constant, so the lift is increased and the weight is taken off the wheels. The reaction from the ground is reduced, but there is still no movement of the aircraft. when the lift exactly balances the weight, a new state of equilibrium has been created, with the aircraft at rest and with no reaction from the ground. As pitch is further increased, lift exceeds weight and the excess force creates an acceleration upwards (F
=
Ma).
That
is, the aircraft will now climb vertically, given perfect still-air conditions.
As the fuselage starts to move, parasite
drag results and must be added to the weight. A new stage of equilibrium will be reached at the climbing speed, where parasite drag is equal to the excess of lift over weight. To achieve hover, pitch is reduced until the lift again equals weight.
The parasite drag then decelerates the rate of
climb,at the same time, itself reducing to zero. A new state of equilibrium is then reached, with lift equal to weight and the aircraft stationary at the required height.
10/9i
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This, then, represents the perfect hover, assuming no outside interference, wind, and so on, with lift exactly balancing weight. Ideally, there should be no further movement of the collective, cyclic, throttle, or tail rotor controls to maintain a constant position. In practice, however, small control corrections must constantly be made to keep an accurate hover.
If power is now reduced, the aircraft will descend as a result of excess weight over lift. The descent will again be an acceleration until the parasite drag from the fuselage once again equates the forces. The descent is then at constant speed. Ground effect: As a slowly descending helicopter nears the ground, its rate of descent reduces, and it may even come to a hovering attitude ——-even though no changes to the collective and throttle controls are made.
This phenomenon is caused by
ground effect.
The effect is brought about by an increased pressure area being created between the rotor disc and the ground as a result of the normal downward flow of air through the disc being slowed by the ground immediately below the rotor. The effect is sometimes called ground cushion because the impression is of the aircraft "sitting" on an air cushion. The closer the rotor is to the ground, the more the air will tend to be trapped and slowed and therefore "cushion" the aircraft. That is, the closer the aircraft is to the ground, the greater will be the ground effect and therefore the lesser the power required to hover. Because ground effect decreases with height above the ground, it is not easy to state positively a height at which the effect will be negligible. For practical purposes, the ground cushion is taken as the rotor height above ground equivalent to the length of one rotor blade or one half of the rotor diameter.
Thus, the larger the rotor,
the thicker the ground cushion.
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_ 3 _ The second factor affecting ground effect is the nature of the ground itself. Because the effect depends upon the slowed air maintaining a streamline flow, the smoother, firmer, and more level the ground is, the greater the effect will be. Thus, a level stretch of tarmac or concrete will give maximum effect. Long grass, small bushes, or uneven ground tend to break up the smooth flow of air and reduce the effect.
Sloping ground causes
an inequality of ground effect round the disc and hence some tendency for the aircraft to "slide" down the slope. A similar result will occur if the disc is not parallel to the ground, for example, when hovering in a wind or in the transition from the hover to translational flight. A wind tends to displace the "cushion" downwind of the helicopter. Re-circulation: Some energy is lost by the spillage of air around the tips of the blades. This can be aggravated when hovering near the ground, particularly if some obstruction fairly near to the rotor, such as a hangar door or a high building, which causes the air, after passing through the rotor, to re-circulate down through the rotor again. This detracts from the ground effect and, when the obstruction is only on one side of the rotor, causes an inequality of lift around the disc
so that the aircraft tends to "creep in" towards the
obstruction.
Transition from Hover to Forward Flight This transition, which will be made nearly every time the helicopter is flown, is usually accomplished as the helicopter is climbing from its take-off site. However, it can be a manoeuvre during flying training, and so we'll consider the theory from the point of view of keeping the lift factor constant. ' In Fig. l, the centre line shows the flight path. The parallel dotted lines show that lift is equal to weight so that a constant height is maintained.
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- u _ Figure l (a) shows the helicopter in a perfect hover (still air) within the ground cushion. The lift is shown as a combination of power and ground effect. To achieve forward flight from hover, you need to tilt the disc forward with a forward movement of the cyclic stick, and so create a thrust force in the required direction. See Fig. l (b) You then need to increase the size of the useful force to keep the lift component equal to weight. The tilting of the disc also causes some loss of ground effect, requiring yet more power to compensate. In fact, power is normally increased to maximum to ensure no loss of height. This can be a hazard operationally if the power margin is small. It may be possible to hover, as the result of ground effect, when at nearly full power, but the sink caused by the change of disc attitude may be such that not enough power is left to prevent the aircraft sliding off the ground cushion and so striking the ground. The thrust force created by tilting the disc now causes the fuselage to accelerate in the direction of disc tilt, that is, forwards in this case. Acceleration will continue until the parasite drag of the fuselage balances the thrust component —— Fig. l (c). An equilibrium state is established, and speed will now remain constant. The speed at which this occurs will depend on the amount the disc is tilted and whether there is enough power to provide the necessary useful force at this disc attitude.
Translational Lift As speed increases, power may be reduced as a result of translational lift. This is additional lift created by the rotor, at given pitch and power settings, when moving forward, as a result of the increased mass flow of air now passing through the disc in a given time.
Less power is required to
produce a given force if a large mass is given a small acceleration compared with a small mass being given large acceleration. If power is not reduced but level flight is maintained by moving the stick forward, not only will the forward
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_ 5 _
speed increase, but the rotor rev/min will also increase as a result of the extra power available. Any further increase in speed will require a disproportionate increase in power to compensate for parasite drag, which rises as the square of the speed —— Fig. l (d). 1 GE—wsroonde§kxi
P—Power
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Transistion from hover to forward flight
Transition from Forward Flight to Hover Figure 2 shows the change of rotor attitude from forward flight condition to induce rapid deceleration called a flare. The effects of the flare are:
l.
An increase in the useful force as a result of the increase in the angle of attack of the disc (forward movement is maintained). This is comparable to the increase in angle of attack of a fixed—wing aircraft in a steep turn or pulling out of a dive.
2.
A reversal of the direction of the thrust component, causing rapid deceleration.
3.
Some tendency to cut off the inflow of air, hence partly offsetting the effect described in l.
4.
An increase in rotor rev/min. This is surprising because an increase in angle of attack would suggest an increase in drag and therefore a decrease in rev/min. However, the important factor is the relationship of the direction of total reaction to the plane of rotation. Figure 2 shows that, as a result of the flare, total reaction has moved forward relative to the plane of rotation, thus causing an increase in rotor rev/min.
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_ 5 -
RA — Rclaiive, airflow’
PR - Plal/H2. of roiaiion
L — LIFT‘
L
A — Pitch angle
TR
RPD — Rotor profiie drag
TR - Total FQQCIIOH
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In forward powered flight, the relative airflow enters the rotor disc from above the plane of rotation. it enters from below.
During a flare,
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TO return to a state of hover from a cruise state, you could just move the cyclic control back to tilt the rotor disc to the hovering attitude. The aircraft would then lose speed as a result of parasite drag. However, a more rapid deceleration is usually required so the helicopter is normally flared. As the result of the temporary increase in angle of attack, the aircraft will tend to climb unless power is reduced.
You
may also need to throttle back to prevent the rotor from overspeeding due to the increase in rev/min. However, the air inflow soon becomes stabilised, and translational lift decreases as the aircraft decelerates rapidly as a result of both parasite drag and reverse thrust. Power must then be increased considerably as the aircraft is coming to rest so as to prevent it from sinking. At this stage, you must maintain rotor rev/min. Finally, as all forward speed is lost, you need to restore the disc to the correct hovering attitude to prevent the aircraft from moving backwards and you should reduce power again, not only because of the disappearance of the thrust component but also because of the re—establishment of the ground cushion. See Fig.
3. 555/3/10
_ 7 _ We have assumed throughout the transition that there is no wind effect. In practice, the transitions will normally be made into wind.
NOTE:
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Transition from forward flight to hover
Power Requ ired Large changes of power must be made to maintain level flight
with varyi ng forward speeds. In practice, only one change of power is made, by either collective lever or hand throttle or both, while in theory, the power required for level flight can be divided into three components
l.
¢ ¢
Rotor profile power:
The component Of total
reaction acting in the plane of rotation, called rotor profile drag, must be overcome if rotor rev/min are to be maintained. The power to do this is called rotor profile power.
2.
Induced power:
TO create lift, you need
to cause a flow of air through the rotor by applying pitch, that is, giving the air an induced velocity. The power needed to cause this airflow is called induced power. As pitch is increased, the rotor profile power will also increase. Provided enough power is available to produce lift and still maintain rotor rev/min, the helicopter should be able to hover. The factors of power needed to drive the tail rotor and the cooling fan are included in the total of rotor profile power.
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_ 3 _ 3.
Parasite power: When the helicopter moves, the airflow over the fuselage meets a resistance to its passage. This is called parasite drag, and the power needed to overcome this drag is called parasite power. Figure H shows the variations of these three components of power required through the speed range. Although the buildup of forward speed causes complications of the airflow relative to the
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rotor disc, the total effect on rotor profile drag is not pronounced until the higher speeds are reached.
FIG. 4
Components of power required
Because of this,
rotor profile power is shown as
a slight curve, gradually increasing with speed. As forward speed is increased, a greater mass of air will pass through the rotors.
Because a large mass of air needs
less acceleration to provide effective force than a small mass, the induced power can be decreased. In other words, as forward speed increases, rotor efficiency also increases. In the transition stage, however, more induced power is needed to supply the required thrust force and to compensate for the loss of ground effect. Parasite drag tends to increase in ratio to V2. For example, if we have one unit of drag at 20 m/s (V), we would have four units of drag at HO m/s and nine units at 60 m/s. Thus, parasite power is shown as an ever-increasing curve with speed increase. Figure 5 shows the total power required for level flight. The power required curve shows the initial demand for induced power in the transition from the hover, the subsequent decrease as forward speed is built up, and the rapid increase of parasite drag with higher speed.
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_ 9 _
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Power
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Ia) Lightly loaded helicopter (b) Heavily loaded helicopter FIG. 5 Comparison of power required with power available
Power Available with fixed-wing, piston—engined and turbine~engined aircraft, the power available curve rises to a maximum and then falls or keeps rising respectively.
with the helicopter's
constant induction manifold pressure or fuel flow, rev/min, and altitude, there is no variation of power available throughout the speed range.
Hence, it appears as a straight line.
More important is its relative vertical position in terms of the power required curve, that is, whether the availability meets the requirement. This may be regarded as one of the essential factors in all helicopter operations (the power/weight ratio). If you compare the two curves in Fig. 5, you can see that if the power available is above the power required throughoutFig. 5 (a)-—-thennot only are hover, transition, and full speed possible, but a climb is also possible at any speed, the rate of climb being determined by the excess power and the weight of the aircraft. If, however, the power available curve cuts the power required curve—Fig. 5 (b)-then it is not possible to hover, and you may need to make a "running" take—off or its equivalent,that is, make use of translational lift, to compensate far the pgwep shortage, The maximum forward speed will also be reduced.
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_ 19 _
Overpitching: Where power may be marginal, during take-off, hover, or transition, overpitching may occur. This means that so much collective pitch has to be applied to produce the required lift that not enough power is left to overcome the high rotor profile drag. As a result, the rotor rev/min will fall in spite of the hand throttle being fully open, and the aircraft will start to sink. The application of more pitch would obviously only aggravate the situation, and the only remedy is to lower the lever, keeping the throttle fully open, and accept the resultant sink until correct rev/min are recovered. You may even have to put the aircraft back on the ground and reduce weight. NOTE:
True overpitching occurs only when the throttle is fully open, though similar symptoms can result from a rapid lever movement without adequate throttle lead, for example, in the transition to the hover. In this case, the remedy is to open the throttle.
The factors governing power available and power required are l.
Altitude, temperature, and humidity:
These
three factors can be dealt with under one heading because the basic problem is one of air density. The reason for applying pitch to the rotor is to accelerate a mass flow of air through the rotor. A given pitch setting at constant rev/min will, however, accelerate a given column of air, and the mass flow will therefore depend on the density mass
=
volume X density
Any reduction in air density will therefore reduce its mass flow and the resulting lift, so requiring an increase in pitch and, consequently, power to balance the weight. An increase in altitude, temperature, or humidity will, in each case, cause a decrease in density and therefore in the performance of the rotor.
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_ 11 _ Each of these factors also affects the engine performance resulting in a decrease in the power available at the same time that the power required is increasing. The humidity used is the specific humidity, that is, the actual amount of water vapour present, as opposed to relative humidity. 2.
Wind effect: The effect of moving the rotor disc through the air is to increase the mass airflow through the rotor and so decrease the power needed. Wind blowing through the rotors will have much the same effect as rotation of the rotors.
3.
A11—up weight: Because more pitch must be applied to supply an increase in lift (assuming constant rev/min), the factor most affecting power required is all—up weight. Thus, the ratio of weight to power is of great importance in all helicopter operations.
SUMMARY
Ground effect, or ground cushion, is said to extend vertically upward to a rotor height above ground of half the rotor diameter. Power may be reduced once translational lift is gained. Parasite drag increases very rapidly with increasing air speed. In forward—powered flight, the relative airflow enters the rotor disc from above the plane of rotation. when flared it enters from below.
PRACTICE EXERCISE A State whether each of the following statements is true or false. l.
In a hover in still air, lift thrust = drag.
2.
The type of ground surface has a great influence on the strength of the ground cushion.
3.
Translational lift will increase parasite drag, and so more power will be needed.
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=
weight, and
_ 12 _
4.
Induced power required increases with forward speed.
5.
When the power available is just enough for hover, a transition to level flight is safely possible.
6.
An increase in AUW will lower the maximum forward speed.
7.
Air density affects power available but not power required.
8.
Wind effect increases the mass airflow through the rotor.
9.
In level flight, the relative airflow enters the rotor from below the plane of rotation.
10.
Ground effect extends vertically to a rotor height above ground of one rotor diameter, and the ground cushion extends to one half-rotor diameter.
(Answers on page 39)
Forward Flight
As soon as forward speed is gained, the effect of dissymetry of lift is felt. This effect, which has an important bearing on the cyclic control of the helicopter, also imposes a limit on the forward speed.
Dissymetry of Lift When the aircraft is moving forward, a rotor blade in the 180° of rotation from the tail cone to the nose is said to be advancing, and this side of the rotor disc is called the advancing gigg.
From the nose back to the tail cone, the blade is said to
be retreating around the retreating iide of the disc.
If a
two-bladed rotor is considered, the maximum effect of the forward speed will be experienced with the blades athwartships (Pig. 6). On the advancing blade, the relative airflow is the sum of the
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_ 13 -
effects of rotational velocity and forward speed (V1 + V2). On the retreating blade, it is the difference between the two (V1 — V2).
Because airflow affects lift, then, given equal pitch,
the advancing blade has more lift than the retreating blade, and the disc therefore tends to roll to the retreating side.
The
formulas are
Lift (advancing blade)
=
CL kc (V1 + V2)2 S
Lift (retreating blade)
=
CL tp
(V1 - V2)2 S
This is known as dissymmetry of lift.
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Relative airflow on advancing and retreating blades
In fact, the disc would not roll to the retreating side because the blades have been given freedom to move about either a gimbal ring (two-bladed rotor) or flapping hinges (three or more bladed rotor). As soon as the dissymmetry of lift occurs, it causes flapping to take place, the advancing blade flapping up with increase in lift, and the retreating blade downwards due to decreased lift.
This movement about the flapping hinge
causes a further element of relative airflow either up or down, thus altering the angle of attack to compensate for the effect of V2 and so restoring equality of lift (Fig. 7). The flapping caused by dissymmetry of lift will also cause a change of disc
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_ lu _ attitude, which will be maximum 90° on from the athwartships position (phase lag), that is, fore and aft, the advancing blade rising in front and the retreating blade falling at the tail. This tilting back of the disc as a result of forward speed is called flap—back.
See Fig.
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Change of angle of attack due to flapping
If flap-back is allowed to occur, the direction of useful force will change, as will the components of lift and thrust,and equilibrium in forward flight would be impossible. To maintain a steady forward flight attitude, you therefore need to prevent flap—back from occurring because of V2. The equality of lift between advancing and retreating blades is achieved by a cyclic pitch change. When the cyclic control is moved forward to begin forward flight, you must move the stick further forward as soon as the effect of V2 becomes apparent This means that pitch is reduced where (V1 + V2) is maximum and increased where (V1 - V2) is minimum. '
Thus, an equality of lift is effectively maintained
throughout, so preventing any flapping from taking place. There is now an angular difference between the plane of the control orbit and that of the rotor disc, which corresponds to the angle of flap—back of the disc when flapping was allowed to occur. See Fig. 8. This means that the correction for V2 has been made by feathering and not by flapping. If the speed is such that the forward limit of cyclic stick is reached, then it is impossible to prevent flap-back resulting from any further increase in speed.
This could represent a
limitation of forward speed.
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_ 15 _
Disc
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Limits of Forward Speed
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The factors limiting the maximum forward speed of the helicopter are
1.
Limit of cyclic control,
2.
Reversal of airflow,
3.
Stalling of the retreating blade,
H.
Compressibility,
5.
All-up weight, and
6.
Altitude.
Limit of cyclic control:
When the forward limit of the
cyclic control is reached, it is no longer possible to counteract flap-back, and so any further increase in forward speed is prevented. Offset flapping hinges, a stabiliser at the tail, and a delta hinge effect incorporated in the pitch operating arm, have all been used to overcome cyclic control limits, with the result that, under normal conditions, the cyclic control is unlikely to run out of movement before other factors impose their own limits.
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_ 15 _ gr
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Reversal of airflow:
ewe
considering dissymmetry of lift, we compared rotational speed (V1) and forward speed (V2). However, this comparison is true only at any one station along the blade
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the retreating blade, the relative airflow (V1 — V2) can become very small, and at a high V2, a negative
imoo
quantity.
Reversal of airflow
The area of the blade
affected in this fashion spreads from the root as V2 increases.
Figure 9 shows the exaggerated effect. Stalling of the retreating blade:
Probably the most important
factor limiting forward speed is the stalling of the retreating blade by the cyclic pitch applied to prevent flap-back.
Because
this cyclic pitch change is added to the collective pitch already
applied to provide lift, it is easy to reach the critical angle of the blade and so cause it to stall.
The symptoms and effects of this stall are similar to those of the usual fixed-wing stall, that is, judder, loss of lift, and an increase in drag. No two helicopters react to blade stall in exactly the same way. Usually, the onset of vibration and erratic cyclic control forces signals the start of the condition. As stalling continues to move inward from the tip area, vibration increases, followed by a partial loss of control and a nose-up pitching tendency.
Severe stalling may result
in large rolling tendencies and complete loss of control. Retreating blade stall imposes a limiting V2 on all helicopters for any given conditions.
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It
_ 17 _
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Compressibilitg: The upper limit for rotor rev/min is reached when the advancing blade tip speed reaches the compressibility region, with vibration and loss of efficiency. To provide a given amount of lift, either a small rotor may rotate quickly, or a large rotor rotate slowly, the latter being normally the more efficient. In either case, the tip speed will be similar, and compressibility is usually inevitable when the forward speed (VQ reaches about 200 knots. All—up weight: The higher the all-up weight of the helicopter, the greater must be the collective pitch applied to lift that weight. The greater the collective pitch applied, the less cyclic pitch can be used before the stall angle on the retreating blade is reached. That is, the higher the all-up weight, the lower the limiting speed V2. Altitude: As the rotor rev/min must remain more or less constant, the helicopter will lift less as the altitude increases because of the reduced air density, and so a greater pitch angle is needed to lift a given weight. Thus, the higher the altitude, the greater the collective pitch for a given weight and the lower the limiting speed at which the critical angle will be reached, that is, the lower the IAS for retreating blade stall. For the same reasons, the cyclic control movement at altitude becomes less effective in terms of disc movement.
To
counteract the flap-back arising from a given forward speed (IAS) at increased altitude, you need to make a larger forward control movement. Consequently, the limit of forward control will be reached at a lower limiting speed. Stability If an aircraft in flight tends to return to its original %
~§
position after being disturbed, it is said to be stable. If it remains in its new position, it has neutral stability, and if it departs farther and farther from its original patch, it is unstable. The helicopter is unstable in flight so far as changes of discattitude are concerned because a change of attitude (flare)
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_ 18 _ of the disc causes a change in the useful force in both magnitude and direction. This will, in turn, have a further effect on the disc.
In general, the discattitude must be controlled by the
pilot at all times to prevent the helicopter from going out of control.
Cyclic Control Forces In forward flight, a lateral force, arising from the reaction of the control orbit to the forces acting in the pitch operating arms whenever the control orbit is displaced from the perfect hover attitude, can cause the aircraft to roll. Forward flight implies that the control orbit is tilted forward, and so the maximum forces being exerted in the pitch operating arms
to cause a pitch change are more or less on the lateral
axis of the aircraft. Because there is a downwards force on the advancing side and an upwards force on the retreating side, the equal and opposite reaction from the control orbit will tend to tilt it towards the retreating side, which would cause the aircraft to roll in that direction. To prevent this, the pilot would need to push the cyclic control in the opposite direction. See Pig. l0. MAX In aircraft with manual control, this load can be relieved with an adjustable
M‘
CONTROL ORBIT TILTED FORWARD FOR FORWARD FLIGHT
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Control orbit reaction in forward flight
spring loading in the control system to hold the control column in a given position. The amount of spring loading needed to relieve the load will vary with the forward speed. Another way of preventing these loads being . felt by the pilot is to use irreversibles in the fore and aft and lateral control runs.
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_ 19 _ In aircraft with servo controls, the load is not normally
felt by the pilot because the forces cannot be fed back through the servo jacks. In the event of a hydraulic system supply failure, however, the load is felt suddenly and the helicopter tends to roll violently to the side of the retreating blade. One helicopter maker used lock—load valves on the servos to prevent those feedback forces being felt, but other helicopter makers relied on the pilot reducing speed to where the forces became acceptable. To prevent or reduce the effects of a hydraulic failure, a duplicate servo system may be provided and powered by a different hydraulic system.
Vortex Ring Another name for vortex ring is settling with power. names are freely used to describe the same condition.
Both
In normal powered flight, there is an induced flow of air downwards through the rotor. In the event of a fuselage movement normal to the rotor disc, it is possible to set up an airflow relative to the disc directly opposed to the induced flow, which therefore causes a very confused pattern of flow round the rotor. This movement of the air on rotor blades at high angles of attack will stall the blades at the hub. This stalling can move outward along the blade as the rate of descent increases. In particular, a turbulent vortex called a vortex ring, is created around the periphery of the disc. See Fig. ll. The combination of conditions in which a vortex ring is likely to occur are during
l.
Powered flight with induced flow through the rotor,
2.
Movement of the aircraft causing a relative flow normal to the disc from the opposite direction, and
3.
Relatively still air conditions.
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_ 29 _
..
The first example in which the conditions could be met is
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FIG. ll_
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a vertical descent with power. Vortex ring is unlikely to occur in normal conditions until the rate of descent exceeds 300 ft/min. As a safety margin, a rate of 250 ft/min should not be exceeded in a vertical descent in still air.
Vortex ring
The second example is in a fairly steep flare, where the aircraft is also being allowed to descend. If power is applied in this condition, vortex ring can result. Another example is when power is applied in recovery from an autorotation. If the path of movement is normal to the disc, for example, during descent in autorotation, and power is applied without a change of disc attitude, then vortex ring can result. In all cases, the question of wind must be considered. A head wind of 10 knots or more will normally be enough to prevent a vortex state from being reached. Remember, however, that it is possible to lose this headwind component quickly, for example, descending into a clearing or below high buildings when a vortex state may result if the rate of descent is too rapid. Conversely, it is possible to be descending downwind with ground speed clearly apparent, but with the rotor descending in relatively still air and therefore subject to vortex ring. The effects of vortex ring are similar to those of a fixedwing aircraft in a stalled condition. Initially, the disturbed flow may result in buffeting and vibration, but these may disappear in a complete vortex state. There will be a considerable loss of lift, which will result in the aircraft accelerating in its original direction of movement, and a large increase in drag, which will cause loss of rotor rev/min and possible loss
555/3/10
_ 21 _ of directional control.
Finally, the high rate of descent may
have enough weathercock effect on the tail cone to cause sharp nose—down pitching, which will, in fact, destroy the vortex condition. As the vortex condition arises from the confused pattern of airflow, the introduction of a new component of airflow will help to remove the vortex state.
This is most easily achieved
by a change of disc attitude, that is, by a cyclic control movement. In practice, the control column is normally moved forward to induce translational flight, but if this is impossible a lateral or backward movement can be used. As soon as the disc attitude is altered, full power may be applied to minimise loss of height. Power should never be increased before the control column is moved because this may aggravate the vortex condition and so increase the rate of descent. It would also be possible to change the flow by lowering the lever and autorotating, but the loss of height in effecting recovery in this way would be considerable. Correct recovery from a full vortex state will require approximately 300 feet. Thus, vortex ring below this height is very dangerous, although, in practice, it is at such heights that the danger is most likely to arise, for example, descending into confined space, "quick stop", recovery from autorotation, and so on. It may be compared with the stall on the approach at low altitude in the fixed-wing aircraft.
1
SUMMARY‘ Dissymetry of lift occurs imediately a horizontal airflow passes across the rotor disc. Dissymetry of lift is countered by blade flapping plus cyclic feathering, which restores equality of lift. Flap-back, which occurs because of blade flapping, is controlled by moving the cyclic control column forward.
555/3/10
-22..
Forward speed is limited by l.
Range of movement of the cyclic control,
2.
Reversal of airflow in the rotor disc,
3.
Retreating blade stall,
4.
Compressibility at the rotor blade tips,
5.
All—up weight, and
6.
Altitude.
V
The helicopter is unstable in flight and must be controlled at all times. A vortex ring state is an unsafe condition of flight.
PRACTICE EXERCISE B State whether each of the following statements is true
or false. l.
An advancing blade of a helicopter in forward flight experiences a decreased airflow over its surfaces.
2.
Reversal of airflow occurs on the advancing half of the rotor disc.
3.
An upward—flapping blade has a decreased angle of attack because of the flapping motion.
4.
Blade stall does not affect the limit of the forward—flight speed.
5.
Compressibility at the advancing blade tip will limit the rotor rev/min.
6.
Because the air is less dense at altitude, the helicopter will experience less drag and will have a higher limiting speed.
7.
Because the helicopter is "suspended" from its rotor, it has natural stability.
8.
Vortex ring state is likely to occur if the rate of descent is more than 300 ft/min.
9.
Flap—back is corrected by a slight increase in collective pitch.
sss/s/10
...23...
10.
When the all-up weight of the helicopter is increased, the limiting speed is reduced.
Lanswers on page 40)
CONTROL ON THE GROUND When the helicopter is on the ground with its rotors turning several serious problems can arise. Chief among them, for helicopters with articulated rotor heads, is ground resonance, which, if left unchecked, will cause the complete destruction of the helicopter in a few action—packed seconds. Ground Resonance Ground resonance is a severe low-frequency vibration resulting from a forced or self-induced vibration of a mass in contact with the ground. In the case of the helicopter (the mass), the vibration can originate either as a disturbance in the rotor transmitted to an undercarriage in contact with the ground, or in the undercarriage itself due to mislanding, rough ground, and so on. In either case, an element of sympathy must exist between the original vibration and the natural frequency of vibration of the other system. Rotor vibration can arise from any basic unbalance of the rotor, for example, blades of unequal weight or with their centres of gravity unequal distances from the centre of the rotor, or blades producing unequal lift or with their centres of lift unequal distances from the centre of the rotor. However, before being mounted on the helicopter, a set of blades is usually balanced, and so the most likely cause of rotor vibration is faulty drag dampers. Drag dampers are incorporated to control the rate of movement about the drag hinges. If they are set incorrectly, the blades will move about the drag hinges at different rates and so cause blade unbalance.
555/3/10
See Fig. 12.
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(b) Vibration from disc unbalance
A likely cause of rotor vibration is mishandling of the cyclic control by the pilot. "Stirring the stick" while the wheels are on or very near the ground should therefore be avoided. A similar effect can result from an inexperienced pilot trying to be too careful about his landing and only succeeding in touching first one wheel and then the other, so setting up a "padding" of the undercarriage. It is easier for a vibration to be set up in the undercarriage, when there is no weight on the wheels, that is, when the collective pitch is quite high. This state should be avoided as much as possible by lowering the collective lever as soon as the wheels touch on landing, and by making a smooth progressive increase in pitch when taking off, to take the aircraft well clear of the ground.
555/3/10
g
_ 35 _
The nature of the ground that the helicopter is resting on can influence ground resonance. For example, on landing, one wheel may slip into a concealed hole or rut, and even this small movement might set up the required initial vibration. Forward movement over rough ground would naturally increase the risk of vibration, and landing across sloping ground can also have the same effect, particularly if the pilot is unaware of the slope when landing. Because the sympathetic frequency of vibration of the rotor and the undercarriage is an essential feature of ground resonance, designers choose undercarriage systems that minimise the possibility of such a sympathy being set up.
If oleo
extensions and type pressures are kept to the correct figures,
ground resonance becomes less likely. In the event of resonance occurring, the best recovery action is to take off immediately to hover, where the vibration should die out quite rapidly. To allow for immediate take-off, keep the rotor rev/min up to the take-off figure all the time there is any possibility of ground resonance.
Should it be
impossible to take off, then the sympathy between rotor and undercarriage should be destroyed by reducing the rotor rev/min as quickly as possible, that is,collective down, throttle closed, switch off, rotor brake on.
Taxying The thrust for-taxying is provided by the main rotor, with
the lift component kept to a minimum to avoid ground resonance. Aim to have the best thrust/lift ratio without having
an exaggerated forward attitude of the disc and the possibility of the blades striking their lower stops.
Keep rotor rev/min
at the flight figure and steer by using the wheel brakes. Keep taxying speed down to a walking pace, and avoid rough ground.
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_26_
Blade Sailing Blade sailing occurs when the rotor is either starting up or slowing down in strong, gusty winds. In this event, a dissymmetry of lift is experienced between advancing and retreating blades similar to the effect in forward flight. The advancing blade flaps up at the front and the retreating blade down at the rear. If this motion becomes exaggerated (particularly if it becomes in phase with the natural frequency of vibration of the blade), it can result in damage to the fuselage. In extreme cases, blades have been known to strike the ground. The flapping can be countered to some extent by a small forward movement of the stick, but be careful, particularly in gusty conditions, because a sudden reduction of wind might cause the blade to flap down in front to a dangerous extent. Because the blades pass nearest to the fuselage when
crossing the tail cone and the blade will be at its lowest point downwind, it may be advisable to turn the aircraft H5° out of wind, so that the blades pass their lowest point well clear of the tail cone. Most helicopters incorporate droop stops, which are held out of position by centrifugal reaction above approximately 100 rotor rev/min but fall into position below this figure and restrict the downward droop of the blades. Many helicopters also have flapping restrainers, which prevent the blades from flapping up or down below a fixed rotor rev/min. However, even with droop stops and flapping restrainers, be careful when
engaging and disengaging rotors in a high or blustery wind.
CENTRE OF GRAVITY (c.e.) The centre of gravity of an aircraft is the point through which the total of weight forces act.
It is normally calculated
by reference to the moments of the various weight forces around a given datum, which in the case of the helicopter, is usually the centre line of the rotor. The position of the c.g. is then
555/3/10
_ 27 _
quoted as so many inches or centimetres fore or aft of the datum. In the helicopter, the relationship between the useful force and the weight will affect the behaviour of the aircraft in all stages of flight. So far, we have assumed that the centre line of the rotor passes through the c.g. and that the useful force directly opposes weight. If, however, the c.g. lies either fore or aft of the datum, then the resulting couple will cause the aircraft to adopt a nose-up or nose-down attitude. The disc has to be maintained in its correct position in space by movement of the cyclic stick until the line of useful force passes through the new c.g. and a new state of balance is reached. See Fig. I3.
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.
(a) forward FIG. 13
,/
. (b) aft Fuselage attitude with extremes of c.g.
t
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Two results follow from this. First, the range of movement of c.g. may be limited by the amount of cyclic stick control. Secondly,an incorrect position of c.g. will limitinanoeuvrability in a given direction, for example, limitation of forward speed.
Centre-of-gravity Limits The theoretical limit of movement of c.g. will be governed by the extremes of disc attitude because it must be possible for the line of useful force to pass through the c.g. if control is to be maintained. See Pig. lH. In the case of the twin—rotor helicopter, the range lies midway between the rotors and is longer than in the single-rotor type.
555/3/10
- Q8 _ The limit increases with the distance between the rotor head and the c.g. That is, the lower the load and the higher the rotor head, the wider the limit. A similar effect is achieved by the use of offset flapping hinges, which give a more effective disc response in terms of stick movement and therefore increase the c.g. limits. The practical range laid down by the designer is naturally much less than the theoretical limit in order to ensure manoeuvrability and is specified for each type
of aircraft in terms of the datum.
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(b) Fore and aft range
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FIG. 14
Limits of c.g.
AUTOROTATION
Autorotation is the condition of flight where the rotor is being driven by aerodynamic forces derived from an induced upwards airflow through the rotor as a result of the aircraft descending with no power applied to the rotor shaft.
It is the safety
factor in the event of engine failure and is similar to the ability of the fixed-wing aircraft to glide by maintaining a given airflow over the aerofoils. In the event of an engine failure, the rotor profile drag must be reduced as rapidly as possible, and the angle of attack must be adjusted in terms of the new relative airflow caused by the aircraft descending. Both these requirements are met by lowering the collective pitch lever to its low pitch stops immediately power is lost. Further alterations to the collective pitch will have to be made once autorotative rev/min have been established as the helicopter descends.
555/3/10
-29..
Autorotative Force if
The autorotative force is the component of total reaction acting forward in the plane of rotation,which opposes rotor a
profile drag. It depends on the direction of total reaction relative to the perpendicular to the plane of rotation. See Fig. l5. . \ \
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TOTAL REACTION
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vfi“ (a) normal flight FIG. 15
(b) autorotative flight Autorotative force
The angle of attack (the angle formed between the chord of the blade and the relative airflow) determines the direction of the total reaction.
On a rotor blade, it depends on
l.
The rate of descent of the helicopter,
2.
The forward speed of the helicopter,
3.
The rotational speed of the blade, and
H.
The collective pitch applied.
Figure l6 shows these four factors.
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555/3/10
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Plane of rotation
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Rotational speed '
FIG. 16
Forward speed
Angle of attack in autorotation
Three of these factors are common to the whole blade, but the fourth, the rotational speed, varies with the span, with a consequent variation in angle of attack. Thus, the autorotative characteristics will alter and must be summed up for the whole blade in order to arrive at its overall autorotative performance. There is usually an autorotative section of the blade corresponding to the area of highest L/D ratio, the force from which will balance the rotor profile drag from the remainder of the blade and so maintain constant rev/min. Figure 17 compares the forces acting at three different stations along the span of a blade.
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Comparison of forces acting at three different stations along the span of a blade _
555/3/1O
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An alternative method of deciding the autorotative section is by comparing the angles between lift and the perpendicular to the plane of rotation LA, and lift and total reaction LB. See Fig. 18. The effect on rotor rev/min is as follows.
e
l.
If LA is greater than LB, then rotor rev/min will increase.
2-
If LA is equal to LB, then rotor rev/min will be constant.
3.
If LA is less than LB, then rotor rev/min will decrease. Z CD -'lF7‘!
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[_A=lNFLOW ANGLE
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FIG. 18
Comparison of angles between lift, total reaction, and the perpendicu— lar to the plane of rotation
The blades are rigged to give normal rev/min under given conditions of flight, namely, a known all—up weight giving a certain rate of descent, a given forward speed, and collective pitch setting. Under these conditions, the autorotative section will correspond to that shown in Fig. l7 (b) and I9 (a). This is not the most efficient autorotative state, which would be achieved with the autorotative section closer to the tip of the
T
blade, but it provides a safety factor. In the event of external forces tending to slow the blades down, the angle of attack is increased, hence, the autorotative section moves out,
.
speeding up the blade to the original rev/min. Should the blades tend to speed up, angle of attack is decreased, autorotative efficiency decreases, and rev/min return to normal.
555/3/lO
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FIG. l9
Variation of autorotative rev/min with collective control lever
This increased efficiency can be achieved by increasing the collective pitch. In practice, the pilot can control the rotor performance with the collective control. However, this is limited by the danger of reaching the peak of autorotative efficiency beyond which rev/min will fall off-rapidly.
_
Minimum
permissible rev/min in autorotation are therefore governed by this factor, subject to possible coning angle limitation and an added safety factor.
-
Forward Speed As in powered flight, where the transition into forward flight produces translational lift and enables the power to be reduced, so in autorotation, increased mass flow through the blades improves performance, and results in a reduction in rate of descent. The graph of rate of descent in terms of forward speed compares very much with the power curve in level flight, the initial gain being offset by the sharp increase in parasite drag as speed increases.
See Fig. 20.
There is a small increase in rate of descent initially as a result of reducing the effective disc area to the relative airflow when tilting the disc forward. The optimum speeds for minimum rate of descent and maximum distance over the ground can then be found from the graph shown in Fig. 21. In the latter case,
555/3/10
-33..
Tan 9
Rate of descent velocity
=
u
Thus, when minimum, it gives a maximum ratio of speed to rate of descent, which is equivalent to distance covered from any given height.
RATE OF DESCENT
RATE BF
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PARASITE DRAG
EFFECT OF
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Effect of forward speed on rate of descent
ISPEED FOR IMAX. DIST
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FIG. 21
V 533;?“
The optimum speeds for minimum rate of descent and maximum distance
In theory, when using either of these speeds, maximum performance would be obtained by setting the lower figure of rotor rev/min, but in practice, the higher figure of rev/min is maintained at the lower speed. Additional speed can, to some extent, be converted into rev/min by the use of the flare. This relatively large change of disc attitude has already been discussed when dealing with transitions in powered flight and the basic effects remain the same:
l.
An increase in useful force as a result of increased angle of attack;
2.
Change in inflow (in the autorotative case, this is increased as a result of increasing the effective disc area);
3.
Reversal of thrust component;
H.
Increase in rotor rev/min.
555/3/IO
and
...3I.I_
The first two factors result in an increase in lift and therefore decrease the rate of descent to some extent. The amount by which rate of descent decreases varies considerably with aircraft type. The reversal of thrust causes a rapid decrease in forward speed just before touchdown. The increase in rev/min appear surprising in view of the increase in angle of attack and lift, which implies an increase in drag.
'
However, the
important factor is the relationship of total reaction to the i plane of rotation as shown in Fig. 22. Because total reaction moves forward relative to the perpendicular to the plane of rotation, the autorotative force is increased, causing a rise in rev/min. TO T4 .
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Blade Section in FIare (a) Blade section in autorotation FIG. 22
(b) Blade section in flare
Increase of rev/min in flare
AII—up weight The effect on autorotative performance of an increase in all-up weight is to increase the rate of descent, thus increasing the mass flow and so causing the rev/min to rise. This increase in rev/min must be controlled by increasing the collective pitch, which will restore the descent to a normal rate.
555/3/lO
'
_ 35 _ Altitude
The problem of autorotation at altitude is the reduced density. As in level flight, because the rev/min have to remain constant, lift will be reduced and therefore the rate of descent increased. The more important factor is the considerable increase in rev/min due to the decrease in drag, which must be controlled by use of the collective lever. If a helicopter base is sited at a high altitude, you may need to re-rig the collective control and to reset the collective pitch/low pitch stops to a higher pitch angle to get efficient flight at that altitude. If this is done, the low pitch stops mast be reset before a flight is made to a lower altitude to ensure that normal autorotative rev/min are available at that lower altitude.
NOTE:
If autorotation is being continued from high altitude to sea level, rev/min are controlled by a gradual lowering of the lever as density increases.
Safety Height Should the engine fail during hover, there will be a loss of height of approximately 300 feet before a full autorotative airflow can be established.
Allowing another lOO feet to make a safe landing, it is unsafe to be hovering below MOO feet, except that, up to about IO feet from the ground, a safe landing should be made simply by cushioning the impact by raising thecollective lever.
Forward speed will help to establish the inflow, so
this safety height can be reduced as speed is increased, until at approximately H5 knots, you should be able to make a safe landing from any altitude. Don't, however, fly too low at high speeds because, in the event of an engine failure, the aircraft might strike the ground before the speed could be reduced by flaring.
555/3/IO
_ 35 _
400'
/
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FIG. 23
.
45 K. B 5 K. FORWARD SPEED
‘Safety height and speed
RANGE AND ENDURANCE
The factors influencing range and endurance are similar to those for fixed-wing aircraft.
.1
Work
=
force X distance
Distance
=
work force
To obtain maximum distance, if work is constant, force must be minimal. The speed (V) to achieve this is obtained from the power curve, Fig. 24. Because POWEI
Tan 6
=
————~r-— velocity
Power
=
force X velocity
...
and
Combining (1) and (2), we get
555/3/1O
..
(2)
-37..
Tan 6
=
force X velocity velocity
=
Force
Thus, when tan 6 is a minimum (tangent to the curve), force will be minimum and we have the best speed to give the greatest range.
V; = Max endurome speed Va = Max range speed ed rzqur
Power ‘D <_-. N
_. _.- (:5
FIG. 24
5peed
A power curve (exaggerated)
To get the greatest endurance, the work available must be spread over the longest possible time.
. .
POWGI
=
. Time
=
work
. tlmé ‘
work —————— power
If the work is regarded as constant, power must be minimal to ensure maximum time. The forward speed to obtain this is immediately under the lowest point of the power curve. In Fig. 2H, VI shows the speed for greatest range, and Ve, the speed for maximum endurance.
555/3/1O
-33-
SUMMARY Ground resonance is a vibration caused by the interaction of an unbalanced articulated rotor head and its undercarriage. Blade sailing is controlled by flapping restrainers.
Autorotation is the helicopter’s equivalent of a fixed—wing aircraft's gliding. Autorotation rev/min increase with altitude.
PRACTICE EXERCISE C State whether each of the following statements is true or false. l.
Faulty dampers often give rise to blade sailing.
2.
Ground resonance can rapidly lead to the destruction
of the helicopter. 3.
The position of the centre of gravity will have no influence on the effectiveness of the cyclic pitch control.
4.
Autorotation is to the helicopter as spinning is to a fixed-wing aircraft.
5.
If the collective pitch control is raised when the engine power decreases, the rotor rev/min will increase.
6.
An increase in forward speed in autorotation results in a reduced rate of descent.
7.
Flaring results in a marked decrease in rotor rev/ min.
8.
Autorotation rev/min decrease with decreasing altitude.
9.
Oleo leg extensions have no effect on ground resonance
10.
The higher the AUW, the greater the rate of descent and the higher the autorotation rev/min.
(Answers on page 40)
555/3/10
f
_ 39 _
ANSWERS T0 PRACTICE EXERCISES
EXERCISE A Statements 2, 6, and 8 are True. 1.
False: In a hover in still air, there is no drag other than rotor profile drag because neither the aircraft nor the air surrounding it are moving.
3.
False: Greater efficiency is had from the rotor at translational speed due to the increased mass of air flowing through the rotor. This means that less power is needed, although parasite drag will have increased by a small amount.
H.
False: The increased mass of air flowing through the rotor means that induced power required will decrease as forward speed increases.
5.
False: To move forward, the helicopter needs some thrust. If this thrust is taken from the power available, that is just enough for hover, the helicopter will descend.
7.
False: When the air density decreases, the mass flow of air through the rotor disc decreases and so the lift also decreases. To maintain the lift, the collective pitch must be increased, which calls for more power.
9.
False: In all powered flight, the relative airflow enters the rotor from above the plane of rotation.
10.
False: Ground effect, which is another name for ground cushion, is taken to extend to a rotor height above ground of one half of the rotor diameter.
555/3/10
- ug _ EXERCISE B
Statements 3, 5, 8, and 10 are True. 1.
e
False: An advancing blade experiences an increased airflow. That is, it has V + Vfimwmrd flight
F
over its surfaces.
rotor
2.
False: Reversal of airflow occurs on the retreating half of the rotor disc.
H.
False: Stall of the retreating blade because of the cyclic pitch used to prevent flap-back places an upper limit to the forward speed.
6.
False: Because of the decreased air density, an increase in collective pitch will be needed to maintain lift. This will cause the stalling angle of the retreating blade to occur at a lower forward speed.
7.
False: Any change in the rotor disc attitude produces an immediate change in the useful force in both size and direction. This has a further effect on the disc, and so the helicopter is unstable in flight.
9.
False: A slight increase in collective pitch will affect all blades in exactly the same way. A forward movement of the cyclic control will correct flap-back.
EXERCISE C Statements 2, 6, 8, and 10 are True. l.
False: Faulty dampers cause the blades to move erratically about their drag hinges to give rotor imbalance. A low rotor rev/min in a gusting wind will cause blade sailing.
555/3/10
_ ul _
3.
False: If the c. g. is too far forward or too far aft, the cyclic control will run out of aft—and—forward movement sooner than normal. An incorrect c. g. position thus reduces the effectiveness of the cyclic control.
4.
False: Gliding is the fixed—wing equivalent of helicopter autorotation.
5.
False: Raising the collective control increases the pitch on all of the blades. If this is done with decreasing engine power, the rotor rev/min will quickly decrease.
7.
False: The increased angle of attack in the flare causes an increase in the autorotative force and this, in turn, increases the rotor rev/min.
9.
False: Incorrect oleo leg extensions will allow the fuselage to rock from side to side in unison with an out-of-balance rotor head. This rocking motion is ground resonance.
TEST PAPER 10 Make a sketch of a helicopter blade section in autorotation showing (a
The plane of rotation and a perpendicular to the plane of rotation at the trailing edge of the blade section,
(b)
The pitch angle,
(c
The inflow angle,
(d)
The angle of attack,
(e
The total reaction,
(f)
The lift vector,
555/3/10
_ ug -
(g)
The rotor profile drag vector, and
(h)
The autorotative force vector.
With the aid of a sketch, describe the airflow through
the rotors in a vortex ring state. Explain how the helicopter can enter and recover from this state and why it is a hazard to flight.
The power needed by a helicopter for horizontal flight can be considered in three parts. Name these parts, and state the use of each power vector.
What is flap-back, and how is it controlled?
What is translational lift, and why does it occur?
What are the effects of a c.g. position that is (a)
Outside of the forward limit, and
(b)
Outside of the aft limit.
(a)
What is airflow reversal?
(b)
When does it occur, and
(c)
What will it do to the helicopter when it becomes large?
What is ground resonance? Does it affect all types of helicopter? When is it most likely to occur, and what must be done when it does occur? What maintenance work can be done to lessen the possibility of ground resonance occurring?
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What factor has the most effect in limitin g th e f orward speed of the helicopter? What will happen if this facto r is ' ' ' ignored? List two other factors that also limit the forward speed.
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