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VCLMAX VSR

VS1g VSW

5 kt or 5%

CAS

1, 2 & 3 on page 148

Figure 7.4 Aircraft without stick pusher

STICK
PUSH
			VSR				VSW
		2 kt				3 kt
		or				or
		2%				3%				CAS
VC
		LMAX

Figure 7.5 Aircraft with stick pusher

Artificial Stall Warning Devices

Adequate stall warning may be provided by the airflow separating comparatively early and giving aerodynamic buffet by shaking the wing and by buffeting the tailplane, perhaps transmitted up the elevator control run and shaking the control column, but this is not usually sufficient, so a device which simulates natural buffet is usually fitted to all aircraft.

Artificial stall warning on small aircraft is usually given by a buzzer or horn. The artificial stall warning device used on modern large aircraft is a stick shaker, in conjunction with lights and a noisemaker.

Stick Shaker

A stick shaker represents what it is replacing; it shakes the stick and is a tactile warning. If the stick shaker activates when the pilot’s hands are not on the controls, when the aircraft is on autopilot, for example, a very quiet stick shaker could not function as a stall warning so a noisemaker is added in parallel.

The stick shaker is a pair of simple electric motors, one clamped to each pilot’s control column, rotating an out of balance weight. When the motor runs, it shakes the stick.

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7 Stalling

An artificial stall warning device can receive its signal from a number of different types of detector switch, all activated by changes in angle of attack.

7
Stalling	FLAPPER SW ITCH
	( activated by movement of stagnation point )
	STAGNATION POINT
	( has moved downwards and backwards around leading edge )
	Figure 7.6 Flapper switch

Flapper Switch (Leading Edge StallWarningVane)

Figure 7.6. As angle of attack increases, the stagnation point moves downwards and backwards around the leading edge. The flapper switch is so located that, at the appropriate angle of attack, the stagnation point moves to its underside and the increased pressure lifts and closes the switch.

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AS ANGLE OF ATTACK INCREASES, VANE ROTATES RELATIVE TO FUSELAGE

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VANE

FUSELAGE

SKIN

Figure 7.7 Angle of attack vane

Angle of AttackVane

Figure 7.7. Mounted on the side of the fuselage, the vane streamlines with the relative airflow and the fuselage rotates around it. The stick shaker is activated at the appropriate angle of attack.

Angle of Attack Probe

Also mounted on the side of the fuselage, it consists of slots in a probe, which are sensitive to changes in angle of relative airflow.

All of these sense angle of attack and, therefore, automatically take care of changes in aircraft mass; the majority also compute the rate of change of angle of attack and give earlier warning in the case of faster rates of approach to the stall. The detectors are usually datum compensated for configuration changes and are always heated or anti-iced. There are usually sensors on both sides to counteract any sideslip effect.

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7 Stalling

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Basic Stall Requirements (EASA and FAR)

•It must be possible to produce and to correct roll and yaw by unreversed use of aileron and rudder controls, up to the time the aeroplane is stalled. No abnormal nose-up pitching may occur. The longitudinal control force must be positive up to and throughout the stall. In addition, it must be possible to promptly prevent stalling and to recover from a stall by normal use of the controls.

•For level wing stalls, the roll occurring between the stall and the completion of the recovery may not exceed approximately 20°.

•For turning flight stalls, the action of the aeroplane after the stall may not be so violent or extreme as to make it difficult, with normal piloting skill, to effect a prompt recovery and to regain control of the aeroplane. The maximum bank angle that occurs during the recovery may not exceed:

•Approximately 60 degrees in the original direction of the turn, or 30 degrees in the opposite direction, for deceleration rates up to 1 knot per second; and

•Approximately 90 degrees in the original direction of the turn, or 60 degrees in the opposite direction, for deceleration rates in excess of 1 knot per second.

Wing Design Characteristics

It has been shown that stalling is due to airflow separation, characterized by a loss of lift, and an increase in drag, that will cause the aircraft to lose height. This is generally true, but there are aspects of aircraft behaviour and handling at or near the stall which depend on the design of the wing aerofoil section and planform.

The Effect of Aerofoil Section

Shape of the aerofoil section will influence the manner in which it stalls. With some sections, stall occurs very suddenly and the drop in lift is very marked. With others, the approach to stall is more gradual, and the decrease in lift is less disastrous.

In general, an aeroplane should not stall too suddenly, and the pilot should have adequate warning, in terms of handling qualities, of the approach of a stall. This warning generally takes the form of buffeting and general lack of response to the controls. If a particular wing design stalls too suddenly, it will be necessary to provide some sort of artificial pre-stall warning device or even a stall prevention device.

A given aerofoil section will always stall at the same angle of attack

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Features of aerofoil section design which affect behaviour near the stall are:

•leading edge radius,

•thickness-chord ratio,

•camber, and particularly the amount of camber near the leading edge, and

•chordwise location of the points of maximum thickness and maximum camber.

Generally, the sharper the nose (small leading edge radius), the thinner the aerofoil section, or the further aft the position of maximum thickness and camber, the more sudden will be the stall. e.g. an aerofoil section designed for efficient operation at higher speeds, Figure 7.8.

The stall characteristics of the above listed aerofoil sections can be used to either encourage a stall to occur, or delay stalling, at a particular location on the wingspan.

CL	1.	ROUNDED LEADING EDGE
	2.	HIGHER THICKNESS-CHORD RATIO
	3.	MAX. THICKNESS AND CAMBER MORE FWD.

1. SHARP LEADING EDGE

2. LOW THICKNESS-CHORD RATIO

3. AFT MAXIMUM THICKNESS AND CAMBER

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Figure 7.8

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