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Key Facts 1 (Completed)

Stability is the tendency of an aircraft to return to a steady state of flight, after being disturbed by an external force, without any help from the pilot.

There are two broad categories of stability: static and dynamic.

An aircraft is in a state of equilibrium (trim) when the sum of all forces is zero and the sum of all moments is zero.

The type of static stability an aircraft possesses is defined by its initial tendency, following the removal of some disturbing force.

The three different types of static stability are:

•Positive static stability exists if an aircraft is disturbed from equilibrium and has the tendency to return to equilibrium.

•Neutral static stability exists if an aircraft is subject to a disturbance and has neither the tendency to return nor the tendency to continue in the displacement direction.

•Negative static stability exists if an aircraft has a tendency to continue in the direction of disturbance.

The longitudinal axis passes through the CG from nose to tail.

The normal axis passes “vertically” through the CG at 90° to the longitudinal axis.

The lateral axis is a line passing through the CG, parallel to a line passing through the wing tips. The three reference axes all pass through the centre of gravity.

Lateral stability involves motion about the longitudinal axis (rolling).

Longitudinal stability involves motion about the lateral axis (pitching).

Directional stability involves motion about the normal axis (yawing).

We consider the changes in magnitude of lift force due to changes in angle of attack, acting through a stationary point; the aerodynamic centre.

The aerodynamic centre (AC) is located at the 25% chord position.

The negative pitching moment about the AC remains constant at normal angles of attack. A wing on its own is statically unstable because the AC is in front of the CG.

An upward vertical gust will momentarily increase the angle of attack of the wing. The increased lift force magnitude acting through the AC will increase the positive pitching moment about the CG. This is an unstable pitching moment.

The tailplane is positioned to generate a stabilizing pitching moment about the aircraft CG.

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If the tail moment is greater than the wing moment, the sum of the moments will not be zero and the resultant nose-down moment will give an angular acceleration about the CG.

The greater the tail moment relative to the wing moment, the greater the rate of return towards the original equilibrium position.

The tail moment is increased by moving the aircraft CG forwards, which increases the tail arm and decreases the wing arm.

If the nose-down (negative) tail moment is greater than the nose-up (positive) wing moment, the aircraft will have static longitudinal stability.

The position of the CG when changes in the sum of the tail moment and wing moment due to a disturbance is zero, is known as the neutral point.

The further forward the CG, the greater the nose-down angular acceleration about the CG - the greater the degree of static longitudinal stability.

The distance the CG is forward of the neutral point will give a measure of the static longitudinal stability; this distance is called the static margin.

The greater the static margin, the greater the static longitudinal stability.

The aft CG limit will be positioned some distance forward of the neutral point.

The distance between the aft CG limit and the neutral point gives the required minimum static stability margin.

An aircraft is said to be trimmed if all moments in pitch, roll, and yaw are equal to zero.

Trim (equilibrium) is the function of the controls and may be accomplished by:

a)pilot effort,

b)trim tabs,

c)moving fuel between the wing tanks and an aft located trim tank, or

d)bias of a surface actuator (powered flying controls).

The term controllability refers to the ability of the aircraft to respond to control surface displacement and achieve the desired condition of flight.

A high degree of stability tends to reduce the controllability of the aircraft.

The stable tendency of an aircraft resists displacement from trim equally, whether by pilot effort on the controls (stick force) or gusts.

If the CG moves forward, static longitudinal stability increases and controllability decreases (stick force increases).

If the CG moves aft, static longitudinal stability decreases and controllability increases (stick force decreases).

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With the CG on the forward limit, static longitudinal stability is greatest, controllability is least and stick force is high.

With the CG on the aft limit, static longitudinal stability is least, controllability is greatest and stick force is low.

The aft CG limit is set to ensure a minimum degree of static longitudinal stability.

The fwd CG limit is set to ensure a minimum degree of controllability under the worst circumstance.

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Key Facts 2 (Completed)

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Positive static longitudinal stability is indicated by a negative slope of CM versus CL. The degree of static longitudinal stability is indicated by the slope of the curve.

The net pitching moment about the lateral axis is due to the contribution of each of the component surfaces acting in their appropriate flow fields.

In most cases, the contribution of the fuselage and nacelles is destabilizing.

Noticeable changes in static stability can occur at high CL (low speed) if:

a)the aeroplane has sweepback,

b)there is a large contribution of ‘power effect’, or

c)there are significant changes in downwash at the horizontal tail,

The horizontal tail usually provides the greatest stabilizing influence of all the components of the aeroplane. (page 259).

Downwash decreases static longitudinal stability.

If the thrust line is below the CG, a thrust increase will produce a positive or nose-up moment and the effect is destabilizing.

High lift devices tend to increase downwash at the tail and reduce the dynamic pressure at the tail, both of which are destabilizing.

An increase in TAS, for a given pitching velocity, decreases aerodynamic damping.

The aeroplane with positive manoeuvring stability should demonstrate a steady increase in stick force with increase in load factor or “g”.

The stick force gradient must not be excessively high or the aeroplane will be difficult and tiring to manoeuvre. Also, the stick force gradient must not be too low or the aeroplane may be overstressed inadvertently when light control forces exist.

When the aeroplane has high static stability, the manoeuvring stability will be high and a high stick force gradient will result. The forward CG limit could be set to prevent an excessively high manoeuvring stick force gradient. As the CG moves aft, the stick force gradient decreases with decreasing manoeuvring stability and the lower limit of stick force gradient may be reached.

At high altitudes, the high TAS reduces the change in tail angle of attack for a given pitching velocity and reduces the pitch damping. Thus, a decrease in manoeuvring stick force stability can be expected with increased altitude.

A flying control system may employ centring springs, down springs or bob weights to provide satisfactory control forces throughout the speed, CG and altitude range of an aircraft.

While static stability is concerned with the initial tendency of an aircraft to return to equilibrium, dynamic stability is defined by the resulting motion with time.

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An aircraft will demonstrate positive dynamic stability if the amplitude of motion decreases with time.

When natural aerodynamic damping cannot be obtained, artificial damping must be provided to give the necessary positive dynamic stability.

The longitudinal dynamic stability of an aeroplane generally consists of two basic modes of oscillation:

a)long period (phugoid)

b)short period

The phugoid oscillation occurs with nearly constant angle of attack.

The period of oscillation is so great, the pilot is easily able to counteract long period oscillation.

Short period oscillation involves significant changes in angle of attack.

Short period oscillation is not easily controlled by the pilot.

The problems of dynamic stability can become acute at high altitude because of reduced aerodynamic damping.

To overcome the directional instability in the fuselage it is possible to incorporate into the overall design dorsal or ventral fins.

The fin is the major source of directional stability for the aeroplane.

A T - tail makes the fin more effective by acting as an “end plate”.

Because the dorsal fin stalls at a very much higher angle of attack, it takes over the stabilizing role of the fin at large angles of sideslip.

Sweepback produces a directional stabilizing effect, which increases with increase in CL.

Ventral fins increase directional stability at high angles of attack. Landing clearance requirements may limit their size, require them to be retractable, or require two smaller ventral fins to be fitted instead of one large one.

Generally, good handling qualities are obtained with a relatively light, or weak positive, lateral stability.

The principal surface contributing to the lateral stability of an aeroplane is the wing. The effect of geometric dihedral is a powerful contribution to lateral stability.

A low wing position gives an unstable contribution to static lateral stability.

A high wing location gives a stable contribution to static lateral stability.

The magnitude of “dihedral effect” contributed by the vertical position of the wing is large and may require a noticeable dihedral angle for the low wing configuration. A high wing position, on the other hand, usually requires no geometric dihedral at all.

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The swept-back wing contributes a positive “dihedral effect”.

An aircraft with a swept-back wing requires less geometric dihedral than a straight wing.

The fin contribution to purely lateral static stability, is usually very small.

Excessive “dihedral effect” can lead to “Dutch roll,” difficult rudder coordination in rolling manoeuvres, or place extreme demands for lateral control power during crosswind take-off and landing.

Deploying partial span flaps gives a reduced dihedral effect.

A swept-back wing requires much less geometric dihedral than a straight wing. If a requirement also exists for the wing to be mounted on top of the fuselage, an additional “dihedral effect” is present. A high mounted and swept-back wing would give excessive “dihedral effect”, so anhedral is used to reduce “dihedral effect” to the required level.

When an aeroplane is placed in a sideslip, the lateral and directional response will be coupled, i.e. sideslip will simultaneously produce a rolling and a yawing moment.

Spiral divergence will exist when static directional stability is very large when compared to the “dihedral effect”.

The rate of divergence in the spiral motion is usually so gradual that the pilot can control the tendency without difficulty.

Dutch roll will occur when the “dihedral effect” is large when compared to static directional stability.

Aircraft which Dutch roll are fitted with a Yaw Damper. This automatically displaces the rudder proportional to the rate of yaw to damp-out the oscillations.

If the Yaw Damper fails in flight, it is recommended that the ailerons be used by the pilot to damp-out Dutch roll.

If the pilot uses the rudder, pilot induced oscillation (PIO) will result and the Dutch roll may very quickly become divergent, leading to loss of control.

When the swept wing aeroplane is at low CL the “dihedral effect” is small and the spiral tendency may be apparent. When the swept wing aeroplane is at high CL the “dihedral effect” is increased and the Dutch Roll oscillatory tendency is increased.

When pilot induced oscillation is encountered, the most effective solution is an immediate release of the controls. Any attempt to forcibly damp the oscillation simply continues the excitation and amplifies the oscillation.

Higher TAS common to high altitude flight reduces the angle of attack changes and reduces aerodynamic damping.

Mach Tuck is caused by loss of lift in front of the CG and reduced downwash at the tail due to the formation of a shock wave on a swept-back wing at high Mach numbers.

The Mach trim system will adjust longitudinal trim to maintain the required stick force gradient and operates only at high Mach numbers.

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