Limitations

The Manoeuvre Envelope (V - n Diagram)

The maximum load factors which must be allowed for during manoeuvres are shown in an envelope of load factor against speed (EAS). Figure 14.1 shows a typical manoeuvre envelope or V - n diagram.

The limit load factors will depend on the design category of the aircraft.

The EASA regulations state that:

a)For normal category aircraft, the positive limit load factor may not be less than 2.5 and need not be more than 3.8.

(So that structural weight can be kept to an absolute minimum, a manufacturer will not design an aircraft to be any stronger than the minimum required by the regulations).

The positive limit load factor for modern high speed jet transport aircraft is 2.5.

b)For utility category aircraft the positive limit load factor is 4.4

c)For aerobatic category aircraft the positive limit load factor is 6.0

The negative limit load factor may not be less than:

d)-1.0 for normal category aircraft

e)-1.76 for utility category aircraft

f)-3.0 for aerobatic category aircraft

The CLMAX Boundary

The line OA in Figure 14.1 is determined by the CLMAX of the aircraft. In theory, the lift, and hence the load factor for a given weight, depends on the angle of attack of the wing and the

airspeed. The maximum possible lift will occur at the angle of attack where CL is a maximum. At this angle of attack the lift will increase with speed as shown by the line OA.

For level (1g) flight the speed at CLMAX will be the stalling speed (VS), represented by point S in

Figure 14.1.

At Point A, the load factor reaches its positive limit.

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14 Limitations

Limitations 14

It can be seen from Figure 14.2 that at speeds below point A the wing cannot produce a lift force equal to the limit load factor, whereas at speeds above point A the limit load factor can be exceeded. Manoeuvres at speeds above point A therefore have the potential to cause permanent deformation to the structure or structural failure if the ultimate load is exceeded.

This does not mean that any manoeuvre at a speed greater than point A will always cause structural damage; manoeuvres may be performed safely provided that the limit load factor is not exceeded.

PERMANENT DEFORMATION OF STRUCTURE POSSIBLE

	4	STRUCTURAL FAILURE		ULTIMATE
				LOAD FACTOR
			UNSAFE	FACTOR
	3			OF
		A	MANOEUVRE	SAFETY (1 5)
		10	POSITIVE LIMIT LOAD FACTOR
	2
		STALL
		ANGLE	SAFE
		5
	1		MANOEUVRE
	O			SPEED (EAS)

Figure 14.2 Loads imposed during manoeuvres

There is, of course, a safety factor on the airframe of 1.5 so complete failure of the structure will not occur at the load factor of 2.5 but at 2.5 × 1.5 = 3.75.

However, permanent deformation of the structure may occur at load factors between 2.5 and 3.75, so it is not safe to assume that the load factor may be increased above the limiting value just because there is a safety factor.

Design Manoeuvring Speed, VA

The highest speed at which sudden, full elevator deflection (nose-up) can be made without exceeding the design limit load factor.

When establishing VA the aeroplane is assumed to be flying in steady level flight, at point A1 in Figure 14.3, and the pitch control is suddenly moved to obtain extreme positive pitch

acceleration (nose-up). VA is slower than the speed at the intersection of the CLMAX line and the positive limit load factor line (point A) to safeguard the tail structure because of the higher load

on the tailplane during the pitch manoeuvre (Ref. Page 274, Chapter 10, Manoeuvre Stability).

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			Limitations	14
3	C L MAX WING FLAPS UP
		A
2
1		A1
		VC	VD
0	Vs 1	VA
-1

Figure 14.3 Design manoeuvring speed VA

Line OA in Figure 14.3 represents the variation of stalling speed with load factor. Stalling speed increases with the square root of the load factor, therefore;

VA = VS1g √ n

For example an aircraft with a 1g stalling speed of 60 kt and limit load factor of 2.5 would have a VA of:

60 √ 2.5 = 95 kt

Limitations 14

Effect of Altitude on VA

At high altitude the equivalent stalling speed increases with ‘g’ rather more rapidly than at sea level because of the Mach number effect on CLMAX. Above a certain altitude the buffet boundary may intersect the stall boundary at a value of ‘g’ lower than the structural limit, thus VA will become more limiting at high altitude (Please refer to Figure 13.25 for a diagram).

Effect of Aircraft Weight on VA

The 1g stalling speed depends on the weight of the aircraft. The line OA is drawn for the maximum design weight, so for lower weights the stalling speed will be less.

For the same limit load factor VA will therefore decrease. For the example considered above, if VA is 95 kt at 2500 lb weight, then at 2000 lb weight it will be:

2000
95 √2500	= 85 kt

Note: 20% decrease in weight has given approximately 10% decrease in VA.

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14 Limitations

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	3	POSITIVE	A	C
		CLMAX
	2			D
	1	S
				E	SPEED
	O	VS			(EAS)
	1	H		F
NEGATIVE	CLMAX		VA
				VC
				VD

Figure 14.4 The manoeuvre envelope

Design Cruising Speed VC

Point ‘C’ in Figure 14.4 is the design cruise speed VC. This is a speed selected by the designer and used to assess the strength requirements in the cruise. Its value is determined by the requirements CS-25.335 and CS-23.335. It must give adequate spacing from VB (see page 467) and VD to allow for speed upsets. For example CS-25 requires VC to be at least 43 kt above VB, and not greater than 0.8 VD. CS-23 has similar requirements. VC need not exceed the maximum speed in level flight at maximum continuous power (VH) or in CS-23, 0.9 VH at sea level

Design Dive Speed VD

Point ‘D’ in Figure 14.4 is the design dive speed VD. This is the maximum speed which has to be considered when assessing the strength of the aircraft. It is based on the principle of an upset occurring when the aircraft is flying at VC, resulting in a shallow dive, during which the speed increases, until recovery is effected. If the resulting speed is not suitable because of buffet or other high speed effects, a demonstrated speed may be used. This is called VDF, the flight demonstrated design dive speed.

Negative Load Factors

In normal flying and manoeuvres it is not likely that very large negative ‘g’ forces will be produced; however, some negative ‘g’ forces may occur during manoeuvres and the aircraft must be made strong enough to withstand them.

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