In fact, the useful limit Mach numbers in operation are the ones for which buffeting occurs. 

Figure F22 represents the buffet limit, and for n = 1 (level straight flight), a minimum Mach appears for low speed buffet and a maximum Mach for high speed buffet. When n increases, the Mach number range decreases, so that when n = n max, Mmin = Mmax. 

So, nmax is the maximum admissible load factor at this weight and altitude, and the corresponding Mach number M allows the highest margin regarding buffet limit. 

3.3.2.3. Pressure altitude effect

 Figure F23 illustrates the effects of pressure altitude on the lift area. It appears that, for a given weight: 

When nmax = 1, the aircraft has reached the lift ceiling. For example, in Figure F23, PA3  corresponds to the lift ceiling at a given weight. 

At pressure altitude PA1 (Figure F23), nmax = 1.3. That is to say, it is possible to bear a load factor equal to 1.3, or make a 40° bank turn before buffeting occurs. 

In order to maintain a minimum margin against buffeting and ensure good aircraft maneuverability, it is necessary to determine an acceptable load factor limit below which buffeting shall never occur. This load factor limit is generally fixed to 

1.3. This value is an operating limitation, but not a regulatory one. The corresponding altitude is called the “1.3g buffet limited altitude” or “buffet ceiling”. 

For a given Mach number, Figure F24 represents the 1.3g buffet limited altitude versus weight. At a given Mach number, when weight ì. the buffet limited altitude ê. 

As a result, the maximum recommended altitude indicated by the FMGS, depending on aircraft weight and temperature conditions, is the lowest of the: 

Maximum certified altitude, 

Maximum cruise altitude, 

1.3g buffet limited altitude, 

Climb ceiling (see the “Climb” chapter). 

3.3.2.4. A320 example

 Figure F25 shows how buffet limitations are illustrated in an A320 FCOM. 

 In practice, for a given weight, the load factor limitation (1.3g) is taken into account as follows: 

At a fixed FL, the cruise Mach number range is determined for n = 1.3g, 

At a fixed cruise Mach number, the maximum FL (buffet ceiling) is determined for n = 1.3g. 

3.4. Cruise Optimization: Step Climb 

Ideal cruise should coincide with optimum altitude. As a general rule, this altitude is not constant, but increases as weight decreases during cruise. On the other hand, ATC restrictions require level flight cruise. Aircraft must fly by segments of constant altitude which must be as close as possible to the optimum altitude. 

In accordance with the separation of aircraft between flight levels, the level segments are established at ± 2,000 feet from the optimum altitude. In general, it is observed that in such conditions: 

SR ≥ 99% SR max 

As a result, the following profile is obtained for a step climb cruise (Figure F26). 

Flight levels are selected in accordance with temperature conditions. Usually, the first step is such that it starts at the first usable flight level, compatible with maximum cruise altitude. This is the case with the ISA condition cruise example in Figure F26. 

4. FCOM CRUISE TABLE 

In the FCOM, cruise tables are established for several Mach numbers in different ISA conditions with normal air conditioning and anti-icing off. Aircraft performance levels are presented in Figure F27. 

G. CLIMB 

1. FLIGHT MECHANICS 

1.1. Definitions 

The following Figure (G1) shows the different forces applied on an aircraft in 

climb. 

The angle of attack (α) represents the angle between the aircraft axis and the aerodynamic axis (speed vector axis tangent to the flight path). .

The climb gradient (γ) represents the angle between the horizontal axis and the aerodynamic axis. 

The aircraft attitude (θ) represents the angle between the aircraft axis and the horizontal axis (in a ground reference system). 

The rate of climb (RC) represents the vertical component of the aircraft’s speed. It is positive and expressed in feet per minute. 

1.2. Climb Equations 

During climb at constant speed, the balance of forces is reached. Along the aerodynamic axis, this balance can be expressed as : 

(1) Thrust cosα= Drag + Weight sinγ 

1 In order to simplify, the thrust vector is represented parallel to the aircraft longitudinal axis.