Page 257 - Aircraft Stuctures for Engineering Student

238 Airworthiness and airframe loads

of time. Such an instrument was developed by J. Taylor in 1950 and was designed so
that the response fell off rapidly above 10 Hz. Crossings of g thresholds from 0.2g to
1.8g at 0.lg intervals were recorded (note that steady level flight is 1g flight) during
experimental flying at the RAE on three different aircraft over 28 000 km, and the
best techniques for extracting information from the data established. Civil airlines
cooperated by carrying the instruments on their regular air services for a number
of years. Eight different types of aircraft were equipped so that by 1961 records
had been obtained for regions including Europe, the Atlantic, Africa, India and the
Far East, representing 19 000 hours and 8 million km of flying.
Atmospheric turbulence and the cabin pressurization cycle are only two of the
many fluctuating loads which cause fatigue damage in aircraft. On the ground the
wing is supported on the undercarriage and experiences tensile stresses in its upper
surfaces and compressive stresses in its lower surfaces. In flight these stresses are
reversed as aerodynamic lift supports the wing. Also, the impact of landing and
ground manoeuvring on imperfect surfaces cause stress fluctuations while, during
landing and take-off, flaps are lowered and raised, producing additional load cycles
in the flap support structure. Engine pylons are subjected to fatigue loading from
thrust variations in take-off and landing and also to inertia loads produced by lateral
gusts on the complete aircraft.
A more detailed investigation of fatigue and its associated problems is presented in
Section 8.7 after the consideration of basic manoeuvre and gust loads.

The maximum loads on the components of an aircraft’s structure generally occur
when the aircraft is undergoing some form of acceleration or deceleration, such as
in landings, take-offs and manoeuvres within the flight and gust envelopes. Thus,
before a structural component can be designed, the inertia loads corresponding to
these accelerations and decelerations must be calculated. For these purposes we
shall suppose that an aircraft is a rigid body and represent it by a rigid mass, 111,
as shown in Fig. 8.3. We shall also, at this stage, consider motion in the plane of
the mass which would correspond to pitching of the aircraft without roll or yaw.
We shall also suppose that the centre of gravity (CG) of the mass has coordinates
2, 3 referred to x and y axes having an arbitrary origin 0; the mass is rotating
about an axis through 0 perpendicular to the +XJ’ plane with a constant angular
velocity w.
The acceleration of any point, a distance r from 0, is w2r and is directed towards 0.
Thus, the inertia force acting on the element, bm, is w’rSm in a direction opposite to
the acceleration, as shown in Fig. 8.3. The components of this inertia force, parallel to
the x and y axes, are w2rSm cos 6 and w2rSn? sin 6 respectively, or, in terms of .Y and J’,
w2xSm and w2ySm. The resultant inertia forces, F, and F,., are then given by
S’
F, = w xdm =
s?
’J’
F,. = w ydm = wL ydm