OPERATING STRENGTH. LIMITATIONS

The weight of the structural components of an aircraft is an extremely important factor in the development of an efficient aircraft con­figuration. In no other field of mechanical design is there such necessary importance assigned to structural weight. The efficient aircraft and powerplant structure is the zenith of highly refined minimum weight design. In
order to obtain the required service life from his aircraft, the Naval Aviator must under­stand, appreciate, and observe the operating strength limitations. Failure to do so will incur excessive maintenance costs and a high incidence of failure during the service life of an aircraft.

GENERAL DEFINITIONS AND STRUC­TURAL REQUIREMENTS

There are strength requirements which are common to all aircraft. In general, these re­quirements can be separated into three particu­lar areas. These are detailed in the following discussion.

STATIC STRENGTH

The static strength requirement is the con­sideration given to the effect of simple static loads with none of the ramifications of the repetition or cyclic variation of loads. An important reference point in the static strength requirement is the “limit load’’ condition. When the aircraft is at the design configura­tion, there will be some maximum of load which would be anticipated from the mission requirement of the airplane. For example, a fighter or attack type aircraft, at the design Configuration, may encounter a very peak load factor of 7.5 in the accomplishment of its mis­sion. Of course, such an aircraft may be sub­ject to load factors of 3, 4, 5, 6, 1, etc., but no more than 7 5 should be required to accom­plish the mission. Thus, the limit load condi­tion is the maximum of loads anticipated in normal operation of the aircraft. Various types of aircraft will have different limit load factors according to the primary mission of the aircraft. Typical values are tabulated below:

Type of aircraft:

Fighter or attack…………………………………………….. 7 5

Trainer……………………………………………………………. 7.5

Transport, patrol, antisubmarine…………….. 3.0 or 2.5

Of course, these examples are quite general and it is important to note that there may be varia­tions according to specific mission require­ments.

Since the limit load is the maximum of the normally anticipated loads, the aircraft struc­ture must withstand this load with no ill effects. Specifically, the primary structure of the aircraft should experience no objectionable
permanent deformation when subjected to the limit load. In fact, the components must with­stand this load with a positive margin. This requirement implies that the aircraft should withstand successfully the limit load and then return to the original unstressed shape when the load is removed. Obviously, if the air­craft is subjected to some load which is in excess of the limit load, the overstress may incur an objectionable permanent deformation of the primary structure and require replace­ment of the damaged parts.

Many different flight and ground load condi­tions must be considered to define the most critical conditions for the structural com­ponents. In addition to positive lift flight, negative lift flight must be considered. Also, the effect of flap and landing gear configura­tion, gross weight, flight Mach number, sym­metry of loading, c. g. positions, etc., must be studied to account for all possible sources of critical loads. To verify the capability of the structure, ground static tests are conducted and flight demonstrations are required.

To provide for the rare instances of flight when a load greater than the limit is required to prevent a disaster, an “ultimate factor of safety” is provided. Experience has shown that an ultimate factor of safety of 1-5 is suf­ficient for piloted aircraft. Thus, the aircraft must be capable of withstanding a load which is 1.5 times the design limit load. The primary structure of the aircraft must withstand the “ultimate load” (1.5 times limit) without failure. Of course, permanent deformation may be expected with this “overstress” but no actual failure of the major load-carrying components should take place at ultimate load Ground static tests are necessary to verify this capability of the structure.

An appreciation of the static strength re­quirements may be obtained by inspection of the basic properties of a typical aircraft metal. Figure 5-1 illustrates the typical static strength properties of a metal sample by a plot of applied stress versus resulting strain. At low values

T

CYCLIC

STRESS

(PSD

of stress the plot of stress and strain is essen­tially a straight line, i. e., the material in this range is elastic. A stress applied in this range incurs no permanent deformation and the ma­terial returns to the original unstressed shape when the stress is released. At higher values of stress the plot of stress versus strain develops a distinct curvature in the strain direction and the material incurs disproportionate strains.

High levels of stress applied to the part and then released produce a permanent deforma­tion. Upon release of some high stress, the metal snaps back—but not all the way. The stress defining the limit of tolerable permanent strain is the “yield stress" and stresses applied above this point produce objectionable per­manent deformation. The very highest stress the material can withstand is the “ultimate stress." Noticeable permanent deformation usually occurs in this range, but the material does have the capability for withstanding one application of the ultimate stress.

The relationship between the stress-strain diagram and operating strength limits should be obvious. If the aircraft is subjected to a load greater than the limit, the yield stress may be exceeded and objectionable permanent deformation may result. If the aircraft is subject to a load greater than the ultimate, failure is imminent.