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the test load on the towing hook. Care should be taken in testing for these conditions to guard against overloading such portions of the fuselage. For instance, if the side load is resisted only by the front strut fitting and the tail post, the loading in the rear part of the fuselage might be higher than the design load and failure would occur. The solution in this case would be to apply a moment at the wing-root fittings and at the strut points as well as at the tail, each of which would be less than the ultimate loads for which the fuselage is designed.

In actual flight, the loads on the tow line, especially side loads, are resisted by inertia loads as well as by air loads. Accordingly, this should be given careful consideration in determining the magnitude and location of the fuselage test reactions. For example, much of the side load will be resisted by the inertia of the wing, through the wing-root fittings, while the vertical components of the towing loads will be resisted mainly by the inertia of the various items of mass in the fuselage, the loads being applied through their points of attachment to the fuselage.

Reference is made in Chap. 1 to the resultant effects on the strength of the wing root fittings and adjacent structure from unsymmetrical loads.

The comments of the preceding section regarding testing for towing loads apply also to landing gear static tests. Care should be exercised so that no part of the fuselage is overloaded locally at points where high loads would not normally

be expected during actual landings. 5. Deflections.—Deflection readings should be taken at each

increment of load, at various strategic locations, including the wing attachment fittings, the rear end of the fuselage

and, for the torsion test, the tip of the fin. C. Test report.-In all cases the manufacturer making a test is

required to submit a complete report covering details of tests. The report should include photographs or drawings of the test setup and the test specimen; photographs of failed parts or sections; records of deflections and readings taken; date of test; identification number of report; serial and model number of glider and signature of responsible witnesses and/or test personnel. In addition, the following points should be covered when applicable: (a) Substantiation by references or computations of the selec

tion of critical test conditions and loadings. (b) Loading schedule used in the test. (c) Description of test setup, with reference to drawing numbers.

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Figure 2-XII. Setup for test of att portion of a fuselage for sidewise acting fin and rudder loads.

FLIGHT LOAD TESTS In general, the demonstration of strength by means of flight load tests is not recommended. Any proposals for such tests should be brought to the attention of the regional office concerned for special study and recommendations.

FLUTTER AND VIBRATION Flutter and vibration prevention tests.-Wings, tail, control surfaces, tabs, wing spoilers, and dive brakes should be demonstrated to be free from flutter at all airspeeds and altitudes within the design speed range up to Vd. Freedom from flutter may be demonstrated either by a formal flutter analysis, by compliance with acceptable simplified flutter prevention criteria or by an acceptable flight flutter test.

Ground vibration tests.—The natural frequencies of all main structural components should be determined by ground vibration tests. These tests should determine at least the following resonant modes of vibration: Wing first symmetrical and unsymmetrical bending and wing torsion, symmetrical rotation of ailerons about their hinge lines, wing spoiler rotation about its hinge line, dive brake rotation about its hinge line, fuselage torsion, fuselage side bending, fuselage vertical bending, fin bending, stabilizer bending, rotation of tabs about their hinge lines.

Simplified flutter criteria.-Studies have indicated that for a conventional airfoil in which the center of gravity of the airfoil section is not too far back, wing flutter could be prevented by designing for a certain degree of wing torsional rigidity and by control surface dynamic balance, whereas empennage flutter could be prevented by providing a degree of control surface dynamic balance. Satisfactory rational analytic methods have been available for a number of years which would permit an engineer to carry through computations to determine the flutter stability of a specific design. In view of the fact that flutter is an aeroelastic phenomenon which is caused by a combination of aerodynamic, inertia and elastic effects, any criteria which does not consider all three effects is bound to have severe limitations. That this is so is evidenced by the fact that in almost all cases where rational analyses have been carried through for specific designs it has been found that the balance requirements specified by the simple criteria have been too severe. Although a rational flutter analysis is to be preferred to the use of the simplified criteria contained herein (since in most cases a better design may be achieved by reducing or eliminating the need for nonstructural balance weights), the application of these criteria to conventional aircraft is believed to be adequate to insure freedom from flutter.

WING FLUTTER CRITERIA The following wing flutter criteria should be applicable to all conventional gliders which do not have large mass concentrations on the wings.

Wing torsional stiffness.—The wing torsional flexibility factor
F defined below should be equal to or less than 200.

Where: F = SOC?ds

01 = Wing twist at station i, per unit torsional moment

applied at a wing station outboard of the end of

the aileron. (radians/ft.-lb.)
Ci = Wing chord length at station i, (ft.)
ds = Increment of span (ft.)

Vo=Design dive speed mph (IAS) Integration to extend over the aileron span only. The value of the above integral can be obtained either by dividing the wing into a finite number of spanwise increments AS over the aileron span and summing the values of 0,CAS or by plotting the variation of O_C12 over the aileron span and determining the area under the resulting curve.

In order to determine the wing flexibility factor F, a pure torsional couple should be applied near the wing tip (outboard of the end of the aileron span) and the resulting angular deflection at selected intervals along the span measured. The test can best be performed by applying simultaneously equal and opposite torques on each side of the glider and measuring the torsional deflection with respect to the glider center line. The twist in radians per unit torsional moment in ft.-lbs. should then be determined. If the aileron portion of the wing is divided into four spanwise elements and the deflection determined at the midpoint of each element the flexibility factor F can be determined by completing a table similar to table 2-I. Fig. 2-XV illustrates a typical setup for the determination of the parameters C and AS.

Aileron balance criterion.—The dynamic balance coefficient

K/I should not be greater than the value obtained from fig.
2–XVI wherein K/I is referred to the wing fundamental
bending node line and the aileron hinge line. If no knowledge
exists of the location of the bending node line the axis parallel
to the fuselage center line at the juncture of the wing and
fuselage center line at the juncture of the wing and fuselage
can be used.
Wherein: K =product of inertia

I =mass moment of inertia of aileron about its

hinge line

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