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Where: f, =lowest natural frequency of the tab as installed in the
glider (c.p.m.)—either tab rotation about the hinge
line or tab torsion whichever is lower. Ce=Chord of movable control surface aft of the hinge line,
at the tab midspan position (ft.). S, =Span of tab (ft.) Sc =Span of movable control surface to which tab is at
tached (both sides of elevator, each aileron and rudder)
(ft.) Particular care should be taken in the detail design to minimize the possibility of fatigue failures which might allow the tab to become free and flutter violently.
Balance weight attachment criteria.–Balance weights should be distributed along the span of the control surface so that the static unbalance of each spanwise element is approximately uniform. However, where a single external concentrated balance weight is attached to a control surface of high torsional rigidity the natural frequency of the balance weight attachment should be at least 50 percent above the highest frequency of the fixed surface with which the control surface may couple in a flutter mode. For example the aileron balance weight frequency should be at least 50 percent above the wing fundamental torsional frequency. The balance weight supporting structure should be designed for a limit load of 24g normal to the plane of the surface and 12g in the other mutually perpendicular directions.
It should be noted that the dynamic balance coefficient K/I can be reduced by (1) reducing K, (2) increasing I or (3) reducing K and increasing I. Since an increase in I results in a reduced control surface natural frequency with possible adverse flutter effects, the primary purpose of ballast weights used to reduce K/1, should be to decrease the product of inertia K and not to increase the mass moment of inertia I.
Flight flutter testing.–Various procedures are available for demonstrating freedom from flutter, namely: Performance of a flutter analysis, application of the simplified criteria as discussed above, when applicable, and performance of a flight flutter test by the applicant. It is not recommended that flight flutter tests be used as a general procedure for substantiating freedom from flutter. The performance of a flutter analysis or the application of the simplified flutter criteria are considered preferred procedures due to the hazards involved in flight flutter tests. It is recommended that flight flutter tests be performed only when allied investigations or engineering evaluations give some assurance that the tests may be performed safely. A flight flutter test program may involve flutter substantiation of two or more control surfaces whose static unbalance values are in excess of the allowable values in the simplified flutter criteria. As a precautionary procedure, it is recommended that tests be performed for one surface at a time with the remaining unsubstantiated control surfaces balanced to at least the degree indicated necessary by the simplified flutter criteria.
(a) Acceptability.–Flight flutter tests will be acceptable as substantiation of freedom from flutter when it can be demonstrated by such tests that proper and adequate attempts to induce flutter have been made within the speed range up to Vp, and the vibratory response of the structure during the tests indicates freedom from flutter.
(b) Records.-Flight recording instrumentation of either the electrical or photographic type should be installed to provide a permanent record of the control surface and/or fixed surface response to the applied flutter exciting forces, as well as a record of the associated test airspeed.
(c) Test procedure.—The following is an outline of an acceptable procedure for demonstrating freedom from flutter by flight tests in which rapid control surface deflections are applied to induce flutter: (1) The tests should cover the flight speed range with excitation
applied at small incremental increases of airspeed up to Vo. The incremental speed increases between 0.8 Vo and Vo should not be more than 5 m.p.h. At lower flight speeds, larger
increments of airspeed may be used. (2) The controls should be deflected to attempt to excite flutter as
follows: • Aileron control to induce wing and aileron flutter. . Rudder control to induce rudder and vertical tail flutter. • Elevator control to induce elevator and horizontal tail flutter,
and symmetric wing flutter. . Control surface to which the tab is attached to induce tab
flutter. (3) Attempts to induce flutter should be made by abrupt rotational
deflections of the respective control surfaces. These deflections should be obtained by striking the corresponding control with the free hand, or foot, and the disturbed control should be allowed to stabilize without restraint by the pilot. The force applied should be sufficient to produce an impulsive deflection
of the control surface of at least 3 degrees. (4) At each test speed, at least 3 attempts to induce flutter should
be made for each of the surfaces being investigated.
(5) A permanent record* at each test speed should be obtained as
follows: • In flutter tests of the rudder, elevator and tab surfaces; a time
history of the control surface rotational deflection, and the
tory response, and the associated test airspeed. (6) The tests should consider significant variations in mass and
rigidity values which might be expected in service. The aileron and tab control systems should be freed to the extent necessary
to be representative of what might be expected in service. (7) The tests should be conducted at altitudes of approximately
50 to 75 percent of the maximum operating altitude. (8) The vibration survey should be conducted prior to performing
the flight flutter tests. The structural frequency and vibration mode data should be evaluated to determine what structural
modes are most likely to be flutter critical. (9) The resulting oscillations of the wing and/or control surfaces
in response to the excitation applied shall be damped with no tendencies to persist at any test speed.
* These data can be obtained by installing a control surface position indicating device at the control surface, a vibration pickup in the vicinity of one wing tip to detect the wing response, and a suitable pressure transducer connected to the aircraft's pitot-static system to measure the airspeed with the electrical signals from these instruments connected to a recording oscillograph.
Alternatively, photographic methods using cameras may also be used for recording the flight test data. However, the photographic records obtained should be of a nature that will permit satisfactory evaluation of the degree of control surface deflection applied, the flutter stability and a correlation of the airspeed record with the associated flutter test point.
Chapter 3—DESIGN, CONSTRUCTION AND
FABRICATION The primary structure should not incorporate design details which experience has shown to be unreliable or otherwise unsatisfactory. The suitability of all design details should be firmly established. Products such as bolts, pins, screws, tie-rods, wires, terminals, et cetera, used in the primary structure should be of aircraft standards as established by the SAE or as established by government aircraft standards and specifications or TSO's. Strength values contained in MIL-HDBK-5, ANC-18, and MIL-HDBK-17, and ANC-23 Part II, should be used unless shown to be inapplicable in a particular case. The primary structure should be made from materials which experience or conclusive tests have proved to be uniform in quality and strength and to be otherwise suitable for glider construction. The methods of fabrication employed in constructing the primary structure should be such as to produce a uniformly sound structure which should also be reliable with respect to maintenance of the original strength under reasonable service conditions. The workmanship of the primary structure should be of sufficiently high grade to insure proper continued functioning of all parts. All members of the structure should be suitably protected against deterioration or loss of strength in service due to weathering, corrosion, abrasions and other causes. Adequate provision for ventilation and drainage of all parts of the structure should be made. (See also FAA Technical Manual No. 103 "Aircraft Design Through Service Experience.")
PROCESSES Detailed information on assembly processes will be found in CAM 18, "Maintenance, Repair and Alteration of Airframes, Powerplants, Propellers and Appliances,” available from the Government Printing Office, Washington, D.C. Bulletin ANC-19, “Wood Aircraft Inspection and Fabrication" also includes detailed information on the properties and fabrication techniques of various woods and wood materials and associated data on aircraft glues.
WING DESIGN The more elementary types of gliders usually call for inexpensive wing construction, with simple parts which are easy to construct and assemble. Typical construction of this type is the two-spar wing having a drag truss composed of fore and aft drag struts and single drag and antidrag wires. Where double drag wires, that is, top and bottom drag wires, are employed, the torsional stiffness of the wing is greater than if single drag wires are employed.
If the general planform of the wing is rectangular, and if the spars are parallel, construction is relatively simple. For instance, in the layout of fittings, most of the angles would be right angles, and the problems of load determination would be two dimensional. In wings with rectangular planform, most of the ribs would be of identical construction. In contrast, if the general planform of the wing is tapered, then the construction becomes more complicated, as evidenced by the fact that the spars would necessarily taper in depth and each rib would be different due to the effect of taper. However, there are many advantages to the tapered wings. Usually, they are more efficient structurally and aerodynamically. The wing structures of most gliders of advanced design are currently comprised of a single spar with a D-nose and a special drag strut intersecting the main spar at about the 14-span point. The single spar with D-nose type of construction requires careful workmanship, especially in laminating the spar and in attaching the nose cover over the supporting noseribs.
The single spar D-nose type of wing construction lends itself to metal as well as wood construction. Metal has the disadvantages that special skill, tools and machines are required to work it, and the fact that the structure is usually not highly enough loaded to permit the most efficient use of the material. Also, metal working tools and machinery are more expensive than wood working tools and machinery. The minimum gauges of material usable from the standpoint of handling, workability, and corrosion resistance and, in some cases, the minimum gauges available commercially, often exceed that needed to carry the design loads. Consequently, the structure if of metal may be somewhat heavy. For a particular wing of relatively small thickness, but having a high aspect ratio and high design loading, it is probable that metal construction would be as light and efficient as wood.
Particular care must be used in the design of the wing-to-fuselage attachment fittings of single-spar wings. In some cases a secondary member is run back from the spar at an angle connecting into the rear root fitting at a chordal point corresponding approximately to the rear spar position in a two-spar wing. The nose covering is carried back to this member; to give structural stability, two pins at the root, one forward, and one aft are necessary for an externally braced wing. Full cantilever wings require horizontal pins at the top and bottom of the spar, or one vertical pin full depth at this point, and another pin aft to take out the drag and torsion reactions. Some wings have another pin at the leading edge to increase the rigidity, but this is not