• Generally, the class of airplanes considered under § 23.629(f)(2), have three or more control surface hinges.
• Tab hinges are often the piano hinge type, which are inherently redundant.
• A review of the regulatory history including the preamble, NPRM, and the 1974- 1975 Airworthiness Review discussions, does not show anything addressing whether the hinges should be included as part of the flight control or tab control system in evaluating § 23.629(f). The implication is that it was not the intent of the rule to include the hinges and supporting structure.
(2) It is, therefore, concluded that the control surface hinges and tab hinges, their attachments, and local portions of structure should not be included as part of the flight control system or tab control system in determining compliance with the flutter requirements of § 23.629(f).
The control surface hinges, tab hinges, their attachments, and local portions of structure are not considered part of the control “system”, regardless of whether the system is fail-safe or not.
(3) In conjunction with this policy, special emphasis should be given to ensure that the rules applying specifically to hinges (§§ 23.393, 23.415, 23.623, 23.625, 23.657, 23.659) are fully evaluated and documented. In addition, hinges should also be evaluated for good design practices with respect to number of hinges, redundancy of fasteners, adequacy of structure adjacent to hinges, methods of providing hinge pin retention if the nut backs off, etc. Similarly, control surface or tab balance weights, their attachments, and local portions of the supporting structure are part of the control surface. They should not be included as part of the primary flight control system or tab control system in determining compliance with the flutter requirements of § 23.629(f).
12 CHAPTER 4. DIVERGENCE AND CONTROL REVERSAL
11. GENERAL. Static aeroelastic instabilities of lifting surfaces are avoided by providing adequate torsional rigidity. Methods to determine the adequacy of torsional rigidity are outlined in references 2 and 3 of Appendix 3.
12. DIVERGENCE. Divergence occurs when the aerodynamic torque exceeds the torque resisting capability of the wing. Because the aerodynamic torque is a function of speed as well as deflection, whereas the resisting torque depends on the torsional rigidity of the lifting surface which is a constant, there exists a limiting divergence speed. Divergence may occur with no warning.
13. CONTROL REVERSAL. Control reversal will often be preceded by pilot comments of “heavy” or “sluggish” ailerons. A limiting reversal speed is reached when the change in lift due to control surface rotation is nullified by the change in lift due to twist of the lifting surface.
A1-1 APPENDIX 1. GROUND TESTING
1. INTRODUCTION. The adequacy of the methods used to show compliance with § 23.629, as discussed in the main body of this document, is dependent upon the availability of reliable ground test data to verify the analytical data used and/or to serve as a basis for flutter substantiation per the simplified criteria of reference 1. This appendix, therefore, presents guidelines in conducting the more significant tests required to accomplish this objective. However, in keeping with the general purpose of this advisory circular, the information provided is not intended to be mandatory, nor is it to be considered an exhaustive treatment of the subject.
2. CONTROL SURFACE AND TAB MASS PROPERTIES. The experimental mass properties of control surfaces and tabs (weight, static moments, moments of inertia, and
c.g.) are important ingredients in flutter substantiation. These properties form a basis for verification of the analytical data used in the rational analysis and provide the necessary parameters for use in the simplified criteria. Reference 1 presents a detailed procedure for the experimental determination of these properties.
3. TAB FREE PLAY. Free play tests provide the necessary data for determining the effectiveness of a tab in fulfilling the requirements for irreversibility as specified in the main body of this document. In addition to demonstrating the maximum free play available, these tests provide the stiffness of the actuating system for use in computing tab rotational frequency.
a. Free play and stiffness may best be measured by a simple static test wherein “upward” and “downward” (or “leftward” and ”rightward”) point forces are applied near the trailing edge of the tab at the spanwise attachment of the actuator (so as not to twist the tab). The control surface should be blocked to its main surface. Rotational deflection readings are then taken near the tab trailing edge using an appropriate measuring device, such as a dial gauge. Several stepwise load and deflection readings should be taken using loads first applied in one direction, then in the opposite.
b. A plot of these load deflections typically appears as shown in Figure A-1.
Figure A-1. Load Deflections Plot
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c. Free play is then defined by extending th e best straight lines through zero. System stiffness may then be obtained from the slopes of the curves away from the zero point.
4. INFLUENCE COEFFICIENT TESTS. Bending and/or torsion influence coefficient test results form the basis for the definition of component stiffness distributions. The extent of the tests depends on the intended use of the data. A full scale test program, wherein the coefficients of each spanwise mass strip are defined, may be desired if experimental data is the primary source for defining component stiffness. In contrast, calculated influence coefficients, based on analytical bending (EI) and torsion (GJ) stiffness distributions, may be adjusted reliably with considerably less test data. A method is outlined below for determining influence coefficients for conventional structure, i.e., aspect ratio greater than four and unswept elastic axis.
a. The test article, wing, tailplane, or fin, is generally mounted at its root, without control surfaces, in a rigid test fixture for these tests. However, wing stiffness tests, particularly torsion as required for simplified criteria, may be successfully conducted with the wing mounted on the fuselage restrained in a cradle. This type of setup requires duplicate loading fixtures for right and left wing to balance the aircraft under load and thus minimize “jig rotation” effects.
b. The chordwise location of the elastic axis is determined by applying a torque load at selected stations and plotting the deflection versus chord shear center or elastic axis at that station.
c. Torsional influence coefficients (radians twist about the elastic axis per unit torque load) are obtained by applying a pure torque load about the elastic axis at the tip and measuring the resulting spanwise twist. The twist per unit torque applied at intermediate inboard stations will be the same inboard of the load point. Thus, it is necessary to load only one additional inboard station, say 75 percent span, to check for data repeatability only. To insure that the load applied is a pure torque load, the deflections of the elastic axis should be monitored during the loading process. Zero deflections should result.
d. Bending influence coefficients (deflecti ons per unit shear load) are obtained by applying shear load on the elastic axis at a selected station and measuring the resulting deflections at a sufficient number of spanwise locations to define the influence line for that load point. The procedure is repeated for each load station. To insure that the shear load is applied on the elastic axis, no appreciable chordwise variation in the measured deflections should be evident.
e. The experimental determination of fuselage stiffness properties can be accomplished essentially the same way as for the aerodynamic surfaces. In this test the fuselage is treated as two beams, forward and aft fuselage, each cantilevered from the wing-root attachment. It is extremely important that the fixture at this attachment be very rigid; and, any displacement of the test jig during loading must be monitored, regardless of how small, throughout the test for inclusion in the data analysis. Small displacements can be quite influential in a rather complex data reduction procedure, and if improperly done, can lead to erroneous and troublesome conclusions. On this basis it is often the practice to compute fuselage stiffness properties for the fuselage, then use ground vibration test results
A1-3 to tune calculated modes and, in turn, stiffness as required.
f. Thin-skinned structure may buckle at a ve ry low load, reducing actual stiffness in flight considerably from that determined by the above procedure and the analyst is cautioned to investigate such conditions.
5. GROUND VIBRATION TESTS. Ground vibration testing has as its fundamental objective the definition of vibration mode frequencies, mode shapes, and damping characteristics of an aircraft. These data then become the basis for the analytical development of a mathematical vibration model of the airplane or serve as a check on such a model once it is developed. The results ultimately become the basis for rational flutter analyses. If the simplified flutter prevention criteria of reference 1, discussed in the main body of this advisory circular, is used, then the results from these tests are used directly to establish a predicted flutter speed of the airplane.
The degree of sophistication required to conduct a resonance test (techniques, recording equipment, suspension system, etc.) depends upon the complexity of the structure being tested. Since it is impossible to cover all test situations that may arise, the discussions presented in this section are fundamental in nature, dealing specifically with sinusodial methods of excitations. They are intended as guidelines for those persons concerned with general type aircraft, which have only the basic test facilities. Other procedures employing random or impulse excitations are being used more frequently. However, these methods are considered beyond the scope of this advisory circular.
a. Test Article and Suspension System. The airplane should be supported in a level attitude such that the rigid body frequencies of the airplane on its support are less than onehalf the frequencies of the lowest elastic wing or fuselage mode to be excited.
One of the following methods of support can generally be used:
(1) Support the airplane on its landing gear with the tires deflated sufficiently to achieve the above result. Fifty percent normal tire pressure usually achieves good results.
It may be necessary to block the landing gear struts to eliminate damping in the oleos.
(2) Suspend the airplane on springs.
(3) Support the airplane on its landing gear resting on spring platforms.
(4) Support the airplane fuselage and wi ngs on large air-filled flotation bags.
(a) The airplane should be equipped wi th all items having appreciable mass such as engines and tip tanks. The weight and c.g. of the test article should be determined to enable proper correlation with the analytical model. Where fuel is located in the outboard 50 percent of the wing semispan, it may be desirable to test a full fuel condition in addition to the empty condition in order to provide additional data for math model correlation.
(b) It is generally advantageous to lock the control surfaces in their neutral position when obtaining airframe modes.
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b. Equipment. Various types of shakers are available, i.e., inertia, elastic, airjet, electromagnetic, etc. Electromagnetic exciters are generally preferred and most commonly used. This type consists of a coil that is attached to the structure with a fixed drive rod, as opposed to a flexible shaft or spring for inertia or elastic type shakers. The coil is surrounded by a magnetic field and is set in motion by an alternating current. Electronic oscillators and amplifiers are used to control this type of system.
(1) Vibration amplitude may be obtained by using either velocity pickups or accelerometers so long as transducer mass is insignificant. Most modern accelerometers weigh only a few ounces. Typically that is insignificant when installed on the fuselage, wings, or vertical and horizontal stabilizers. When used on control surfaces, this same accelerometer may affect the flutter characteristics. Care should be exercised to select a small, light-weight accelerometer when mounting on control surfaces like tabs. The output can be observed using a cathode ray oscilloscope and digital voltmeter. Phase relationship between two transducers can be noted with sufficient accuracy, and by exercising extra care, using an oscilloscope equipped with a grid screen.
(2) Data systems are available that provide the coincident, in- phase or real term, and the quadrature, the imaginary term, responses of the total response frequency (the product of the force and reference signal). Graphical representation of these terms is presented, providing a very accurate identification technique for resonant frequencies and phase relationships. Structural damping is also readily available from these data.
(3) Whatever data system is use d, uniformity is recommended. Piecemeal systems, using velocity pickups and accelerometers, or filters with different characteristics, etc., can give erroneous data and should not be used without careful regard to their calibrations and performance characteristics and limitations.
c. General Procedures for Airframe Modes. It is usually sufficient to apply a harmonic excitation force to the structure provided the force is not applied in the proximity of a node line. For this reason vibrators are usually attached at an extremity such as the nose and/or rear of the fuselage or near the tips of the wing or empennage surfaces where nodes are not likely to occur.
(1) The node line lies along th e elastic axis of the structure (fuselage, wing, stabilizer, etc.). The axis passes through the loci of the shear centers of the cross-sections along the span of the structure. This theoretical location is typically determined mathematically when the dynamic computer model is created, and may be tweaked during the ground vibration tests. It can also be determined during the vibration tests, since an excitation force that is applied in the proximity of a node line will produce a near zero structural response. The point of application of the excitation force can then be moved away from that point and tried again.
(2) With the shaker(s) and a reference pickup mounted at a selected location, frequency is varied upward through the range usually encountered in aircraft structures.
With small increments of frequency, the response of the structure is recorded and the resulting plot of amplitude of response versus forcing frequency is used to determine the resonant frequencies of the system. A typical sweep is shown in Figure A-2.
A1-5
Figure A-2. Typical Sweep
(3) Although duplication of peak responses will result, it is advantageous to obtain frequency response records with a reference pickup positioned on each of the main surfaces and fuselage at a specific shaker location. This will reduce the chances of overlooking modes.
(4) There are several criteria for establishing that the excited response approximates a normal mode of vibration. The most commonly accepted approach requires that all of the criteria below be met:
(a) A relative maximum response per unit input exists. A normal mode of vibration occurs at a resonant frequency. At resonant frequency, the response is at its maximum. One needs to “tune” the excitation frequency very carefully until the maximum response is reached. The frequency at which the maximum response is observed, gives the resonant frequency. If the excitation frequency is slightly off of that frequency, the response will fall well below the maximum response.
(b) Accelerations at all points in the structure are either exactly in phase or 180 degrees out of phase with each other. The accelerations measured at all points on the structure during resonance will be either in phase or out of phase with a reference location but will be at a ± 90 degree phase angle with the force, for small values of damping.
(c) A decay record exhibits a single-frequency, non-beating, low-damped characteristic.
(5) Having established the resonant frequencies, a survey of the aircraft is conducted with the shakers tuned to each frequency in-turn. A roving transducer is used to sense amplitude and phase angle relative to the reference pickup at each airplane location.
An adequate number of points should be surveyed along the span and chord (typically on the spars) of each surface and along the fuselage to define the airplane modal displacements, and the associated node lines. The number of survey points depends on the nature and span of the surface and the number of modes to be surveyed. If one is interested in obtaining a number of modes, (like third bending or fourth bending modes) more accelerometers are needed span-wise. Usually 8 to 10 accelerometers are sufficient for a small airplane. To obtain proper phase relationship additional excitation may be necessary.
A1-6
(6) It may not be necessary to survey identical peak frequency responses although they occur at different locations. In all probability, the mode will be the same. This can be determined by checking only a few stations or simply by visually observing the motion of the aircraft.
(7) Care should be exercised in defi ning component node lines for each mode.
This is particularly important in evaluating the effectiveness of balance weight locations.
d. Aircraft Structural Modes Usually Encountered. The modes excited during ground vibration depend on the type of configuration being tested. The vibration modes of an airplane that carries heavy mass on the wing, such as engines, tip tanks, etc., or has the stabilizer located high on the fin will be highly coupled and generally cannot be described except by diagrams that show the relative shape and phase of each part of the airplane.
Airplanes that do not have these design characteristics usually have relatively uncoupled modes, which can be described by naming the type of motion that is predominant. In general, the following predominant modes should be obtained insofar as is practicable.
(1) Wing Group Modes.
(a) For wings without engines, tip ta nks, or heavy external or internal stores: Wing vertical bending and wing torsion, fundamental and higher modes, symmetric and anti-symmetric.
(b) For wings carrying heavy masses outboard of the fuselage: Wing bending coupled with wing torsion and flexible store (engines) modes, fundamental and higher modes, symmetric and anti-symmetric.
(2) Fuselage - Empennage Group Nodes.
(a) Fuselage Torsion (coupled with stabilizer anti-symmetric bending).
(b) Fuselage lateral bending and fi n bending, fundamental and higher order consisting of two fundamental modes in which the fin tip and aft fuselage are in phase in one mode, and out of phase in the other.
(c) Fin bending - symmetric and anti-symmetric for multi-tail airplanes.
(c) Fin torsion (generally highly coupled with stabilizer yawing if stabilizer is located at the outer span stations of the fin).
(e) Rudder bending and torsion.
(f) Fuselage vertical bending and stabilizer bending, fundamental and higher order consisting of two fundamental modes in which the aft fuselage and stabilizer tips are in phase in one mode, and out of phase in the other.
A1-7
(g) Stabilizer torsion - symmetric and anti-symmetric.
(h) Stabilizer yawing for surface located at the outer span stations of the fin.
(i) All movable horizontal tail - ro tation coupled with bending, torsion.
(3) Engine or External Store Modes. For multiengine aircraft or aircraft carrying “large” pylon-mounted stores (the weight and the cg location with respect to the elastic axis of lifting surface are the significant factors in this instance), the pitch, roll, yaw, and lateral and vertical translation modes should be defined. These modes should also be determined for all turbo propeller engine installations. It may be necessary to excite the engine fore and aft on the propeller blade to obtain the most critical pitch and yaw modes. If this method is used, consideration should be given to possible modal distortion due to propeller blade flexibility. Also, caution should be exercised and the engine manufacturer's instructions followed concerning possible damage to bearings when exciting the engine.
e. General Procedures for Control System Modes. The experimental determination of control surface and tab rotation modes about their hinge lines may be difficult due to inherent friction within the system or the masking of these modes by structural interaction.
(1) On this basis, extra care is required for proper identification of the system's characteristics.
(2) For conventional aileron or elevator systems, the rotation modes may be successfully measured by applying a single excitation force to either the right hand or left hand surface. However, multiple shakers are preferred, particularly if the right and left surfaces are operated from separate control systems. Likely shaker positions are on the trailing edge at midspan or on the horn leading edge. Tab rotation may be determined from the control surface excitation but usually a direct excitation on the tab surface is required with the control surface (aileron, elevator, or rudder) blocked to its main surface.
(3) A transducer placed on the control being excited is used to monitor the response and determine peak frequencies by the same technique described for airframe modes. To define the modes excited, it is generally necessary to follow any or all of the following procedures:
(a) Monitor the phase between the right and left surface, the control column, or the attaching structure.
(b) Conduct a detailed survey of the surface, spanwise and chordwise, to define any structural modes. If the surface has a very long span or wide chord, these modes, bending and torsion, are likely to be dominant.
(c) Visually monitor the surface under excitation.
(d) Simple rationalization to distinguis h the excited modes from previously defined airframe modes.
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(4) In the performance of these tests, the shakers and/or transducers may contribute sufficient weight to the surface being tested to significantly affect the frequency of the surface. This is particularly true for tabs with very small mass and rotational inertias. Dunkerly's equation, presented in Appendix 3, reference 8, provides an acceptable method for correcting the measured frequency to the true surface frequency.
(5) A check on experimentally determined modes may be facilitated by calculating rotational frequencies from measured inertias and system stiffness properties obtained from static tests.
(6) For extremely light-weight structures , another method that may be used to eliminate the shaker influence is to use an air shaker or other device which does not directly attach to the control surface or tab.
f. Control Surface Rotation. Symmetric aileron rotation, the normal opposed operational mode, with control stick fixed or free, is defined as the peak frequency at which both ailerons are rotating in phase. Anti-symmetric rotation, the normal operation mode, generally has zero or very low stiffness.
(1) Rudder rotation in the no rmal operation mode with pedals free occurs when the rudder and pedals are out of phase.
(2) Elevator and all movable tail plane rotation modes should be determined with the pilot's controls fixed and free. Elevator rotation with the stick fixed is defined as the peak frequency at which both elevators are in phase for symmetric rotation, and out of phase for anti-symmetric rotation. For all moving tailplanes or elevators with stability augmentation systems (control column bob weights and down springs), normal opposed operation with stick free will occur when the control stick and elevator are responding out of phase.
(3) The effect of variations in contro l cable tension should be investigated.
Typically a nominal value is specified for the cable tension along with certain tolerances. During the tests, the control surface cable can be, rigged to nominal tension, minimum tension, and maximum tension, and the tests repeated for each case. The variation, if any, can then be observed.
g. Tab Rotation. Rotational modes for irreversible trim and servo tabs are determined experimentally to supplement the calculated frequency obtained from measured stiffness in the free play tests. Tab rotation frequency will usually vary with angular deflection and is determined at maximum trailing edge up, neutral, and maximum trailing edge down positions to determine the range of tab frequencies. For geared tabs, the rotation frequency is usually determined with the control surface at maximum deflections and at neutral.
(1) Large tabs, either wide chord or very long with a single actuator, often tend to be difficult to measure in a resonance test. Wide chord tabs often become significantly involved with "plate modes" of their carrying surfaces, while long narrow tabs may have their lowest frequency in a torsional mode rather than rotation. On this basis, it may be
A1-9 necessary to survey each response frequency rather extensively to properly define each mode.
(2) Test requirements for spring tabs are dependent upon the tab control system design. In general, the following tests should be conducted to provide the required data for a mathematical representation of a spring tab system. (These tests are similar to those discussed in the previous paragraph for all moving tailplane systems.)
(a) For a preloading spring, tests should be performed for several amplitudes including complete removal of the preload, if practical.
(b) Frequency of the control surface, with tab locked to surface and pilots control column blocked, against the elastic restraint of the control system, a stick fixed mode.
(c) Frequency of the control column with the control surface locked to its main surface, against the elastic restraint of the control system, a stick free mode.
(d) Frequency of the tab, with the control system cables disconnected and the control surface blocked to its supporting structure, against the elastic restraint of the springs in the tab system.
(3) Spring loaded tabs are no n-linear systems, which are usually quite sensitive to small parameter changes making the design of these systems to preclude flutter most difficult. It is advisable to avoid their use unless extensive flutter analyses, including detail parameter evaluations, are conducted.
h. Structural Damping Measurements.
(1) Structural damping of each significant mode surveyed should be measured.
The most commonly used procedure is based on the measurement of the rate of decay of oscillation. This is best expressed in terms of logarithmic decrement, the natural logarithm of the ratio of two successive amplitudes. Records of the response of a reference transducer, while driving the structure at a specific frequency and obtained immediately before and after power to the shaker is cut off, provide the amplitude relationships required as shown in Figure A-3.
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Figure A-3. Structural Damping Measurements (X = Displacement, t = time)
The log decrement, δ, is then equal to ln (X1/X2); or as 0.693 Where n=no. of cycles to ½ amplitude; i.e.: X1/X2 = 2. n
For small values of damping, the damping factor, or γC/Co, can be Estimated as δ/2π and the structural damping g= δ/π (Appendix 1, references 2 and 8).
(2) Multiple Input and Multiple Output (MIMO's) Methods. MIMO's have become more or less common in recent times. These methods input random excitation in the structure and samples several outputs simultaneously. These methods enable completion of the ground vibration tests much faster, and several cases can be considered during ground tests.
i. Balance Weight Attachment. For control surfaces with balance weights mounted at one end of a cantilevered moment arm, the resonant frequency of the balance weight attachment arm should be at least 50 percent greater than the highest frequency of the fixed surface with which the control surface may couple. The control surface should be mounted in a jig and the vibrator attached to the balance weight. The input frequency is varied upward and the response of a reference transducer mounted on the balance weight is monitored to define the peak response.
All balance weight supporting structures should be designed for a limit static load of not less than 24g normal to a plane containing the hinge and the weight, and not less than 12g within that plane parallel with the hinge. The balance weight loads should be able to be carried by the control surface and by the fittings and their attachments on both sides of the hinge. Proof of these criteria can be accomplished by relatively simple static tests of the control surface mounted in a jig.