AC 23.629-1B

Subject: Means of Compliance with Title 14 CFR, Part 23, § 23.629, Flutter Date: 9/28/04 Initiated By: ACE-100 AC No:23.629-1B Change:

1. PURPOSE. This advisory circular presents information and guidance to provide one means, but not the only means of complying with § 23.629, Flutter (including divergence, and control reversal) of part 23 of the Federal Aviation Regulations. Accordingly, this material is neither mandatory nor regulatory in nature.

2. CANCELLATION. AC 23.629-1A, Means of Compliance with Title 14 CFR part 23, § 23.629, Flutter, dated October 23, 1985, is cancelled.

3. BACKGROUND. The complexity of flutter analysis has historically prompted endeavors to find simplified methods of flutter substantiation. Although the advent of electronic computers has de-emphasized the need to make drastic assumptions previously necessary to enable mathematical treatment of the flutter phenomenon, there remains a need to simplify flutter solution as much as possible consistent with safety in order to minimize the cost and effort required to show freedom from flutter. Past experiences gained by the necessity to judiciously choose degrees of freedom, and by the need to make essential parametric studies to establish practical boundaries of the effectiveness of the various physical quantities, has resulted in a generally recognized set of good practices. These good practices form the basis for this advisory circular.

Dorenda D. Baker Manager, Small Airplane Directorate Aircraft Certification Service /s/ Dorenda D. Baker

ii

CONTENTS

Paragraph Page

CHAPTER 1. METHODS OF SUBSTANTIATION

1. Ground Tests for Dynamically Similar Aircraft 1

2. Rational Analysis with Flight Flutter Tests 1

3. Rigidity and Mass Balance Criteria (Simp lified Criteria) with Flight Flutter Test 5

4. Whirl Mode 6

CHAPTER 2. MODIFICATIONS TO AIRCRAFT ALREADY CERTIFICATED

5. Re-evaluation 8

CHAPTER 3. CONTROL SURFACES AND TABS

6. Response 9

7. Balance 9

8. Vibratory Modes 9

9. Analyses 9

10. Fail Safe Requirements 10

CHAPTER 4. DIVERGENCE AND CONTROL REVERSAL

11. General 12

12. Airfoil Divergence 12

13. Control Reversal 12

APPENDIXES

APPENDIX 1. GROUND TESTING A1-1 APPENDIX 2. FLIGHT FLUTTER TESTING A2-1 APPENDIX 3. ACKNOWLEDGMENTS, REFERENCES, BIBLIOGRAPHY A3-1

1 CHAPTER 1. METHODS OF SUBSTANTIATION

1. GROUND TESTS FOR DYNAMICALLY SIMILAR AIRCRAFT. (§ 23.629(a)) Comparison of test data may be used in lieu of a totally new analysis and flutter test, in the case of dynamically similar aircraft. Comparison would usually be based upon geometry, mass and stiffness distributions, speed regime, and more importantly, upon a comparison of the measured coupled vibration modes.

a. Ground Testing would normally include:

(1) Ground Vibration Testing

(2) Control Surfaces and Tab Mass Property Determination

(3) Stiffness Tests of Wings, Stabilizers, etc.

(4) Free Play Measurement of All Control Surfaces and Tabs

(5) Rotational Frequency for All Control Surfaces and Tabs

(6) Rotational Stiffness for Control System and Tab System.

b. Appendix 1 presents some guidelines for recommended tests and procedures.

c. The degree of similarity between aircraft that is required for flutter substantiation can vary greatly. Some of the factors, which should be considered are the amount of safety margins available, flutter speed sensitivity to certain parameters, and the thoroughness of the original analysis. There are no hard and fast rules. Each project must be evaluated using engineering judgment. However, consider the following:

• The airplanes should be similar in weight. (You can't use dynamic similarity to compare a 5,000-pound airplane with a 19,000-pound airplane).

• The airplanes should have a similar speed range (You can't use dynamic similarity to compare a 120-knot airplane to a 250-knot airplane).

• The airplanes should be geometrically similar. (You can't use dynamic similarity to compare a V-Tail configuration airplane to a Cruciform or T-Tail configuration airplane).

• The airplanes should be similar in mass and stiffness distribution. (You can't use dynamic similarity to compare an airplane with wing-mounted engines to an airplane with aft fuselage mounted engines).

• The aircraft should have similar control systems and architecture. (You can't use dynamic similarity to compare an airplane with unassisted manual mechanical controls to one with a sophisticated powered automatic flight control system).

2. RATIONAL ANALYSIS WITH FLIGHT FLUTTER TESTS. (§ 23.629(b and c))

a. Flight Flutter Tests. The latest amendment (23-48) of section § 23.629 requires certification based upon flight test, and either rational analysis (§ 23.629(c)) or the rigidity and mass balance criteria (§ 23.629(d)). Earlier amendment levels did not require both analysis/criteria and flight flutter test.

(1) The results of any of the rational an alysis procedures required by § 23.629(c), could be used to provide guidance for formulating a flight flutter test plan. The rigidity and mass balance criteria specified in § 23.629(d) may be used in lieu of a rational analysis for certain airplanes (see paragraph 3) for simplicity.

(2) However, it is not recommended to use ONLY the rigidity and mass balance criteria (§ 23.629(d)) prior to flight flutter tests. Unless a rational analysis is accomplished, the critical mode is unknown, and this mode must be excited during flight flutter tests.

Extensive instrumentation is required to record the vibration responses during flight. An excitation system must be available to excite all the significant modes of the airplane when a rational analysis has not been accomplished prior to the flight flutter test. Each mode, including those affected by airplane weight or fuel, would have to be excited if an analysis has not been conducted to determine the critical modes. Simple pilot induced control impulses may not be adequate. It is highly recommended that a rational analysis be accomplished for all airplanes prior to flight flutter test, even if the airplane qualifies for the use of the rigidity and mass balance criteria, and it was used for the airplane design. In all cases, as required by § 23.629(a), the natural frequencies and the mode shapes of main structural components should be determined by vibration tests or other approved methods prior to conducting any flight testing. A more thorough discussion of flight flutter testing is presented in Appendix 2.

b. Review of Past Analysis. Review of previous flutter analyses conducted upon similar aircraft can provide the engineer with useful information regarding trends, critical modes, etc. Although in general such a review is not used as a substantiation basis for a new aircraft, it can provide a useful tool in evaluating the effect of modifications to existing certified aircraft.

c. Two-Dimensional Analysis. The flutter characteristics of straight wings (or tails) of large aspect ratio can be predicted reasonably well by considering a "representative section” that has two or three degrees of freedom. Translation and pitch are always needed for main surfaces, and the third degree of freedom for analysis with control surfaces is the rotation about the hinge line.

d. Three-Dimensional Analysis. This analysis is based upon consideration of total span, rather than “representative section” discussed in paragraph 2c. The behavior is integrated over the whole structure being analyzed. Some idealization is always necessary; the most common being the division of the span into strips. Other types of modeling such as doublet lattice box methods are also used.

(1) For part 23 airplanes, the wing a nd empennage analyses may be conducted separately; however, this is not always adequate for unconventional configurations or airplanes capable of speeds in excess of 0.6 Mach. Both the symmetric and anti-symmetric modes require

investigation in the flutter analysis.

(2) Calculated mass and stiffness distributions are generally used to calculate uncoupled modes and frequencies of the airplane components such as wing, stabilizer, etc.

These modal parameters are then used to conduct coupled vibration analysis of the whole airplane. The resulting coupled modes and frequencies are then compared with measured natural frequencies and mode shapes obtained from the ground vibration tests.

(3) The calculated stiffness-related inputs are generally adjusted until good agreement is obtained with the ground vibration test data. This analytical model is validated by the ground tests. The validated model is then used in the flutter analysis.

(4) It is suggested that one perform para metric variation of stiffness and other important parameters in the assumed input conditions to see which parameters are critical.

Control surface balance conditions and control system frequencies (especially tab frequencies) are often investigated parametrically. The effect of control system tension values at the low and high ends of the tolerance range should be assessed during flight flutter tests.

(5) It may be advisable to vary certain main surface frequencies (stiffness), especially torsional frequencies and engine mode frequencies, while leaving other frequencies constant, to study the effect of variations in stiffness of the control system.

(6) Sometimes it is desirable to evaluate the effect of a slight shift in span wise node location for a very massive item where the node is located very close to or within the item.

(Test data may not be sufficiently accurate for this assessment.)

(7) It is normal practice to run a density-altitude check to include near-sea-level, maximum and any other pertinent altitudes such as the knee of the airspeed-altitude envelope where the design dive speed becomes MACH-limited.

(8) It is essential to in vestigate aircraft coupled modes for airplanes with VD of 260 KEAS/0.6 Mach or above, as well as for airplanes with unconventional configurations or complex control systems.

(9) Flutter Analysis Evaluation. The resulting output of flutter analysis consists of a number of theoretical damping values (g) with associated airspeeds and flutter frequencies.

(10) Various cross plots of these values among themselves and versus varied input parameters allow a study of trends. Common plots are: damping versus equivalent airspeed (V-g plots), control surface balance versus flutter speed, modal frequency versus flutter speed, altitude versus flutter speed, etc. Normally only the critical items will be extensively compared.

(11) Of particular importance is evaluating the crossing of a damping velocity (V-g) curve toward the unstable region, through the zero damping line. The typical critical V-g curve will first become increasingly stable with increasing speed, then the damping will decrease and finally cross the zero damping line as in curves 3 and 4 in Figure 1-1. Figure 1-1 shows some typical characteristics.

(+ ) (-) g g = 0.03 Figure 1-1. Flutter Analysis Evaluation

(12) Curves 1 and 2 show slight trends toward instability, but do not approach actual instability.

(13) Curve 3 crosses the stability axis but, depending on the inherent structural damping, may or may not actually become unstable. Curve 4 is obviously unstable and probably violent, since its slope is steep as it passes through the zero damping line. In actual flight it may be a few miles an hour or so between completely stable and extremely unstable explosive flutter. Flight tests are not advisable when this type of plot is observed inside or at the boundary of the flight envelope.

(14) Much can be learned from V-g and V-f plots. Absolute values should be viewed with some reserve as there is no perfect one-to-one correspondence of the analytical parameters and flight parameters. Equally important is the rate of approach to instability (slope of curve).

(15) The general practice is to use a damping value of g = +0.03 (as the inherent structural damping) in the V-g plots. In Figure 1, an assumed value for the inherent structural damping value of g = +0.03 would be a positive value. However, this value should be used with caution if the slope of the curve is steep (damping decreases very rapidly with an increase in airspeed) between g = 0 and +0.03. In cases where the slope is steep (generally this would be a decrease in damping of 50 percent for a 5-10 knot airspeed increase), it is suggested that the g = 0 airspeed be at least 1.2 VD. Freedom from flutter should be shown to

1.2Vd/1.2Md.

(16) For damping curves such as (3), which peak out below 1.2 VD, the predicted damping should be no more unstable than g = +0.02 unless justification is provided by other acceptable means.

(17) For a more thorough discussion of flutter analysis refer to texts listed in Appendix 3 or other basic engineering textbooks on aeroelasticity.

3. RIGIDITY AND MASS BALANCE CRITERIA. (§ 23.629(d))

a. Guidelines. Airframe and Equipment Engineering Report No. 45 is intended to serve as a guide to the airplane designer in the prevention of flutter, aileron reversal, and wing divergence.

This method may only be used for airplanes that meet the criteria specified in § 23.629(d)(1),(2), and

(3). The material presented relies upon:

(1) A statistical study of the ge ometric, inertia, and elastic properties of those airplanes, which had experienced flutter in flight, and the methods used to eliminate the flutter.

(2) Limited wind-tunnel tests conducted with semi-rigid models. These were solid models of high rigidity with motion controlled at the root by springs to simulate wing bending and torsion. Springs at the control surface were used to simulate rotation.

(3) Analytic studies based on the two-dimensi onal study of a representative section of an airfoil.

b. Wing and Aileron. Prevention of wing flutter is attempted through careful attention to three parameters; wing torsional flexibility, aileron balance, and aileron free play.

(1) The aileron balance criteria is obtained from the aileron product of inertia, K, about the wing fundamental bending node line and the aileron hinge line; and the aileron mass moment of inertia, I, about its hinge line. A limit of the parameter, K/I, is set as a function of VD.

(2) A wing torsional flexibility factor, F, is defined and a limit established as a function of VD. In order to apply the criteria, one needs to know wing twist distribution per unit applied torque wing platform, and limit dive speed.

(3) The total free play of each aileron with the other aileron clamped to the wing must not exceed the specified maximum.

c. Elevator and Rudder. Dynamic balance criteria for the elevator and rudder (similar to the K/I of the aileron) are defined and limits set as a function of limit dive speed. In order to utilize the criteria, the following information is required:

(1) Geometry Horizontal tail semichord at the midspan Semispan of horizontal tail Distance from fuselage torsion axis to tip of fin Semichord of vertical tail measured at 70 percent span position

(2) Stiffness Fuselage ve rtical bending frequency Fuselage torsional frequency Fuselage lateral bending frequency

(3) Mass Elevator static balance about hinge line Elevator mass moment of inertia about hinge line Elevator product of inertia referred to stabilize centerline and elevator hinge line Rudder static balance about hinge line

6 Product of inertia of rudder referred to fuselage torsion axis and rudder hinge line Rudder mass moment of inertia about hinge line.

d. Tabs. In accordance with reference 1, all reversible tabs should be 100 percent statically mass balanced about the tab hinge line. In practice, most tabs are irreversible, which means:

(1) For any position of the c ontrol surface and tab, no appreciable deflection of the tab can be produced by means of a moment applied directly to the tab when the control surface is held in a fixed position.

(2) The total free play at the tab trailing edge must be less than 2.5 percent of the tab chord aft of the hinge line, at the station where the free play is measured (this is specified in reference 1). However, since in service wear may increase the free play, we recommend the following tighter tolerances as a maximum for the initial design:

(a) If the tab span does not exceed 35 percent of the span of the supporting control surface, the total free play should not exceed two percent of the distance from the tab hinge line to the trailing edge of the tab perpendicular to the tab hinge line.

(b) If the tab span equals or exceeds 35 percent of the span of the supporting control surface, the total free play should not exceed one percent of the distance from the tab hinge line to the trailing edge of the tab perpendicular to the tab hinge line.

(3) The tab natural frequency should be equal to or should exceed the value calculated in reference 1 and expressed as a function of tab and control surface geometry and airplane dive speed.

(4) Spring loaded tabs are free to rotate and t hus are not irreversible. Generally, these tabs will require dynamic as well as static balance. Extensive flutter analysis is always needed to define these requirements.

4. WHIRL MODE. (23.629(e)) Beginning with amendment 23-7, paragraph 23.629(e) required an investigation of the whirl mode phenomena for multiengine turbopropeller airplanes only. The basis being these airplanes characteristically have wing-mounted engines wherein the stability of a flexibly mounted engine/propeller on an elastic wing is of major concern. Amendment 23-31 of § 23.629(e) now requires an investigation of the whirl mode phenomena for both single and multiengine turbopropeller airplanes. Although airframe influence may be negligible for fuselage mounted single engine tractor configurations, the potential for propeller whirl flutter still exists. For pusher configurations, empennage motion may be significantly affected by engine/propeller forces. Stability of either installation is dictated, in part, by engine mount stiffness, damping, mass properties, motion axes, propeller geometry and propeller advance ratio. Therefore, to assure freedom from whirl mode flutter, all turbopropeller installation investigations should include, in addition to the appropriate airframe degrees of freedom:

a. Whirl mode degree of freedom, which takes into account the stability of the plane of rotation of the propeller and significant elastic, inertial, and aerodynamic forces.

b. Propeller, engine, engine mount , and airplane structure stiffness and damping variations appropriate to the particular configuration; e.g., deterioration of engine isolators, large cantilevered engine installations) etc.

In addition, references 9, 10, 11, 12 and 13 of Appendix 3 contain technical information for an acceptable means of demonstrating whirl mode stability.

8 CHAPTER 2. MODIFICATIONS TO AIRCRAFT ALREADY CERTIFICATED

5. RE-EVALUATION. Considerable judgment is often required to determine the degree of re-evaluation necessary. If the mass, mass distribution, or the stiffness distribution are affected sufficiently to result in possible significant changes in resonant frequencies of major modes, mode shapes, or mass coupling terms in the flutter equations, then some re-evaluation, such as pre-mod and post-mod Ground Vibration Test (GVT) data comparison, or analysis may be required. Some examples of significant changes are:

a. Engine (Propeller). A change in mass or mass moment of inertia of the powerplant or in its mounting system (bushings, etc.) or a Center of Gravity (cg.) shift should be investigated.

On single-engine airplanes, such changes will most likely affect fuselage and empennage frequencies and mode shapes. For engines mounted on the wings, the entire airplane may be affected.

For changes in existing designs which entail significant increases in engine power, special assessments of the effect on primary and secondary control systems should be made.

If,

(1) Tabs are exposed to the propeller slip stream, And;

(2) the certification basis of the change is to amendment level 23 or later, or;

(3) the changed product rule is applied in order to bring the original certification basis for the change up to amendment level 23 or later, then,

the fail-safe criterion described and discussed in Chapter 3 applies.

b. Structural Cutouts. Severing or bridging across major structural members, such as fuselage bulkheads and ribs or stringers of aerodynamic surfaces, may produce discontinuities in stiffness parameters that significantly alter the vibratory response of the structure.

The significance of a change may be ascertained by flutter analysis.

c. Change in the Mass. Adding of concentrated masses such as a tip tank or an external pod to a lifting surface could have a profound influence on the flutter speeds. This influence needs to be investigated by analysis and/or flight testing.

9 CHAPTER 3. CONTROL SURFACES AND TABS

6. RESPONSE. The aerodynamic force on an airfoil is very sensitive to control surface displacement, which in turn is responsive to both control motions and aerodynamic forces from tab displacement. Control surface displacement may result from deflection of the control system, deflection of the control surface attachment, or structural deflection of the control surface itself under forces from control application, aerodynamic force due to position or rate of position change, and inertia force.

7. BALANCE. Control surfaces and tabs are mass balanced to prevent rotation about their hinges resulting from inertial response to motion of the main (primary) surface, in any flutter mode. When the flutter mode consists of motion about some axis perpendicular to the control surface hinge axis, a concentrated ballast is most efficiently used. Caution should be used to assure that its location is in a high response area of the mode, which is difficult to determine when the mode is complex. Caution should also be used to assure that its attachment is secure to resist the inertia loads. Because the attachment is subjected to oscillatory loads, which cause fatigue failures and because a distributed ballast achieves balance against all flutter modes, it is conservative to distribute the ballast in accordance with the spanwise weight distribution of the surfaces. If less than static balance is provided, the effect of variations in the amount of balance should be evaluated. To guard against unintended balance changes in service, sealing and proper drain holes should be provided to minimize the risk of water, ice, or dirt accumulation in a control surface or tab. Excessive accumulation of these substances could alter the static and/or dynamic balance of the control sufficiently to adversely affect flutter characteristics.

8. VIBRATORY MODES. Control surface rotation about its hinge line is affected by various constraints. Control system stiffness, moment of inertia about the hinge line, and the rigidity of interconnection between control surfaces determine the control surface rotational modes. Both symmetric and anti-symmetric modes should be considered in the vibration analysis. Vibration mode changes resulting from modifications to the control system such as the addition of a bob weight must be assessed for their effect on flutter. Secondary rotations may result from flexure of the attaching structure or bending of the control surface. This is a major consideration for long short-chord tabs and may affect their effective irreversible characteristics. When it is necessary to raise a tab frequency by redesign, consideration should be given to the contributions of: hinge bending perpendicular to the surface especially near the horn-actuator station, horn length, axial stiffness of the push-pull rod or link, mounting flexibility, and lateral stability at push-rod attachment of the tab actuating mechanism.

9. ANALYSES. In most cases involving control surfaces, the flutter speeds are largely governed by the mass balance weights and their distributions. It is wise for the flutter analyst to cover a range of balance values and distributions to determine the most satisfactory ones. It is common to find that a change, which improves one mode degrades another. When conducting a multi-degree-of-freedom analysis, it is advisable to investigate the effect of control system frequency from zero to about 1-1/2 times the system frequency measured in test.

Due to friction, etc., it may be difficult to excite and measure control system frequency accurately. The stiffness can be measured at the surface with the control column or rudder pedal locked in the cockpit and, using the inertia of the end items, the system frequency can be calculated.

10 Theoretical values of tab and control surface aerodynamic derivatives have, for some configurations, produced higher flutter speeds than flutter model testing. Analytically derived tab and control surface aerodynamic coefficients based on strip theory have for some configurations produced higher flutter speeds than wind tunnel tests. Therefore, flutter speed sensitivity to variations in the theoretical coefficients should be evaluated in all control surface/tab investigations.

10. FAIL SAFE REQUIREMENTS. Amendment 23-23 of § 23.629(f) requires flutter free operation after failure, malfunction, or disconnect of any single tab element. This fail-safe requirement is extended to include a failure, malfunction, or disconnect of any element in the primary flight control system or flutter damper on airplanes that do not meet the criteria specified in §§ 23.629(d)(1) through (d)(3).

a. Potential failures that re quire investigation include, but are not limited to, tab or primary control trim actuating system, primary control actuating system (both of which includes bellcranks, pulleys, brackets, and their attachments), and control cables or push rods.

b. Possible means of compliance to actuating sy stem failures (i.e., actuators, cables, rods) may be achieved by incorporating dual systems, mass balancing the controls to counter the rotation of a disconnected or free surface, or by incorporating a combination of the above two.

Proper mass balancing, particularly for tabs, requires considerable care and knowledge of the flutter mechanism to assure adequacy of the design in suppressing flutter. Dual load path designs should include an assessment of residual strength with a single failure to assure that the remaining path will not fail before the single failure is detected during appropriate specified inspection intervals.

c. The question has arisen whethe r the hinges attaching to the control surfaces of the primary flight controls and tabs are part of the control system. Further, are the hinge fasteners and local portions of the fixed-surface structure to which the hinge attaches considered to be part of the control system and therefore to be considered in the failure, malfunction or disconnection of any single element? It could be argued that from a literal or traditional definition, the control system ends at the control horn. On the other hand, if a single element failure of a control surface or tab hinge can cause the surface to flutter, then it may be argued that the objective of the § 23.629(f) rule has not been met.

(1) These questions were investigated thoroughl y in the early 1980s, shortly after the rule change (amendment 23-23) was adopted. The following is a summary of that investigation and the resulting current policy on the matter.

• Interpretation of § 23.651 indicated that the hinges are considered part of the “control surface” rather than the “system”.

• The number of other rules pertaining to hinges and fitting factors that must be applied, generally dictate robust hinge designs.

• A review of service difficulty reports (SDRs) accomplished at that time did not indicate a significant number of failures of control surface or tab hinges.